Metallic Wall Hall Thrusters

NASA Technical Reports Server (NTRS)

Goebel, Dan Michael (Inventor); Hofer, Richard Robert (Inventor); Mikellides, Ioannis G. (Inventor)

2018-01-01

A Hall thruster apparatus having walls constructed from a conductive material, such as graphite, and having magnetic shielding of the walls from the ionized plasma has been demonstrated to operate with nearly the same efficiency as a conventional nonmagnetically shielded design using insulators as wall components. The new design is believed to provide the potential of higher power and uniform operation over the operating life of a thruster device.

Electron Transport in Hall Thrusters

NASA Astrophysics Data System (ADS)

McDonald, Michael Sean

Despite high technological maturity and a long flight heritage, computer models of Hall thrusters remain dependent on empirical inputs and a large part of thruster development to date has been heavily experimental in nature. This empirical approach will become increasingly unsustainable as new high-power thrusters tax existing ground test facilities and more exotic thruster designs stretch and strain the boundaries of existing design experience. The fundamental obstacle preventing predictive modeling of Hall thruster plasma properties and channel erosion is the lack of a first-principles description of electron transport across the strong magnetic fields between the cathode and anode. In spite of an abundance of proposed transport mechanisms, accurate assessments of the magnitude of electron current due to any one mechanism are scarce, and comparative studies of their relative influence on a single thruster platform simply do not exist. Lacking a clear idea of what mechanism(s) are primarily responsible for transport, it is understandably difficult for the electric propulsion scientist to focus his or her theoretical and computational tools on the right targets. This work presents a primarily experimental investigation of collisional and turbulent Hall thruster electron transport mechanisms. High-speed imaging of the thruster discharge channel at tens of thousands of frames per second reveals omnipresent rotating regions of elevated light emission, identified with a rotating spoke instability. This turbulent instability has been shown through construction of an azimuthally segmented anode to drive significant cross-field electron current in the discharge channel, and suggestive evidence points to its spatial extent into the thruster near-field plume as well. Electron trajectory simulations in experimentally measured thruster electromagnetic fields indicate that binary collisional transport mechanisms are not significant in the thruster plume, and experiments

Influence of the magnetic field configuration on the plasma flow in Hall thrusters

NASA Astrophysics Data System (ADS)

Andreussi, T.; Giannetti, V.; Leporini, A.; Saravia, M. M.; Andrenucci, M.

2018-01-01

In Hall propulsion, the thrust is provided by the acceleration of ions in a plasma generated in a cross-field configuration. Standard thruster configurations have annular channels with an almost radial magnetic field at the channel exit. A potential difference is imposed in the axial direction and the intensity of the magnetic field is calibrated in order to hinder the electron motion, while leaving the ions non-magnetised. Magnetic field lines can be assumed, as a first approximation, as lines of constant electron temperature and of thermalized potential. In typical thruster configurations, the discharge occurs inside a ceramic channel and, due to plasma-wall interactions, the electron temperature is typically low, less than few tens of eV. Hence, the magnetic field lines can be effectively used to tailor the distribution of the electrostatic potential. However, the erosion of the ceramic walls caused by the ion bombardment represents the main limiting factor of the thruster lifetime and new thruster configurations are currently under development. For these configurations, classical first order models of the plasma dynamics fail to grasp the influence of the magnetic topology on the plasma flow. In the present paper, a novel approach to investigate the correlation between magnetic field topology and thruster performance is presented. Due to the anisotropy induced by the magnetic field, the gradients of the plasma properties are assumed to be mainly in the direction orthogonal to the local magnetic field, thus enabling a quasi-one-dimensional description in magnetic coordinates. Theoretical and experimental investigations performed on a 5 kW class Hall thruster with different magnetic field configurations are then presented and discussed.

Hall Thruster Thermal Modeling and Test Data Correlation

NASA Technical Reports Server (NTRS)

Myers, James; Kamhawi, Hani; Yim, John; Clayman, Lauren

2016-01-01

The life of Hall Effect thrusters are primarily limited by plasma erosion and thermal related failures. NASA Glenn Research Center (GRC) in cooperation with the Jet Propulsion Laboratory (JPL) have recently completed development of a Hall thruster with specific emphasis to mitigate these limitations. Extending the operational life of Hall thursters makes them more suitable for some of NASA's longer duration interplanetary missions. This paper documents the thermal model development, refinement and correlation of results with thruster test data. Correlation was achieved by minimizing uncertainties in model input and recognizing the relevant parameters for effective model tuning. Throughout the thruster design phase the model was used to evaluate design options and systematically reduce component temperatures. Hall thrusters are inherently complex assemblies of high temperature components relying on internal conduction and external radiation for heat dispersion and rejection. System solutions are necessary in most cases to fully assess the benefits and/or consequences of any potential design change. Thermal model correlation is critical since thruster operational parameters can push some components/materials beyond their temperature limits. This thruster incorporates a state-of-the-art magnetic shielding system to reduce plasma erosion and to a lesser extend power/heat deposition. Additionally a comprehensive thermal design strategy was employed to reduce temperatures of critical thruster components (primarily the magnet coils and the discharge channel). Long term wear testing is currently underway to assess the effectiveness of these systems and consequently thruster longevity.

Magnetic Field Tailored Annular Hall Thruster with Anode Layer

NASA Astrophysics Data System (ADS)

Lee, Seunghun; Kim, Holak; Kim, Junbum; Lim, Youbong; Choe, Wonho; Korea Institute of Materials Science Collaboration

2016-09-01

Plasma propulsion system is one of the key components for advanced missions of satellites as well as deep space exploration. A typical plasma propulsion system is Hall effect thruster that uses crossed electric and magnetic fields to ionize a propellant gas and to accelerate the ionized gas to generate momentum. In Hall thruster plasmas, magnetic field configuration is important due to the fact that electron confinement in the electromagnetic fields affects both plasma and ion beam characteristics as well as thruster performance parameters including thrust, specific impulse, power efficiency, and life time. In this work, development of an anode layer Hall thruster (TAL) with magnetic field tailoring has been attempted. The TAL is possible to keep discharge in 1 to 2 kilovolts of anode voltage, which is useful to obtain high specific impulse. The magnetic field tailoring is used to minimize undesirable heat dissipation and secondary electron emission from the wall surrounding the plasma. We will report 3 W and 200 W thrusters performances measured by a pendulum thrust stand according to the magnetic field configuration. Also, the measured result will be compared with the plasma diagnostics conducted by an angular Faraday probe, a retarding potential analyzer, and a ExB probe.

Modeling a Hall Thruster from Anode to Plume Far Field

DTIC Science & Technology

2005-01-01

Hall thruster simulation capability that begins with propellant injection at the thruster anode, and ends in the plume far field. The development of a comprehensive simulation capability is critical for a number of reasons. The main motivation stems from the need to directly couple simulation of the plasma discharge processes inside the thruster and the transport of the plasma to the plume far field. The simulation strategy will employ two existing codes, one for the Hall thruster device and one for the plume. The coupling will take place in the plume

NASA's 2004 Hall Thruster Program

NASA Technical Reports Server (NTRS)

Jacobson, David T.; Manzella, David H.; Hofer, Richard R.; Peterson, Peter Y.

2004-01-01

An overview of NASA's Hall thruster research and development tasks conducted during fiscal year 2004 is presented. These tasks focus on: raising the technology readiness level of high power Hall thrusters, developing a moderate-power/ moderate specific impulse Hall thruster, demonstrating high-power/high specific impulse Hall thruster operation, and addressing the fundamental technical challenges of emerging Hall thruster concepts. Programmatic background information, technical accomplishments and out year plans for each program element performed under the sponsorship of the In-Space Transportation Program, Project Prometheus, and the Energetics Project are provided.

NASA's Hall Thruster Program 2002

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; Jacobson, David T.; Pinero, Luis R.; Manzella, David H.; Hofer, Richard R.; Peterson, Peter Y.

2002-01-01

The NASA Hall thruster program currently supports a number of tasks related to high power thruster development for a number of customers including the Energetics Program (formerly called the Space-based Program), the Space Solar Power Program, and the In-space Propulsion Program. In program year 2002, two tasks were central to the NASA Hall thruster program: 1) the development of a laboratory Hall thruster capable of providing high thrust at high power-, and 2) investigations into operation of Hall thrusters at high specific impulse. In addition to these two primary thruster development activities, there are a number of other on-going activities supported by the NASA Hall thruster program. These additional activities are related to issues such as high-power power processor architecture, thruster lifetime, and spacecraft integration.

Hybrid-PIC Computer Simulation of the Plasma and Erosion Processes in Hall Thrusters

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Katz, Ira; Mikellides, Ioannis G.; Gamero-Castano, Manuel

2010-01-01

HPHall software simulates and tracks the time-dependent evolution of the plasma and erosion processes in the discharge chamber and near-field plume of Hall thrusters. HPHall is an axisymmetric solver that employs a hybrid fluid/particle-in-cell (Hybrid-PIC) numerical approach. HPHall, originally developed by MIT in 1998, was upgraded to HPHall-2 by the Polytechnic University of Madrid in 2006. The Jet Propulsion Laboratory has continued the development of HPHall-2 through upgrades to the physical models employed in the code, and the addition of entirely new ones. Primary among these are the inclusion of a three-region electron mobility model that more accurately depicts the cross-field electron transport, and the development of an erosion sub-model that allows for the tracking of the erosion of the discharge chamber wall. The code is being developed to provide NASA science missions with a predictive tool of Hall thruster performance and lifetime that can be used to validate Hall thrusters for missions.

NASA's Hall Thruster Program

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; Jacobson, David T.; Rawlin, Vincent K.; Mason, Lee S.; Mantenieks, Maris A.; Manzella, David H.; Hofer, Richard R.; Peterson, Peter Y.

2001-01-01

NASA's Hall thruster program has base research and focused development efforts in support of the Advanced Space Transportation Program, Space-Based Program, and various other programs. The objective of the base research is to gain an improved understanding of the physical processes and engineering constraints of Hall thrusters to enable development of advanced Hall thruster designs. Specific technical questions that are current priorities of the base effort are: (1) How does thruster life vary with operating point? (2) How can thruster lifetime and wear rate be most efficiently evaluated? (3) What are the practical limitations for discharge voltage as it pertains to high specific impulse operation (high discharge voltage) and high thrust operation (low discharge voltage)? (4) What are the practical limits for extending Hall thrusters to very high input powers? and (5) What can be done during thruster design to reduce cost and integration concerns? The objective of the focused development effort is to develop a 50 kW-class Hall propulsion system, with a milestone of a 50 kW engineering model thruster/system by the end of program year 2006. Specific program wear 2001 efforts, along with the corporate and academic participation, are described.

Plasma Properties in the Plume of a Hall Thruster Cluster

DTIC Science & Technology

2003-06-04

The Hall thruster cluster is an attractive propulsion approach for spacecraft requiring very high-power electric propulsion systems. This article...probes in the plume of a low-power, four-engine Hall thruster cluster. Simple analytical formulas are introduced that allow these quantities to be

A Preliminary Investigation of Hall Thruster Technology

NASA Technical Reports Server (NTRS)

Gallimore, Alec D.

1997-01-01

A three-year, NASA/BMDO-sponsored experimental program to conduct performance and plume plasma property measurements on two Russian Stationary Plasma Thrusters (SPTs) has been completed. The program utilized experimental facilitates at the University of Michigan's Plasmadynamics and Electric Propulsion Laboratory (PEPL). The main features of the proposed effort were as follows: We Characterized Hall thruster [and arcjet] performance by measuring ion exhaust velocity with probes at various thruster conditions. Used a variety of probe diagnostics in the thruster plume to measure plasma properties and flow properties including T(sub e) and n(sub e), ion current density and ion energy distribution, and electric fields by mapping plasma potential. Used emission spectroscopy to identify species within the plume and to measure electron temperatures.

Electron dynamics in Hall thruster

NASA Astrophysics Data System (ADS)

Marini, Samuel; Pakter, Renato

2015-11-01

Hall thrusters are plasma engines those use an electromagnetic fields combination to confine electrons, generate and accelerate ions. Widely used by aerospace industries those thrusters stand out for its simple geometry, high specific impulse and low demand for electric power. Propulsion generated by those systems is due to acceleration of ions produced in an acceleration channel. The ions are generated by collision of electrons with propellant gas atoms. In this context, we can realize how important is characterizing the electronic dynamics. Using Hamiltonian formalism, we derive the electron motion equation in a simplified electromagnetic fields configuration observed in hall thrusters. We found conditions those must be satisfied by electromagnetic fields to have electronic confinement in acceleration channel. We present configurations of electromagnetic fields those maximize propellant gas ionization and thus make propulsion more efficient. This work was supported by CNPq.

High Power Hall Thrusters

NASA Technical Reports Server (NTRS)

Jankovsky, Robert; Tverdokhlebov, Sergery; Manzella, David

1999-01-01

The development of Hall thrusters with powers ranging from tens of kilowatts to in excess of one hundred kilowatts is considered based on renewed interest in high power. high thrust electric propulsion applications. An approach to develop such thrusters based on previous experience is discussed. It is shown that the previous experimental data taken with thrusters of 10 kW input power and less can be used. Potential mass savings due to the design of high power Hall thrusters are discussed. Both xenon and alternate thruster propellant are considered, as are technological issues that will challenge the design of high power Hall thrusters. Finally, the implications of such a development effort with regard to ground testing and spacecraft intecrati'on issues are discussed.

Modeling of Hall Thruster Lifetime and Erosion Mechanisms (Preprint)

DTIC Science & Technology

2007-09-01

Hall thruster plasma discharge has been upgraded to simulate the erosion of the thruster acceleration channel, the degradation of which is the main life-limiting factor of the propulsion system. Evolution of the thruster geometry as a result of material removal due to sputtering is modeled by calculating wall erosion rates, stepping the grid boundary by a chosen time step and altering the computational mesh between simulation runs. The code is first tuned to predict the nose cone erosion of a 200 W Busek Hall thruster , the BHT-200. Simulated erosion

Effect of Background Pressure on the Plasma Oscillation Characteristics of the HiVHAc Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Lobbia, Robert B.; Brown, Daniel L.

2014-01-01

During a component compatibility test of the NASA HiVHAc Hall thruster, a number of plasma diagnostics were implemented to study the effect of varying facility background pressure on thruster operation. These diagnostics characterized the thruster performance, the plume, and the plasma oscillations in the thruster. Thruster performance and plume characteristics as functions of background pressure were previously published. This paper focuses on changes in the plasma oscillation characteristics with changing background pressure. The diagnostics used to study plasma oscillations include a high-speed camera and a set of high-speed Langmuir probes. The results show a rise in the oscillation frequency of the "breathing" mode with rising background pressure, which is hypothesized to be due to a shortening acceleration/ionization zone. An attempt is made to apply a simplified ingestion model to the data. The combined results are used to estimate the maximum acceptable background pressure for performance and wear testing.

Plasma oscillations in a 6-kW magnetically shielded Hall thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Jorns, Benjamin A., E-mail: benjamin.a.jorns@jpl.nasa.gov; Hofer, Richard R.

2014-05-15

Plasma oscillations from 0–100 kHz in a 6-kW magnetically shielded Hall thruster are experimentally characterized with a high-speed, optical camera. Two modes are identified at 7–12 kHz and 70–90 kHz. The low frequency mode is found to be azimuthally uniform across the thruster face, while the high frequency oscillation is peaked close to the centerline-mounted cathode with an m = 1 azimuthal dependence. An analysis of these results in the context of wave-based theory suggests that the low frequency wave is the breathing mode oscillation, while the higher frequency mode is gradient-driven. The effect of these oscillations on thruster operation is examined through an analysismore » of thruster discharge current and a comparison with published observations from an unshielded variant of the thruster. Most notably, it is found that although the oscillation spectra of the two thrusters are different, they exhibit nearly identical steady-state behavior.« less

Non-Intrusive, Time-Resolved Hall Thruster Near-Field Electron Temperature Measurements

DTIC Science & Technology

2011-08-01

With the growing interest in Hall thruster technology, comes the need to fully characterize the plasma dynamics that determine performance. Of...instabilities characteristic of Hall thruster behavior, time resolved techniques must be developed. This study presents a non-intrusive method of

Optical Boron Nitride Insulator Erosion Characterization of a 200 W Xenon Hall Thruster

DTIC Science & Technology

2005-05-01

Hall thruster boron nitride insulator is evaluated as a diagnostic for real-time evaluation of thruster insulator erosion. Three Hall thruster plasma control variables are examined: ion energy (discharge potential), ion flux (propellant flow), and plasma conductivity (magnetic field strength). The boron emission, and hence the insulator erosion rate, varies linearly with ion energy and ion flux. A minimum erosion rate appears at intermediate magnetic field strengths. This may indicate that local plasma conductivity significantly affects the divergence

Effects of wall electrodes on Hall effect thruster plasma

DOE Office of Scientific and Technical Information (OSTI.GOV)

Langendorf, S., E-mail: samuel.langendorf@gatech.edu; Walker, M., E-mail: mitchell.walker@ae.gatech.edu; High-Power Electric Propulsion Laboratory, 625 Lambert St NW, Atlanta, Georgia 30318

2015-02-15

This paper investigates the physical mechanisms that cause beneficial and detrimental performance effect observed to date in Hall effect thrusters with wall electrodes. It is determined that the wall electrode sheath can reduce ion losses to the wall if positioned near the anode (outside the dense region of the plasma) such that an ion-repelling sheath is able to form. The ability of the wall electrode to form an ion-repelling sheath is inversely proportional to the current drawn—if the wall electrode becomes the dominant sink for the thruster discharge current, increases in wall electrode bias result in increased local plasma potentialmore » rather than an ion-repelling sheath. A single-fluid electron flow model gives results that mimic the observed potential structures and the current-sharing fractions between the anode and wall electrodes, showing that potential gradients in the presheath and bulk plasma come at the expense of current draw to the wall electrodes. Secondary electron emission from the wall electrodes (or lack thereof) is inferred to have a larger effect if the electrodes are positioned near the exit plane than if positioned near the anode, due to the difference in energy deposition from the plasma.« less

Low Frequency Plasma Oscillations in a 6-kW Magnetically Shielded Hall Thruster

NASA Technical Reports Server (NTRS)

Jorns, Benjamin A.; Hofery, Richard R.

2013-01-01

The oscillations from 0-100 kHz in a 6-kW magnetically shielded thruster are experimen- tally characterized. Changes in plasma parameters that result from the magnetic shielding of Hall thrusters have the potential to significantly alter thruster transients. A detailed investigation of the resulting oscillations is necessary both for the purpose of determin- ing the underlying physical processes governing time-dependent behavior in magnetically shielded thrusters as well as for improving thruster models. In this investigation, a high speed camera and a translating ion saturation probe are employed to examine the spatial extent and nature of oscillations from 0-100 kHz in the H6MS thruster. Two modes are identified at 8 kHz and 75-90 kHz. The low frequency mode is azimuthally uniform across the thruster face while the high frequency oscillation is concentrated close to the thruster centerline with an m = 1 azimuthal dependence. These experimental results are discussed in the context of wave theory as well as published observations from an unshielded variant of the H6MS thruster.

Thrust performance, propellant ionization, and thruster erosion of an external discharge plasma thruster

NASA Astrophysics Data System (ADS)

Karadag, Burak; Cho, Shinatora; Funaki, Ikkoh

2018-04-01

It is quite a challenge to design low power Hall thrusters with a long lifetime and high efficiency because of the large surface area to volume ratio and physical limits to the magnetic circuit miniaturization. As a potential solution to this problem, we experimentally investigated the external discharge plasma thruster (XPT). The XPT produces and sustains a plasma discharge completely in the open space outside of the thruster structure through a magnetic mirror configuration. It eliminates the very fundamental component of Hall thrusters, discharge channel side walls, and its magnetic circuit consists solely of a pair of hollow cylindrical permanent magnets. Thrust, low frequency discharge current oscillation, ion beam current, and plasma property measurements were conducted to characterize the manufactured prototype thruster for the proof of concept. The thrust performance, propellant ionization, and thruster erosion were discussed. Thrust generated by the XPT was on par with conventional Hall thrusters [stationary plasma thruster (SPT) or thruster with anode layer] at the same power level (˜11 mN at 250 W with 25% anode efficiency without any optimization), and discharge current had SPT-level stability (Δ < 0.2). Faraday probe measurements revealed that ion beams are finely collimated, and plumes have Gaussian distributions. Mass utilization efficiencies, beam utilization efficiencies, and plume divergence efficiencies ranged from 28 to 62%, 78 to 99%, and 40 to 48%, respectively. Electron densities and electron temperatures were found to reach 4 × 1018 m-3 ( ∂ n e / n e = ±52%) and 15 eV ( ∂ T e / T e = ±10%-30%), respectively, at 10 mm axial distance from the anode centerline. An ionization mean free path analysis revealed that electron density in the ionization region is substantially higher than the conventional Hall thrusters, which explain why the XPT is as efficient as conventional ones even without a physical ionization chamber. Our findings

Hall Thruster

NASA Image and Video Library

2017-03-06

NASA Glenn engineer Dr. Peter Peterson prepares a high-power Hall thruster for ground testing in a vacuum chamber that simulates the environment in space. This high-powered solar electric propulsion thruster has been identified as a critical part of NASA’s future deep space exploration plans.

Development and Characterization of High-Efficiency, High-Specific Impulse Xenon Hall Thrusters

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Jacobson, David (Technical Monitor)

2004-01-01

This dissertation presents research aimed at extending the efficient operation of 1600 s specific impulse Hall thruster technology to the 2000 to 3000 s range. Motivated by previous industry efforts and mission studies, the aim of this research was to develop and characterize xenon Hall thrusters capable of both high-specific impulse and high-efficiency operation. During the development phase, the laboratory-model NASA 173M Hall thrusters were designed and their performance and plasma characteristics were evaluated. Experiments with the NASA-173M version 1 (v1) validated the plasma lens magnetic field design. Experiments with the NASA 173M version 2 (v2) showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. Comparison of the thrusters showed that efficiency can be optimized for specific impulse by varying the plasma lens. During the characterization phase, additional plasma properties of the NASA 173Mv2 were measured and a performance model was derived. Results from the model and experimental data showed how efficient operation at high-specific impulse was enabled through regulation of the electron current with the magnetic field. The electron Hall parameter was approximately constant with voltage, which confirmed efficient operation can be realized only over a limited range of Hall parameters.

a Permanent Magnet Hall Thruster for Satellite Orbit Maneuvering with Low Power

NASA Astrophysics Data System (ADS)

Ferreira, Jose Leonardo

Plasma thrusters are known to have some advantages like high specific impulse. Electric propulsion is already recognized as a successful technology for long duration space missions. It has been used as primary propulsion system on earth-moon orbit trnsfer missions, comets and asteroids exploration and on commercially geosyncronous satellite attitude control systems. Closed Drift Plasma Thrusters, also called Hall Thrusters or SPT (Stationary Plasma Thruster) was conceived inthe USSR and, since then, they have been developed in several countries such as France, USA, Japan and Brazil. In this work, introductory remarks are made with focus on the most significant contributions of the electric propulsion to the progress of space missions and its future role on the brazillian space program. The main features of an inedit Permanent Magnet Hall Thruster (PMHT) developed at the Plasma Laboratory of the University of Brasilia is presented. The idea of using an array of permanent magnets, instead of an eletromagnet, to produce a radial magnetic field inside the cylindrical plasma drift channel of the thruster is a very important improvement, because it allows the possibility of developing a Hall Thruster with electric power consumption low enough to be used in small and medium size satellites. The new Halĺplasma source characterization is presented with plasma density, temperature and potential space profiles. Ion temperature mesurements based on Doppler broadening of spectral lines and ion energy measurements of the ejected plasma plume are also shown. Based on the mesured parameters of the accelerated plasma we constructed a merit figure for the PMHT. We also perform numerical simulations of satellite orbit raising from an altitude of 700 km to 36000 km using a PMHT operating in the 100 mN to 500 mN thrust range. In order to perform these caculations, integration techniques of spacecraft trajectory were used. The main simulation parameters were: orbit raising time

End-hall thrusters

NASA Technical Reports Server (NTRS)

Kaufman, H. R.; Robinson, R. S.; Day, M. L.; Haag, T. W.

1990-01-01

The end-Hall thruster can provide electric propulsion with fixed masses, specific impulses, and power-to-thrust ratios intermediate of an arcjet and a gridded (electrostatic) ion thruster. With these characteristics, this thruster is a candidate for missions of intermediate difficulty, such as the north-south stationkeeping of geostationary satellites.

Improving Hall Thruster Plume Simulation through Refined Characterization of Near-field Plasma Properties

NASA Astrophysics Data System (ADS)

Huismann, Tyler D.

Due to the rapidly expanding role of electric propulsion (EP) devices, it is important to evaluate their integration with other spacecraft systems. Specifically, EP device plumes can play a major role in spacecraft integration, and as such, accurate characterization of plume structure bears on mission success. This dissertation addresses issues related to accurate prediction of plume structure in a particular type of EP device, a Hall thruster. This is done in two ways: first, by coupling current plume simulation models with current models that simulate a Hall thruster's internal plasma behavior; second, by improving plume simulation models and thereby increasing physical fidelity. These methods are assessed by comparing simulated results to experimental measurements. Assessment indicates the two methods improve plume modeling capabilities significantly: using far-field ion current density as a metric, these approaches used in conjunction improve agreement with measurements by a factor of 2.5, as compared to previous methods. Based on comparison to experimental measurements, recent computational work on discharge chamber modeling has been largely successful in predicting properties of internal thruster plasmas. This model can provide detailed information on plasma properties at a variety of locations. Frequently, experimental data is not available at many locations that are of interest regarding computational models. Excepting the presence of experimental data, there are limited alternatives for scientifically determining plasma properties that are necessary as inputs into plume simulations. Therefore, this dissertation focuses on coupling current models that simulate internal thruster plasma behavior with plume simulation models. Further, recent experimental work on atom-ion interactions has provided a better understanding of particle collisions within plasmas. This experimental work is used to update collision models in a current plume simulation code. Previous

Hall Thruster Plume Measurements On-Board the Russian Express Satellites

NASA Technical Reports Server (NTRS)

Manzella, David; Jankovsky, Robert; Elliott, Frederick; Mikellides, Ioannis; Jongeward, Gary; Allen, Doug

2001-01-01

The operation of North-South and East-West station-keeping Hall thruster propulsion systems on-board two Russian Express-A geosynchronous communication satellites were investigated through a collaborative effort with the manufacturer of the spacecraft. Over 435 firings of 16 different thrusters with a cumulative run time of over 550 hr were reported with no thruster failures. Momentum transfer due to plume impingement was evaluated based on reductions in the effective thrust of the SPT-100 thrusters and induced disturbance torques determined based on attitude control system data and range data. Hall thruster plasma plume effects on the transmission of C-band and Ku-band communication signals were shown to be negligible. On-orbit ion current density measurements were made and subsequently compared to predictions and ground test data. Ion energy, total pressure, and electric field strength measurements were also measured on-orbit. The effect of Hall thruster operation on solar array performance over several months was investigated. A subset of these data is presented.

Hall Thruster Technology for NASA Science Missions

NASA Technical Reports Server (NTRS)

Manzella, David; Oh, David; Aadland, Randall

2005-01-01

The performance of a prototype Hall thruster designed for Discovery-class NASA science mission applications was evaluated at input powers ranging from 0.2 to 2.9 kilowatts. These data were used to construct a throttle profile for a projected Hall thruster system based on this prototype thruster. The suitability of such a Hall thruster system to perform robotic exploration missions was evaluated through the analysis of a near Earth asteroid sample return mission. This analysis demonstrated that a propulsion system based on the prototype Hall thruster offers mission benefits compared to a propulsion system based on an existing ion thruster.

An Inversion Method for Reconstructing Hall Thruster Plume Parameters from the Line Integrated Measurements (Postprint)

DTIC Science & Technology

2007-07-01

Technical Paper 3. DATES COVERED (From - To) 4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER An Inversion Method for Reconstructing Hall Thruster Plume...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 An Inversion Method for Reconstructing Hall Thruster Plume Parameters from Line Integrated Measurements... Hall thruster is a high specific impulse electric thruster that produces a highly ionized plasma inside an annular chamber through the use of high

Magnetic mirror effect in a cylindrical Hall thruster

NASA Astrophysics Data System (ADS)

Jiang, Yiwei; Tang, Haibin; Ren, Junxue; Li, Min; Cao, Jinbin

2018-01-01

For cylindrical Hall thrusters, the magnetic field geometry is totally different from that in conventional Hall thrusters. In this study, we investigate the magnetic mirror effect in a fully cylindrical Hall thruster by changing the number of iron rings (0-5), which surround the discharge channel wall. The plasma properties inside the discharge channel and plume area are simulated with a self-developed PIC-MCC code. The numerical results show significant influence of magnetic geometry on the electron confinement. With the number of rings increasing above three, the near-wall electron density gap is reduced, indicating the suppression of neutral gas leakage. The electron temperature inside the discharge channel reaches its peak (38.4 eV) when the magnetic mirror is strongest. It is also found that the thruster performance has strong relations with the magnetic mirror as the propellant utilisation efficiency reaches the maximum (1.18) at the biggest magnetic mirror ratio. Also, the optimal magnetic mirror improves the multi-charged ion dynamics, including the ion production and propellant utilisation efficiency.

Effect of multiply charged ions on the performance and beam characteristics in annular and cylindrical type Hall thruster plasmas

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kim, Holak; Lim, Youbong; Choe, Wonho, E-mail: wchoe@kaist.ac.kr

2014-10-06

Plasma plume and thruster performance characteristics associated with multiply charged ions in a cylindrical type Hall thruster (CHT) and an annular type Hall thruster are compared under identical conditions such as channel diameter, channel depth, propellant mass flow rate. A high propellant utilization in a CHT is caused by a high ionization rate, which brings about large multiply charged ions. Ion currents and utilizations are much different due to the presence of multiply charged ions. A high multiply charged ion fraction and a high ionization rate in the CHT result in a higher specific impulse, thrust, and discharge current.

Spacecraft Interactions Studies with a 1 Kw Class Closed-Drift Hall Thruster

DTIC Science & Technology

1998-01-31

Closed Drift Hall thruster plume with spacecraft surfaces and systems. Two basic interaction modes were investigated: (1) the influence of the plume...Spectrometer (MBMS) capable of discerning both the mass and energy of Hall thruster plume species, and the ion acoustic wave probe to measure the drift velocity of the plume plasma.

Conducting wall Hall thrusters in magnetic shielding and standard configurations

NASA Astrophysics Data System (ADS)

Grimaud, Lou; Mazouffre, Stéphane

2017-07-01

Traditional Hall thrusters are fitted with boron nitride dielectric discharge channels that confine the plasma discharge. Wall properties have significant effects on the performances and stability of the thrusters. In magnetically shielded thrusters, interactions between the plasma and the walls are greatly reduced, and the potential drop responsible for ion acceleration is situated outside the channel. This opens the way to the utilization of alternative materials for the discharge channel. In this work, graphite walls are compared to BN-SiO2 walls in the 200 W magnetically shielded ISCT200-MS and the unshielded ISCT200-US Hall thrusters. The magnetically shielded thruster shows no significant change in the discharge current mean value and oscillations, while the unshielded thruster's discharge current increases by 25% and becomes noticeably less stable. The electric field profile is also investigated through laser spectroscopy, and no significant difference is recorded between the ceramic and graphite cases for the shielded thruster. The unshielded thruster, on the other hand, has its acceleration region shifted 15% of the channel length downstream. Lastly, the plume profile is measured with planar probes fitted with guard rings. Once again the material wall has little influence on the plume characteristics in the shielded thruster, while the unshielded one is significantly affected.

Performance of a Permanent-Magnet Cylindrical Hall-Effect Thruster

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Sooby, E. S.; Kimberlin, A. C.; Raites, Y.; Merino, E.; Fisch, N. J.

2009-01-01

The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic topologies. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying higher thrust efficiency. Thruster performance measurements on this configuration were obtained over a power range of 70-350 W and with the cathode orifice located at three different axial positions relative to the thruster exit plane. The thrust levels over this power range were 1.25-6.5 mN, with anode efficiencies and specific impulses spanning 4-21% and 400-1950 s, respectively. The anode efficiency of the permanent-magnet thruster compares favorable with the efficiency of the electromagnet thruster when the power consumed by the electromagnets is taken into account.

Electron Transport and Ion Acceleration in a Low-power Cylindrical Hall Thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

A. Smirnov; Y. Raitses; N.J. Fisch

2004-06-24

Conventional annular Hall thrusters become inefficient when scaled to low power. Cylindrical Hall thrusters, which have lower surface-to-volume ratio, are therefore more promising for scaling down. They presently exhibit performance comparable with conventional annular Hall thrusters. Electron cross-field transport in a 2.6 cm miniaturized cylindrical Hall thruster (100 W power level) has been studied through the analysis of experimental data and Monte Carlo simulations of electron dynamics in the thruster channel. The numerical model takes into account elastic and inelastic electron collisions with atoms, electron-wall collisions, including secondary electron emission, and Bohm diffusion. We show that in order to explainmore » the observed discharge current, the electron anomalous collision frequency {nu}{sub B} has to be on the order of the Bohm value, {nu}{sub B} {approx} {omega}{sub c}/16. The contribution of electron-wall collisions to cross-field transport is found to be insignificant. The plasma density peak observed at the axis of the 2.6 cm cylindrical Hall thruster is likely to be due to the convergent flux of ions, which are born in the annular part of the channel and accelerated towards the thruster axis.« less

Magnetic field configurations on thruster performance in accordance with ion beam characteristics in cylindrical Hall thruster plasmas

NASA Astrophysics Data System (ADS)

Kim, Holak; Choe, Wonho; Lim, Youbong; Lee, Seunghun; Park, Sanghoo

2017-03-01

Magnetic field configuration is critical in Hall thrusters for achieving high performance, particularly in thrust, specific impulse, efficiency, etc. Ion beam features are also significantly influenced by magnetic field configurations. In two typical magnetic field configurations (i.e., co-current and counter-current configurations) of a cylindrical Hall thruster, ion beam characteristics are compared in relation to multiply charged ions. Our study shows that the co-current configuration brings about high ion current (or low electron current), high ionization rate, and small plume angle that lead to high thruster performance.

Magnesium Hall Thruster

NASA Technical Reports Server (NTRS)

Szabo, James J.

2015-01-01

This Phase II project is developing a magnesium (Mg) Hall effect thruster system that would open the door for in situ resource utilization (ISRU)-based solar system exploration. Magnesium is light and easy to ionize. For a Mars- Earth transfer, the propellant mass savings with respect to a xenon Hall effect thruster (HET) system are enormous. Magnesium also can be combusted in a rocket with carbon dioxide (CO2) or water (H2O), enabling a multimode propulsion system with propellant sharing and ISRU. In the near term, CO2 and H2O would be collected in situ on Mars or the moon. In the far term, Mg itself would be collected from Martian and lunar regolith. In Phase I, an integrated, medium-power (1- to 3-kW) Mg HET system was developed and tested. Controlled, steady operation at constant voltage and power was demonstrated. Preliminary measurements indicate a specific impulse (Isp) greater than 4,000 s was achieved at a discharge potential of 400 V. The feasibility of delivering fluidized Mg powder to a medium- or high-power thruster also was demonstrated. Phase II of the project evaluated the performance of an integrated, highpower Mg Hall thruster system in a relevant space environment. Researchers improved the medium power thruster system and characterized it in detail. Researchers also designed and built a high-power (8- to 20-kW) Mg HET. A fluidized powder feed system supporting the high-power thruster was built and delivered to Busek Company, Inc.

Development and characterization of high-efficiency, high-specific impulse xenon Hall thrusters

NASA Astrophysics Data System (ADS)

Hofer, Richard Robert

This dissertation presents research aimed at extending the efficient operation of 1600 s specific impulse Hall thruster technology to the 2000--3000 s range. While recent studies of commercially developed Hall thrusters demonstrated greater than 4000 s specific impulse, maximum efficiency occurred at less than 3000 s. It was hypothesized that the efficiency maximum resulted as a consequence of modern magnetic field designs, optimized for 1600 s, which were unsuitable at high-specific impulse. Motivated by the industry efforts and mission studies, the aim of this research was to develop and characterize xenon Hall thrusters capable of both high-specific impulse and high-efficiency operation. The research divided into development and characterization phases. During the development phase, the laboratory-model NASA-173M Hall thrusters were designed with plasma lens magnetic field topographies and their performance and plasma characteristics were evaluated. Experiments with the NASA-173M version 1 (v1) validated the plasma lens design by showing how changing the magnetic field topography at high-specific impulse improved efficiency. Experiments with the NASA-173M version 2 (v2) showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. Between 300--1000 V, total specific impulse and total efficiency of the NASA-173Mv2 operating at 10 mg/s ranged from 1600--3400 s and 51--61%, respectively. Comparison of the thrusters showed that efficiency can be optimized for specific impulse by varying the plasma lens design. During the characterization phase, additional plasma properties of the NASA-173Mv2 were measured and a performance model was derived accounting for a multiply-charged, partially-ionized plasma. Results from the model based on experimental data showed how efficient operation at high-specific impulse was enabled through regulation of the electron current with the magnetic field. The

Evidence of Collisionless Shocks in a Hall Thruster Plume

DTIC Science & Technology

2003-04-25

Triple Langmuir probes and emissive probes are used to measure the electron number density, electron temperature, and plasma potential downstream of a low-power Hall thruster . The results show a high density plasma core with elevated electron temperature and plasma potential along the thruster centerline. These properties are believed to be due to collisionless shocks formed as a result of the ion/ion acoustic instability. A simple model is presented that shows the existence of a collisionless shock to be consistent with the observed phenomena.

Hall thruster with grooved walls

DOE Office of Scientific and Technical Information (OSTI.GOV)

Li Hong; Ning Zhongxi; Yu Daren

2013-02-28

Axial-oriented and azimuthal-distributed grooves are formed on channel walls of a Hall thruster after the engine undergoes a long-term operation. Existing studies have demonstrated the relation between the grooves and the near-wall physics, such as sheath and electron near-wall transport. The idea to optimize the thruster performance with such grooves was also proposed. Therefore, this paper is devoted to explore the effects of wall grooves on the discharge characteristics of a Hall thruster. With experimental measurements, the variations on electron conductivity, ionization distribution, and integrated performance are obtained. The involved physical mechanisms are then analyzed and discussed. The findings helpmore » to not only better understand the working principle of Hall thruster discharge but also establish a physical fundamental for the subsequent optimization with artificial grooves.« less

Advanced Hybrid Modeling of Hall Thruster Plumes

DTIC Science & Technology

2010-06-16

Hall thruster operated in the Large Vacuum Test Facility at the University of Michigan. The approach utilizes the direct simulation Monte Carlo method and the Particle-in-Cell method to simulate the collision and plasma dynamics of xenon neutrals and ions. The electrons are modeled as a fluid using conservation equations. A second code is employed to model discharge chamber behavior to provide improved input conditions at the thruster exit for the plume simulation. Simulation accuracy is assessed using experimental data previously

Experimental and Numerical Examination of a Hall Thruster Plume (Preprint)

DTIC Science & Technology

2007-07-31

Hall thruster has been characterized through measurements from various plasma electrostatic probes. Ion current flux, plasma potential, plasma density, and electron temperatures were measured from the near-field plume to 60 cm downstream of the exit plane. These experimentally derived measurements were compared to numerical simulations run with the plasma plume code DRACO. A major goal of this study was to determine the fidelity of the DRACO numerical simulation. The effect of background pressure on the thruster plume was also examined using ion current flux measurements

Near-Surface Plasma Characterization of the 12.5-kW NASA TDU1 Hall Thruster

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Huang, Wensheng; Kamhawi, Hani

2015-01-01

To advance the state-of-the-art in Hall thruster technology, NASA is developing a 12.5-kW, high-specific-impulse, high-throughput thruster for the Solar Electric Propulsion Technology Demonstration Mission. In order to meet the demanding lifetime requirements of potential missions such as the Asteroid Redirect Robotic Mission, magnetic shielding was incorporated into the thruster design. Two units of the resulting thruster, called the Hall Effect Rocket with Magnetic Shielding (HERMeS), were fabricated and are presently being characterized. The first of these units, designated the Technology Development Unit 1 (TDU1), has undergone extensive performance and thermal characterization at NASA Glenn Research Center. A preliminary lifetime assessment was conducted by characterizing the degree of magnetic shielding within the thruster. This characterization was accomplished by placing eight flush-mounted Langmuir probes within each discharge channel wall and measuring the local plasma potential and electron temperature at various axial locations. Measured properties indicate a high degree of magnetic shielding across the throttle table, with plasma potential variations along each channel wall being less than or equal to 5 eV and electron temperatures being maintained at less than or equal to 5 eV, even at 800 V discharge voltage near the thruster exit plane. These properties indicate that ion impact energies within the HERMeS will not exceed 26 eV, which is below the expected sputtering threshold energy for boron nitride. Parametric studies that varied the facility backpressure and magnetic field strength at 300 V, 9.4 kW, illustrate that the plasma potential and electron temperature are insensitive to these parameters, with shielding being maintained at facility pressures 3X higher and magnetic field strengths 2.5X higher than nominal conditions. Overall, the preliminary lifetime assessment indicates a high degree of shielding within the HERMeS TDU1, effectively

Plasma Perturbations in High-Speed Probing of Hall Thruster Discharge Chambers: Quantification and Mitigation

NASA Technical Reports Server (NTRS)

Jorns, Benjamin A.; Goebel, Dan M.; Hofer, Richard R.

2015-01-01

An experimental investigation is presented to quantify the effect of high-speed probing on the plasma parameters inside the discharge chamber of a 6-kW Hall thruster. Understanding the nature of these perturbations is of significant interest given the importance of accurate plasma measurements for characterizing thruster operation. An array of diagnostics including a high-speed camera and embedded wall probes is employed to examine in real time the changes in electron temperature and plasma potential induced by inserting a high-speed reciprocating Langmuir probe into the discharge chamber. It is found that the perturbations onset when the scanning probe is downstream of the electron temperature peak, and that along channel centerline, the perturbations are best characterized as a downstream shift of plasma parameters by 15-20% the length of the discharge chamber. A parametric study is performed to investigate techniques to mitigate the observed probe perturbations including varying probe speed, probe location, and operating conditions. It is found that the perturbations largely disappear when the thruster is operated at low power and low discharge voltage. The results of this mitigation study are discussed in the context of recommended methods for generating unperturbed measurements of the discharge chamber plasma.

Remote Diagnostic Measurements of Hall Thruster Plumes

DTIC Science & Technology

2009-08-14

This paper describes measurements of Hall thruster plumes that characterize ion energy distributions and charge state fractions using remotely...charge state. Next, energy and charge state measurements are described from testing of a 200 W Hall thruster at AFIT. Measurements showed variation in...position. Finally, ExB probe charge state measurements are presented from a 6-kW laboratory Hall thruster operated at low discharge voltage levels at AFRL

Power Dependence of the Electron Mobility Profile in a Hall Thruster

NASA Technical Reports Server (NTRS)

Jorns, Benjamin A.; Hofery, Richard H.; Mikellides, Ioannis G.

2014-01-01

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.

Electrodeless plasma thrusters for spacecraft: A review

NASA Astrophysics Data System (ADS)

Bathgate, S. N.; Bilek, M. M. M.; McKenzie, D. R.

2017-08-01

The physics of electrodeless electric thrusters that use directed plasma to propel spacecraft without employing electrodes subject to plasma erosion is reviewed. Electrodeless plasma thrusters are potentially more durable than presently deployed thrusters that use electrodes such as gridded ion, Hall thrusters, arcjets and resistojets. Like other plasma thrusters, electrodeless thrusters have the advantage of reduced fuel mass compared to chemical thrusters that produce the same thrust. The status of electrodeless plasma thrusters that could be used in communications satellites and in spacecraft for interplanetary missions is examined. Electrodeless thrusters under development or planned for deployment include devices that use a rotating magnetic field; devices that use a rotating electric field; pulsed inductive devices that exploit the Lorentz force on an induced current loop in a plasma; devices that use radiofrequency fields to heat plasmas and have magnetic nozzles to accelerate the hot plasma and other devices that exploit the Lorentz force. Using metrics of specific impulse and thrust efficiency, we find that the most promising designs are those that use Lorentz forces directly to expel plasma and those that use magnetic nozzles to accelerate plasma.

Discharge Oscillations in a Permanent Magnet Cylindrical Hall-Effect Thruster

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Sooby, E. S.; Raitses, Y.; Merino, E.; Fisch, N. J.

2009-01-01

Measurements of the discharge current in a cylindrical Hall thruster are presented to quantify plasma oscillations and instabilities without introducing an intrusive probe into the plasma. The time-varying component of the discharge current is measured using a current monitor that possesses a wide frequency bandwidth and the signal is Fourier transformed to yield the frequency spectra present, allowing for the identification of plasma oscillations. The data show that the discharge current oscillations become generally greater in amplitude and complexity as the voltage is increased, and are reduced in severity with increasing flow rate. The breathing mode ionization instability is identified, with frequency as a function of discharge voltage not increasing with discharge voltage as has been observed in some traditional Hall thruster geometries, but instead following a scaling similar to a large-amplitude, nonlinear oscillation mode recently predicted in for annular Hall thrusters. A transition from lower amplitude oscillations to large relative fluctuations in the oscillating discharge current is observed at low flow rates and is suppressed as the mass flow rate is increased. A second set of peaks in the frequency spectra are observed at the highest propellant flow rate tested. Possible mechanisms that might give rise to these peaks include ionization instabilities and interactions between various oscillatory modes.

NASA HERMeS Hall Thruster Electrical Configuration Characterization

NASA Technical Reports Server (NTRS)

Peterson, Peter Y.; Kamhawi, Hani; Huang, Wensheng; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Hofer, Richard

2015-01-01

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in-space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This paper describes the electrical configuration testing of the HERMeS TDU-1 Hall thruster in NASA Glenn Research Center's Vacuum Facility 5. The three electrical configurations examined were 1) thruster body tied to facility ground, 2) thruster floating, and 3) thruster body electrically tied to cathode common. The HERMeS TDU-1 Hall thruster was also configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Performance of an 8 kW Hall Thruster

DTIC Science & Technology

2000-01-12

For the purpose of either orbit raising and/or repositioning the Hall thruster must be capable of delivering sufficient thrust to minimize transfer...time. This coupled with the increasing on-board electric power capacity of military and commercial satellites, requires a high power Hall thruster that...development of a novel, high power Hall thruster , capable of efficient operation over a broad range of Isp and thrust. We call such a thruster the bi

Preliminary Study of Arcjet Neutralization of Hall Thruster Clusters (Postprint)

DTIC Science & Technology

2007-01-18

Clustered Hall thrusters have emerged as a favored choice for extending Hall thruster options to very high powers (50 kW - 150 kW). This paper...examines the possible use of an arcjet to neutralize clustered Hall thrusters, as the hybrid arcjet- Hall thruster concept can fill a performance niche...and helium, and then demonstrate the first successful operation of a low power Hall thruster -arcjet neutralizer package. In the surrogate anode studies

Measurements of neutral and ion velocity distribution functions in a Hall thruster

NASA Astrophysics Data System (ADS)

Svarnas, Panagiotis; Romadanov, Iavn; Diallo, Ahmed; Raitses, Yevgeny

2015-11-01

Hall thruster is a plasma device for space propulsion. It utilizes a cross-field discharge to generate a partially ionized weakly collisional plasma with magnetized electrons and non-magnetized ions. The ions are accelerated by the electric field to produce the thrust. There is a relatively large number of studies devoted to characterization of accelerated ions, including measurements of ion velocity distribution function using laser-induced fluorescence diagnostic. Interactions of these accelerated ions with neutral atoms in the thruster and the thruster plume is a subject of on-going studies, which require combined monitoring of ion and neutral velocity distributions. Herein, laser-induced fluorescence technique has been employed to study neutral and single-charged ion velocity distribution functions in a 200 W cylindrical Hall thruster operating with xenon propellant. An optical system is installed in the vacuum chamber enabling spatially resolved axial velocity measurements. The fluorescence signals are well separated from the plasma background emission by modulating the laser beam and using lock-in detectors. Measured velocity distribution functions of neutral atoms and ions at different operating parameters of the thruster are reported and analyzed. This work was supported by DOE contract DE-AC02-09CH11466.

Comparisons and Evaluation of Hall Thruster Models

DTIC Science & Technology

2002-03-20

COVERED (FROM - TO) 20-04-2001 to 20-04-2002 4. TITLE AND SUBTITLE comparisons and Evaluation of Hall Thruster Models Unclassified 5a. CONTRACT NUMBER...TITLE AND SUBTITLE Comparisons and Evaluation of Hall Thruster Models 5c. PROGRAM ELEMENT NUMBER 5d. PROJECT NUMBER 5d. TASK NUMBER 6. AUTHOR(S...evaluation of Hall thruster models G. J. M. Hagelaar, J. Bareilles, L. Garrigues, and J.-P. Boeuf CPAT, Bâtiment 3R2, Université Paul Sabatier 118 Route

A Small Modular Laboratory Hall Effect Thruster

NASA Astrophysics Data System (ADS)

Lee, Ty Davis

Electric propulsion technologies promise to revolutionize access to space, opening the door for mission concepts unfeasible by traditional propulsion methods alone. The Hall effect thruster is a relatively high thrust, moderate specific impulse electric propulsion device that belongs to the class of electrostatic thrusters. Hall effect thrusters benefit from an extensive flight history, and offer significant performance and cost advantages when compared to other forms of electric propulsion. Ongoing research on these devices includes the investigation of mechanisms that tend to decrease overall thruster efficiency, as well as the development of new techniques to extend operational lifetimes. This thesis is primarily concerned with the design and construction of a Small Modular Laboratory Hall Effect Thruster (SMLHET), and its operation on argon propellant gas. Particular attention was addressed at low-cost, modular design principles, that would facilitate simple replacement and modification of key thruster parts such as the magnetic circuit and discharge channel. This capability is intended to facilitate future studies of device physics such as anomalous electron transport and magnetic shielding of the channel walls, that have an impact on thruster performance and life. Preliminary results demonstrate SMLHET running on argon in a manner characteristic of Hall effect thrusters, additionally a power balance method was utilized to estimate thruster performance. It is expected that future thruster studies utilizing heavier though more expensive gases like xenon or krypton, will observe increased efficiency and stability.

Multiply charged ion generation according to magnetic field configurations in Hall thruster plasmas

NASA Astrophysics Data System (ADS)

Kim, Holak; Lee, Seunghun; Kim, Junbum; Lim, Youbong; Choe, Wonho; KIMS Collaboration

2016-09-01

Plasma propulsion is the most promising techniques to operate satellites for low earth orbit as well as deep space exploration. A typical plasma propulsion system is Hall thruster (HT) that uses crossed electromagnetic fields to ionize a propellant gas and to accelerate the ionized gas. In HT the tailoring of magnetic fields is significant due to that the electron confinement in the electromagnetic fields affects thruster performances such as thrust force, specific impulse, power efficiency, and life time. We designed an anode layer HT (TAL) with the magnetic field tailoring. The TAL is possible to keep discharge in 1 2 kilovolts, which voltage is useful to obtain high specific impulse The magnetic field tailoring is adapted to minimize undesirable heat dissipations and secondary electron emissions at a wall surrounding plasma In presentation, we will report TAL performances including thrust force, specific impulse, and anode efficiency measured by a pendulum thrust stand. This mechanical measurement will be compared to the plasma diagnostics conducted by angular Faraday probe, retarding potential analyzer, and ExB probe Grant No. 2014M1A3A3A02034510.

A one-dimensional with three-dimensional velocity space hybrid-PIC model of the discharge plasma in a Hall thruster

NASA Astrophysics Data System (ADS)

Shashkov, Andrey; Lovtsov, Alexander; Tomilin, Dmitry

2017-04-01

According to present knowledge, countless numerical simulations of the discharge plasma in Hall thrusters were conducted. However, on the one hand, adequate two-dimensional (2D) models require a lot of time to carry out numerical research of the breathing mode oscillations or the discharge structure. On the other hand, existing one-dimensional (1D) models are usually too simplistic and do not take into consideration such important phenomena as neutral-wall collisions, magnetic field induced by Hall current and double, secondary, and stepwise ionizations together. In this paper a one-dimensional with three-dimensional velocity space (1D3V) hybrid-PIC model is presented. The model is able to incorporate all the phenomena mentioned above. A new method of neutral-wall collisions simulation in described space was developed and validated. Simulation results obtained for KM-88 and KM-60 thrusters are in a good agreement with experimental data. The Bohm collision coefficient was the same for both thrusters. Neutral-wall collisions, doubly charged ions, and induced magnetic field were proved to stabilize the breathing mode oscillations in a Hall thruster under some circumstances.

Sheath oscillation characteristics and effect on near-wall conduction in a krypton Hall thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zhang, Fengkui, E-mail: fengkuizhang@163.com; Kong, Lingyi; Li, Chenliang

2014-11-15

Despite its affordability, the krypton Hall-effect thruster in applications always had problems in regard to performance. The reason for this degradation is studied from the perspective of the near-wall conductivity of electrons. Using the particle-in-cell method, the sheath oscillation characteristics and its effect on near-wall conduction are compared in the krypton and xenon Hall-effect thrusters both with wall material composed of BNSiO{sub 2}. Comparing these two thrusters, the sheath in the krypton-plasma thruster will oscillate at low electron temperatures. The near-wall conduction current is only produced by collisions between electrons and wall, thereby causing a deficiency in the channel current.more » The sheath displays spatial oscillations only at high electron temperature; electrons are then reflected to produce the non-oscillation conduction current needed for the krypton-plasma thruster. However, it is accompanied with intensified oscillations.« less

NASA HERMeS Hall Thruster Electrical Configuration Characterization

NASA Technical Reports Server (NTRS)

Peterson, Peter; Kamhawi, Hani; Huang, Wensheng; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Hofer, Richard

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This presentation will cover the electrical configuration testing of the TDU-1 HERMeS Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined are the thruster body tied to facility ground, thruster floating, and finally the thruster body electrically tied to cathode common. The TDU-1 HERMeS was configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

The Influence of Current Density and Magnetic Field Topography in Optimizing the Performance, Divergence, and Plasma Oscillations of High Specific Impulse Hall Thrusters

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Jankovsky, Robert S.

2003-01-01

Recent studies of xenon Hall thrusters have shown peak efficiencies at specific impulses of less than 3000 s. This was a consequence of modern Hall thruster magnetic field topographies, which have been optimized for 300 V discharges. On-going research at the NASA Glenn Research Center is investigating this behavior and methods to enhance thruster performance. To conduct these studies, a laboratory model Hall thruster that uses a pair of trim coils to tailor the magnetic field topography for high specific impulse operation has been developed. The thruster-the NASA-173Mv2 was tested to determine how current density and magnetic field topography affect performance, divergence, and plasma oscillations at voltages up to 1000 V. Test results showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. At 1000 V, 10 milligrams per second the total specific impulse was 3390 s and the total efficiency was 60.8%. Plume divergence decreased at 400-1000 V, but increased at 300-400 V as the result of plasma oscillations. The dominant oscillation frequency steadily increased with voltage, from 14.5 kHz at 300 V, to 22 kHz at 1000 V. An additional oscillatory mode in the 80-90 kHz frequency range began to appear above 500 V. The use of trim coils to modify the magnetic field improved performance while decreasing plume divergence and the frequency and magnitude of plasma oscillations.

Hall-effect Thruster Channel Surface Properties Investigation (PREPRINT)

DTIC Science & Technology

2011-03-03

Article 3. DATES COVERED (From - To) 4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER Hall-effect Thruster Channel Surface Properties Investigation 5b...13. SUPPLEMENTARY NOTES For publication in the AIAA Journal of Propulsion and Power. 14. ABSTRACT Surface properties of Hall-effect thruster...incorporated into thruster simulations, and these models must account for evolution of channel surface properties due to thruster operation. Results from

Plasma oscillation effects on nested Hall thruster operation and stability

NASA Astrophysics Data System (ADS)

McDonald, M. S.; Sekerak, M. J.; Gallimore, A. D.; Hofer, R. R.

High-power Hall thrusters capable of throughput on the order of 100 kW are currently under development, driven by more demanding mission profiles and rapid growth in on-orbit solar power generation capability. At these power levels the nested Hall thruster (NHT), a new design that concentrically packs multiple thrusters into a single body with a shared magnetic circuit, offers performance and logistical advantages over conventional single-channel Hall thrusters. An important area for risk reduction in NHT development is quantifying inter-channel coupling between discharge channels. This work presents time- and frequency-domain discharge current and voltage measurements paired with high-speed video of the X2, a 10-kW class dual channel NHT. Two “ triads” of operating conditions at 150 V, 3.6 kW and 250 V, 8.6 kW were examined, including each channel in individual operation and both channels in joint operation. For both triads tested, dual-channel operation did not noticeably destabilize the discharge. Partial coupling of outer channel oscillations into the inner channel occurred at 150 V, though oscillation amplitudes did not change greatly. As a percentage of mean discharge current, RMS oscillations at 150 V increased from 8% to 13% on the inner channel and decreased from 10% to 8% on the outer channel from single- to dual-channel operation. At 250 V the RMS/mean level stayed steady at 13% on the inner channel and decreased from 7% to 6% on the outer channel. The only mean discharge parameter noticeably affected was the cathode floating potential, which decreased in magnitude below ground with increased absolute cathode flow rate in dual-channel mode. Rotating spokes were detected on high-speed video across all X2 operating cases with wavelength 12-18 cm, and spoke velocity generally increased from single- to dual-channel operation.

Ion behavior in low-power magnetically shielded and unshielded Hall thrusters

NASA Astrophysics Data System (ADS)

Grimaud, L.; Mazouffre, S.

2017-05-01

Magnetically shielded Hall thrusters achieve a longer lifespan than traditional Hall thrusters by reducing wall erosion. The lower erosion rate is attributed to a reduction of the high energy ion population impacting the walls. To investigate this phenomenon, the ion velocity distribution functions are measured with laser induced fluorescence at several points of interest in the magnetically shielded ISCT200-MS and the unshielded ISCT200-US Hall thrusters. The center of the discharge channel is probed to highlight the difference in plasma positioning between the shielded and unshielded thrusters. Erosion phenomena are investigated by taking measurements of the ion velocity distribution near the inner and outer wall as well as above the magnetic poles where some erosion is observed. The resulting distribution functions show a displacement of the acceleration region from inside the channel in the unshielded thruster to downstream of the exit plane in the ISCT200-MS. Near the walls, the unshielded thruster displays both a higher relative ion density as well as a significant fraction of the ions with velocities toward the walls compared to the shielded thruster. Higher proportions of high velocity ions are also observed. Those results are in accordance with the reduced erosion observed. Both shielded and unshielded thrusters have large populations of ions impacting the magnetic poles. The mechanism through which those ions are accelerated toward the magnetic poles has so far not been explained.

Plume Characteristics of the BHT-HD-600 Hall Thruster (Preprint)

DTIC Science & Technology

2006-07-01

Hall thruster on spacecraft, a number of plume properties have been measured. These include current density using a Faraday probe, ion energy distribution using a retarding potential analyzer, and ion species fractions using an E x B probe. The BHT-HD-600 Hall thruster is a nominally 600 W xenon Hall thruster developed by Busek Co. Inc. for the U.S. Air Force Research Laboratory. Plume characterization of Hall thrusters is required to fully understand the impacts of thruster operation on spacecraft. Much of these plume data are

Establishment of a Hall Thruster Cluster

DTIC Science & Technology

2004-02-01

DURIP funds were used to develop a Hall thruster cluster test facility centered around the University of Michigan Large Vacuum Test Facility and a 2x2 cluster of BUSEK 600 W BHT-600 Hall thrusters. This capability will facilitate our three-year program to address the issue of high-power CDT operation and to provide insight on how chamber effects influence CDT engine/cluster characteristics.

Experimental test of 200 W Hall thruster with titanium wall

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Sun, Hezhi; Peng, Wuji; Xu, Yu; Wei, Liqiu; Li, Hong; Li, Peng; Su, Hongbo; Yu, Daren

2017-05-01

We designed a 200 W Hall thruster based on the technology of pushing down a magnetic field with two permanent magnetic rings. Boron nitride (BN) is an important insulating wall material for Hall thrusters. The discharge characteristics of the designed Hall thruster were studied by replacing BN with titanium (Ti). Experimental results show that the designed Hall thruster can discharge stably for a long time under a Ti channel. Experiments were performed to determine whether the channel and cathode are electrically connected. When the channel wall and cathode are insulated, the divergence angle of the plume increases, but the performance of the Hall thruster is improved in terms of thrust, specific impulse, anode efficiency, and thrust-to-power ratio. Ti exhibits a powerful antisputtering capability, a low emanation rate of gas, and a large structural strength, making it a potential candidate wall material in the design of low-power Hall thrusters.

Performance of a Low-Power Cylindrical Hall Thruster

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Markusic, Thomas E.; Stanojev, Boris J.; Dehoyos, Amado; Raitses, Yevgeny; Smirnov, Artem; Fisch, Nathaniel J.

2007-01-01

Recent mission studies have shown that a Hall thruster which operates at relatively constant thrust efficiency (45-55%) over a broad power range (300W - 3kW) is enabling for deep space science missions when compared with slate-of-the-art ion thrusters. While conventional (annular) Hall thrusters can operate at high thrust efficiency at kW power levels, it is difficult to construct one that operates over a broad power envelope down to 0 (100 W) while maintaining relatively high efficiency. In this note we report the measured performance (I(sub sp), thrust and efficiency) of a cylindrical Hall thruster operating at 0 (100 W) input power.

One-Dimensional Analysis of Hall Thruster Operating Modes

DTIC Science & Technology

2001-08-01

Hall thruster structure with no screens or other control surfaces makes it difficult to understand the interrelationships which, in the end, localize and shape the various plasma regions existing in the accelerating channel. Since the radial magnetic field is usually shaped with a peak near the channel exit, the plasma structure has often been explained as simply a reflection of the magnetic field distribution. However, this is inadequate to explain the plasma dynamics inside the accelerating channel. We develop a macroscopic model gathering reliability and clarity.

Plasma Potential and Langmuir Probe Measurements in the Near-field Plume of the NASA 300M Hall Thruster

NASA Technical Reports Server (NTRS)

Herman, Daniel A; Shastry, Rohit; Huang, Wensheng; Soulas, George C.; KamHawi, Hani

2012-01-01

In order to aid in the design of high-power Hall thrusters and provide experimental validation for existing modeling efforts, plasma potential and Langmuir probe measurements were performed in the near-field plume of the NASA 300M Hall thruster. A probe array consisting of a Faraday probe, Langmuir probe, and emissive probe was used to interrogate the plume from approximately 0.1 - 2.0 DT,m downstream of the thruster exit plane at four operating conditions: 300 V, 400 V, and 500 V at 20 kW as well as 300 V at 10 kW. Results show that the acceleration zone and high-temperature region were contained within 0.3 DT,m from the exit plane at all operating conditions. Isothermal lines were shown to strongly follow magnetic field lines in the nearfield, with maximum temperatures ranging from 19 - 27 eV. The electron temperature spatial distribution created large drops in measured floating potentials in front of the magnetic pole surfaces where the plasma density was small, which suggests strong sheaths at these surfaces. The data taken have provided valuable information for future design and modeling validation, and complements ongoing internal measurement efforts on the NASA 300 M.

Performance Characteristics of a 5 kW Laboratory Hall Thruster

DTIC Science & Technology

1996-07-01

Characteristics of a 5 kW Laboratory Hall Thruster James M. Haas’, Frank S. Gulczinski III%, and Alec D. Gallimoret Plasmadynamics and Electric Propulsion...the information learned from the study of this thruster applicable to the understanding of its commercial counterparts. INTRODUCTION Hall thrusters are...few in number at this time; and those that do exist are intended primarily Current generation Hall thruster research has for flight qualification

Investigation of excited states populations density of Hall thruster plasma in three dimensions by laser-induced fluorescence spectroscopy

NASA Astrophysics Data System (ADS)

Krivoruchko, D. D.; Skrylev, A. V.

2018-01-01

The article deals with investigation of the excited states populations distribution of a low-temperature xenon plasma in the thruster with closed electron drift at 300 W operating conditions were investigated by laser-induced fluorescence (LIF) over the 350-1100 nm range. Seven xenon ions (Xe II) transitions were analyzed, while for neutral atoms (Xe I) just three transitions were explored, since the majority of Xe I emission falls into the ultraviolet or infrared part of the spectrum and are difficult to measure. The necessary spontaneous emission probabilities (Einstein coefficients) were calculated. Measurements of the excited state distribution were made for points (volume of about 12 mm3) all over the plane perpendicular to thruster axis in four positions on it (5, 10, 50 and 100 mm). Measured LIF signal intensity have differences for each location of researched point (due to anisotropy of thruster plume), however the structure of states populations distribution persisted at plume and is violated at the thruster exit plane and cathode area. Measured distributions show that for describing plasma of Hall thruster one needs to use a multilevel kinetic model, classic model can be used just for far plume region or for specific electron transitions.

Introduction of Shear-Based Transport Mechanisms in Radial-Axial Hybrid Hall Thruster Simulations

NASA Astrophysics Data System (ADS)

Scharfe, Michelle; Gascon, Nicolas; Scharfe, David; Cappelli, Mark; Fernandez, Eduardo

2007-11-01

Electron diffusion across magnetic field lines in Hall effect thrusters is experimentally observed to be higher than predicted by classical diffusion theory. Motivated by theoretical work for fusion applications and experimental measurements of Hall thrusters, numerical models for the electron transport are implemented in radial-axial hybrid simulations in order to compute the electron mobility using simulated plasma properties and fitting parameters. These models relate the cross-field transport to the imposed magnetic field distribution through shear suppression of turbulence-enhanced transport. While azimuthal waves likely enhance cross field mobility, axial shear in the electron fluid may reduce transport due to a reduction in turbulence amplitudes and modification of phase shifts between fluctuating properties. The sensitivity of the simulation results to the fitting parameters is evaluated and an examination is made of the transportability of these parameters to several Hall thruster devices.

Nonlinear ion dynamics in Hall thruster plasma source by ion transit-time instability

NASA Astrophysics Data System (ADS)

Lim, Youbong; Choe, Wonho; Mazouffre, Stéphane; Park, Jae Sun; Kim, Holak; Seon, Jongho; Garrigues, L.

2017-03-01

High-energy tail formation in an ion energy distribution function (IEDF) is explained in a Hall thruster plasma with the stationary crossed electric and magnetic fields whose discharge current is oscillated at the ion transit-time scale with a frequency of 360 kHz. Among ions in different charge states, singly charged Xe ions (Xe+) have an IEDF that is significantly broadened and shifted toward the high-energy side, which contributes to tail formation in the entire IEDF. Analytical and numerical investigations confirm that the IEDF tail is due to nonlinear ion dynamics in the ion transit-time oscillation.

Performance and Thermal Characterization of the NASA-300MS 20 kW Hall Effect Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Shastry, Rohit; Soulas, George; Smith, Timothy; Mikellides, Ioannis; Hofer, Richard

2013-01-01

NASA's Space Technology Mission Directorate is sponsoring the development of a high fidelity 15 kW-class long-life high performance Hall thruster for candidate NASA technology demonstration missions. An essential element of the development process is demonstration that incorporation of magnetic shielding on a 20 kW-class Hall thruster will yield significant improvements in the throughput capability of the thruster without any significant reduction in thruster performance. As such, NASA Glenn Research Center and the Jet Propulsion Laboratory collaborated on modifying the NASA-300M 20 kW Hall thruster to improve its propellant throughput capability. JPL and NASA Glenn researchers performed plasma numerical simulations with JPL's Hall2De and a commercially available magnetic modeling code that indicated significant enhancement in the throughput capability of the NASA-300M can be attained by modifying the thruster's magnetic circuit. This led to modifying the NASA-300M magnetic topology to a magnetically shielded topology. This paper presents performance evaluation results of the two NASA-300M magnetically shielded thruster configurations, designated 300MS and 300MS-2. The 300MS and 300MS-2 were operated at power levels between 2.5 and 20 kW at discharge voltages between 200 and 700 V. Discharge channel deposition from back-sputtered facility wall flux, and plasma potential and electron temperature measurements made on the inner and outer discharge channel surfaces confirmed that magnetic shielding was achieved. Peak total thrust efficiency of 64% and total specific impulse of 3,050 sec were demonstrated with the 300MS-2 at 20 kW. Thermal characterization results indicate that the boron nitride discharge chamber walls temperatures are approximately 100 C lower for the 300MS when compared to the NASA- 300M at the same thruster operating discharge power.

Complementary Density Measurements for the 200W Busek Hall Thruster (PREPRINT)

DTIC Science & Technology

2006-07-12

Hall thruster are presented. Both a Faraday probe and microwave interferometry system are used to examine the density distribution of the thruster plasma at regular spatial intervals. Both experiments are performed in situ under the same conditions. The resulting density distributions obtained from both experiments are presented. Advantages and uncertainties of both methods are presented, as well as how comparison between the two data sets can account for the uncertainties of each method

High Performance Power Module for Hall Effect Thrusters

NASA Technical Reports Server (NTRS)

Pinero, Luis R.; Peterson, Peter Y.; Bowers, Glen E.

2002-01-01

Previous efforts to develop power electronics for Hall thruster systems have targeted the 1 to 5 kW power range and an output voltage of approximately 300 V. New Hall thrusters are being developed for higher power, higher specific impulse, and multi-mode operation. These thrusters require up to 50 kW of power and a discharge voltage in excess of 600 V. Modular power supplies can process more power with higher efficiency at the expense of complexity. A 1 kW discharge power module was designed, built and integrated with a Hall thruster. The breadboard module has a power conversion efficiency in excess of 96 percent and weighs only 0.765 kg. This module will be used to develop a kW, multi-kW, and high voltage power processors.

Throttling Impacts on Hall Thruster Performance, Erosion, and Qualification for NASA Science Missions

NASA Technical Reports Server (NTRS)

Dankanich, John W.; DeHoyos, Amado

2007-01-01

With the SMART-1, Department of Defense, and commercial industry successes in Hall thruster technologies, NASA has started considering Hall thrusters for science missions. The recent Discovery proposals included a Hall thruster science mission and the In-Space Propulsion Project is investing in Hall thruster technologies. As the confidence in Hall thrusters improve, ambitious multi-thruster missions are being considered. Science missions often require large throttling ranges due to the 1/r(sup 2) power drop-off from the sun. Deep throttling of Hall thrusters will impact the overall system performance. Also, Hall thrusters can be throttled with both current and voltage, impacting erosion rates and performance. Last, electric propulsion thruster lifetime qualification has previously been conducted with long duration full power tests. Full power tests may not be appropriate for NASA science missions, and a combination of lifetime testing at various power levels with sufficient analysis is recommended. Analyses of various science missions and throttling schemes using the Aerojet BPT-4000 and NASA 103M HiVHAC thruster are presented.

Coupling intensity between discharge and magnetic circuit in Hall thrusters

NASA Astrophysics Data System (ADS)

Wei, Liqiu; Yang, Xinyong; Ding, Yongjie; Yu, Daren; Zhang, Chaohai

2017-03-01

Coupling oscillation is a newly discovered plasma oscillation mode that utilizes the coupling between the discharge circuit and magnetic circuit, whose oscillation frequency spectrum ranges from several kilohertz to megahertz. The coupling coefficient parameter represents the intensity of coupling between the discharge and magnetic circuits. According to previous studies, the coupling coefficient is related to the material and the cross-sectional area of the magnetic coils, and the magnetic circuit of the Hall thruster. However, in our recent study on coupling oscillations, it was found that the Hall current equivalent position and radius have important effects on the coupling intensity between the discharge and magnetic circuits. This causes a difference in the coupling coefficient for different operating conditions of Hall thrusters. Through non-intrusive methods for measuring the Hall current equivalent radius and the axial position, it is found that with an increase in the discharge voltage and magnetic field intensity, the Hall current equivalent radius increases and its axial position moves towards the exit plane. Thus, both the coupling coefficient and the coupling intensity between the discharge and magnetic circuits increase. Contribution to the Topical Issue "Physics of Ion Beam Sources", edited by Holger Kersten and Horst Neumann.

Study on the influences of ionization region material arrangement on Hall thruster channel discharge characteristics

NASA Astrophysics Data System (ADS)

Xiang, HU; Ping, DUAN; Jilei, SONG; Wenqing, LI; Long, CHEN; Xingyu, BIAN

2018-02-01

There exists strong interaction between the plasma and channel wall in the Hall thruster, which greatly affects the discharge performance of the thruster. In this paper, a two-dimensional physical model is established based on the actual size of an Aton P70 Hall thruster discharge channel. The particle-in-cell simulation method is applied to study the influences of segmented low emissive graphite electrode biased with anode voltage on the discharge characteristics of the Hall thruster channel. The influences of segmented electrode placed at the ionization region on electric potential, ion number density, electron temperature, ionization rate, discharge current and specific impulse are discussed. The results show that, when segmented electrode is placed at the ionization region, the axial length of the acceleration region is shortened, the equipotential lines tend to be vertical with wall at the acceleration region, thus radial velocity of ions is reduced along with the wall corrosion. The axial position of the maximal electron temperature moves towards the exit with the expansion of ionization region. Furthermore, the electron-wall collision frequency and ionization rate also increase, the discharge current decreases and the specific impulse of the Hall thruster is slightly enhanced.

Particle-in-cell numerical simulations of a cylindrical Hall thruster with permanent magnets

NASA Astrophysics Data System (ADS)

Miranda, Rodrigo A.; Martins, Alexandre A.; Ferreira, José L.

2017-10-01

The cylindrical Hall thruster (CHT) is a propulsion device that offers high propellant utilization and performance at smaller dimensions and lower power levels than traditional Hall thrusters. In this paper we present first results of a numerical model of a CHT. This model solves particle and field dynamics self-consistently using a particle-in-cell approach. We describe a number of techniques applied to reduce the execution time of the numerical simulations. The specific impulse and thrust computed from our simulations are in agreement with laboratory experiments. This simplified model will allow for a detailed analysis of different thruster operational parameters and obtain an optimal configuration to be implemented at the Plasma Physics Laboratory at the University of Brasília.

Hybrid-PIC Modeling of a High-Voltage, High-Specific-Impulse Hall Thruster

NASA Technical Reports Server (NTRS)

Smith, Brandon D.; Boyd, Iain D.; Kamhawi, Hani; Huang, Wensheng

2013-01-01

The primary life-limiting mechanism of Hall thrusters is the sputter erosion of the discharge channel walls by high-energy propellant ions. Because of the difficulty involved in characterizing this erosion experimentally, many past efforts have focused on numerical modeling to predict erosion rates and thruster lifespan, but those analyses were limited to Hall thrusters operating in the 200-400V discharge voltage range. Thrusters operating at higher discharge voltages (V(sub d) >= 500 V) present an erosion environment that may differ greatly from that of the lower-voltage thrusters modeled in the past. In this work, HPHall, a well-established hybrid-PIC code, is used to simulate NASA's High-Voltage Hall Accelerator (HiVHAc) at discharge voltages of 300, 400, and 500V as a first step towards modeling the discharge channel erosion. It is found that the model accurately predicts the thruster performance at all operating conditions to within 6%. The model predicts a normalized plasma potential profile that is consistent between all three operating points, with the acceleration zone appearing in the same approximate location. The expected trend of increasing electron temperature with increasing discharge voltage is observed. An analysis of the discharge current oscillations shows that the model predicts oscillations that are much greater in amplitude than those measured experimentally at all operating points, suggesting that the differences in oscillation amplitude are not strongly associated with discharge voltage.

Numerical Study of Current Driven Instabilities and Anomalous Electron Transport in Hall-effect Thrusters

NASA Astrophysics Data System (ADS)

Tran, Jonathan

Plasma turbulence and the resulting anomalous electron transport due to azimuthal current driven instabilities in Hall-effect thrusters is a promising candidate for developing predictive models for the observed anomalous transport. A theory for anomalous electron transport and current driven instabilities has been recently studied by [Lafluer et al., 2016a]. Due to the extreme cost of fully resolving the Debye length and plasma frequency, hybrid plasma simulations utilizing kinetic ions and quasi-steady state fluid electrons have long been the principle workhorse methodology for Hall-effect thruster modeling. Using a reduced dimension particle in cell simulation implemented in the Thermophysics Universal Research Framework developed by the Air Force Research Lab, we show collective electron-wave scattering due to large amplitude azimuthal fluctuations of the electric field and the plasma density. These high-frequency and short wavelength fluctuations can lead to an effective cross-field mobility many orders of magnitude larger than what is expected from classical electron-neutral momentum collisions in the low neutral density regime. We further adapt the previous study by [Lampe et al., 1971] and [Stringer, 1964] for related current driven instabilities to electric propulsion relevant mass ratios and conditions. Finally, we conduct a preliminary study of resolving this instability with a modified hybrid simulation with the hope of integration with established hybrid Hall-effect thruster simulations.

Compact high-speed reciprocating probe system for measurements in a Hall thruster discharge and plume.

PubMed

Dannenmayer, K; Mazouffre, S

2012-12-01

A compact high-speed reciprocating probe system has been developed in order to perform measurements of the plasma parameters by means of electrostatic probes in the discharge and the plume of a Hall thruster. The system is based on a piezoelectric linear drive that can achieve a speed of up to 350 mm/s over a travel range of 90 mm. Due to the high velocity of the linear drive the probe can be rapidly moved in and out the measurement region in order to minimize perturbation of the thruster discharge due to sputtering of probe material. To demonstrate the impact of the new system, a heated emissive probe, installed on the high-speed translation stage, was used to measure the plasma potential and the electron temperature in the near-field plume of a low power Hall thruster.

Hollow Cathode Assembly Development for the HERMeS Hall Thruster

NASA Technical Reports Server (NTRS)

Sarver-Verhey, Timothy R.; Kamhawi, Hani; Goebel, Dan M.; Polk, James E.; Peterson, Peter Y.; Robinson, Dale A.

2016-01-01

To support the operation of the HERMeS 12.5 kW Hall Thruster for NASA's Asteroid Redirect Robotic Mission, hollow cathodes using emitters based on barium oxide impregnate and lanthanum hexaboride are being evaluated through wear-testing, performance characterization, plasma modeling, and review of integration requirements. This presentation will present the development approach used to assess the cathode emitter options. A 2,000-hour wear-test of development model Barium Oxide (BaO) hollow cathode is being performed as part of the development plan. Specifically this test is to identify potential impacts cathode emitter life during operation in the HERMeS thruster. The cathode was operated with a magnetic field-equipped anode that simulates the HERMeS hall thruster operating environment. Cathode discharge performance has been stable with the device accumulating 743 hours at the time of this report. Observed voltage changes are attributed to keeper surface condition changes during testing. Cathode behavior during characterization sweeps exhibited stable behavior, including cathode temperature. The details of the cathode assembly operation of the wear-test will be presented.

A study of cylindrical Hall thruster for low power space applications

DOE Office of Scientific and Technical Information (OSTI.GOV)

Y. Raitses; N.J. Fisch; K.M. Ertmer

2000-07-27

A 9 cm cylindrical thruster with a ceramic channel exhibited performance comparable to the state-of-the-art Hall thrusters at low and moderate power levels. Significantly, its operation is not accompanied by large amplitude discharge low frequency oscillations. Preliminary experiments on a 2 cm cylindrical thruster suggest the possibility of a high performance micro Hall thruster.

Iodine Hall Thruster

NASA Technical Reports Server (NTRS)

Szabo, James

2015-01-01

Iodine enables dramatic mass and cost savings for lunar and Mars cargo missions, including Earth escape and near-Earth space maneuvers. The demonstrated throttling ability of iodine is important for a singular thruster that might be called upon to propel a spacecraft from Earth to Mars or Venus. The ability to throttle efficiently is even more important for missions beyond Mars. In the Phase I project, Busek Company, Inc., tested an existing Hall thruster, the BHT-8000, on iodine propellant. The thruster was fed by a high-flow iodine feed system and supported by an existing Busek hollow cathode flowing xenon gas. The Phase I propellant feed system was evolved from a previously demonstrated laboratory feed system. Throttling of the thruster between 2 and 11 kW at 200 to 600 V was demonstrated. Testing showed that the efficiency of iodine fueled BHT-8000 is the same as with xenon, with iodine delivering a slightly higher thrust-to-power (T/P) ratio. In Phase II, a complete iodine-fueled system was developed, including the thruster, hollow cathode, and iodine propellant feed system. The nominal power of the Phase II system is 8 kW; however, it can be deeply throttled as well as clustered to much higher power levels. The technology also can be scaled to greater than 100 kW per thruster to support megawatt-class missions. The target thruster efficiency for the full-scale system is 65 percent at high specific impulse (Isp) (approximately 3,000 s) and 60 percent at high thrust (Isp approximately 2,000 s).

Transition in discharge plasma of Hall thruster type in presence of secondary electron emissive surface

NASA Astrophysics Data System (ADS)

Schweigert, I. V.; Yadrenkin, M. A.; Fomichev, V. P.

2017-11-01

Modification of the sheath structure near the emissive plate placed in magnetized DC discharge plasma of Hall thruster type was studied in the experiment and in kinetic simulations. The plate is made from Al2O3 which has enhanced secondary electron emission yield. The energetic electrons emitted by heated cathode provide the volume ionization and the secondary electron emission from the plate. An increase of the electron beam energy leads to an increase of the secondary electron generation, which initiates the transition in sheath structure over the emissive plate.

Very-Near-Field Plume Model of a Hall Thruster

DTIC Science & Technology

2003-07-20

UNCLASSIFIED Defense Technical Information Center Compilation Part Notice ADP014988 TITLE: Very-Near-Field Plume Model of a Hall Thruster DISTRIBUTION...numbers comprise the compilation report: ADP014936 thru ADP015049 UNCLASSIFIED am 46 Very-Near-Field Plume Model of a Hall Thruster F. Taccogna’, S. LongoŖ

Study of Energy Loss Mechanisms in the BPT-4000 Hall Thruster

DTIC Science & Technology

2003-06-30

Aerojet has developed a high performance multi-mode flightweight Hall thruster for orbit raising and stationkeeping on geo-synchronous satellites. In...order to further understand and improve upon the performance of this state of the art Hall thruster and other next generation thrusters being planned

Kinetic particle simulation of discharge and wall erosion of a Hall thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cho, Shinatora; Komurasaki, Kimiya; Arakawa, Yoshihiro

2013-06-15

The primary lifetime limiting factor of Hall thrusters is the wall erosion caused by the ion induced sputtering, which is predominated by dielectric wall sheath and pre-sheath. However, so far only fluid or hybrid simulation models were applied to wall erosion and lifetime studies in which this non-quasi-neutral and non-equilibrium area cannot be treated directly. Thus, in this study, a 2D fully kinetic particle-in-cell model was presented for Hall thruster discharge and lifetime simulation. Because the fully kinetic lifetime simulation was yet to be achieved so far due to the high computational cost, the semi-implicit field solver and the techniquemore » of mass ratio manipulation was employed to accelerate the computation. However, other artificial manipulations like permittivity or geometry scaling were not used in order to avoid unrecoverable change of physics. Additionally, a new physics recovering model for the mass ratio was presented for better preservation of electron mobility at the weakly magnetically confined plasma region. The validity of the presented model was examined by various parametric studies, and the thrust performance and wall erosion rate of a laboratory model magnetic layer type Hall thruster was modeled for different operation conditions. The simulation results successfully reproduced the measurement results with typically less than 10% discrepancy without tuning any numerical parameters. It is also shown that the computational cost was reduced to the level that the Hall thruster fully kinetic lifetime simulation is feasible.« less

Design and utilization of a top hat analyzer for Hall thruster plume diagnostics

NASA Astrophysics Data System (ADS)

Victor, Allen Leoraj

Electric propulsion offers new capabilities for ambitious space missions of the future. However, coating, uneven heating, and the charging of spacecraft components have impeded the integration of Hall thrusters for space missions and encouraged plume diagnostics of the thruster plasma environment. Plume diagnostics are also important for the inference of thruster performance through plume properties downstream of the engine. While the top hat analyzer has been available for low-density space plasma diagnostics for over twenty years, the use of this instrument for plasma thruster plume diagnostics has been nonexistent. This thesis describes the development of a new diagnostics tool, the Top Hat Electric Propulsion Plume Analyzer (TOPAZ), which provides unprecedented insight into the physical mechanisms that govern the performance of Hall thrusters. Novel measurements conducted by TOPAZ on the BHT-600 Hall thruster cluster yielded interesting and undocumented phenomena in the far-field plume. SIMION, a commercial ion optics program, was used to design TOPAZ and estimate the energy and angular resolutions as well as the instrument's sensitivity and plate-voltage relationships. TOPAZ was experimentally characterized through an ion beam facility operating on air, xenon, and krypton gases. Measurements on the BHT-600 cluster indicated lower-energy ions emanated from positions closer to the cathode while higher-energy ions were measured from along the discharge channel centerlines. Low-energy ions were also measured from behind the cathodes only during cluster operation. Charge-exchange and ionization outside the primary acceleration region are believed to be the cause of the variance in the energy distributions. Cross pollination of the cathode plume with the opposite thruster is argued to create low-energy ions which emanate from behind the cathode. Time-of-flight measurements through TOPAZ allowed for charge-state and species fraction discriminations as functions of

An End-to-End Model of a Hall Thruster

DTIC Science & Technology

2000-09-01

and deposition of sputtered material, simulation of the operator of a Hall Thruster in a vacuum tank and the extension to the near-plume of a...sophisticated Hall thruster transient hybrid PlC model which had been previously used only to describe the internal flow. The first two items have been

Hybrid-PIC Modeling of the Transport of Atomic Boron in a Hall Thruster

NASA Technical Reports Server (NTRS)

Smith, Brandon D.; Boyd, Iaian D.; Kamhawi, Hani

2015-01-01

Computational analysis of the transport of boron eroded from the walls of a Hall thruster is performed by implementing sputter yields of hexagonal boron nitride and velocity distribution functions of boron within the hybrid-PIC model HPHall. The model is applied to simulate NASA's HiVHAc Hall thruster at a discharge voltage of 500V and discharge powers of 1-3 kW. The number densities of ground- and 4P-state boron are computed. The density of ground-state boron is shown to be a factor of about 30 less than the plasma density. The density of the excited state is shown to be about three orders of magnitude less than that of the ground state, indicating that electron impact excitation does not significantly affect the density of ground-state boron in the discharge channel or near-field plume of a Hall thruster. Comparing the rates of excitation and ionization suggests that ionization has a greater influence on the density of ground-state boron, but is still negligible. The ground-state boron density is then integrated and compared to cavity ring-down spectroscopy (CRDS) measurements for each operating point. The simulation results show good agreement with the measurements for all operating points and provide evidence in support of CRDS as a tool for measuring Hall thruster erosion in situ.

High-Power Hall Thruster Technology Evaluated for Primary Propulsion Applications

NASA Technical Reports Server (NTRS)

Manzella, David H.; Jankovsky, Robert S.; Hofer, Richard R.

2003-01-01

High-power electric propulsion systems have been shown to be enabling for a number of NASA concepts, including piloted missions to Mars and Earth-orbiting solar electric power generation for terrestrial use (refs. 1 and 2). These types of missions require moderate transfer times and sizable thrust levels, resulting in an optimized propulsion system with greater specific impulse than conventional chemical systems and greater thrust than ion thruster systems. Hall thruster technology will offer a favorable combination of performance, reliability, and lifetime for such applications if input power can be scaled by more than an order of magnitude from the kilowatt level of the current state-of-the-art systems. As a result, the NASA Glenn Research Center conducted strategic technology research and development into high-power Hall thruster technology. During program year 2002, an in-house fabricated thruster, designated the NASA-457M, was experimentally evaluated at input powers up to 72 kW. These tests demonstrated the efficacy of scaling Hall thrusters to high power suitable for a range of future missions. Thrust up to nearly 3 N was measured. Discharge specific impulses ranged from 1750 to 3250 sec, with discharge efficiencies between 46 and 65 percent. This thruster is the highest power, highest thrust Hall thruster ever tested.

Low-Mass, Low-Power Hall Thruster System

NASA Technical Reports Server (NTRS)

Pote, Bruce

2015-01-01

NASA is developing an electric propulsion system capable of producing 20 mN thrust with input power up to 1,000 W and specific impulse ranging from 1,600 to 3,500 seconds. The key technical challenge is the target mass of 1 kg for the thruster and 2 kg for the power processing unit (PPU). In Phase I, Busek Company, Inc., developed an overall subsystem design for the thruster/cathode, PPU, and xenon feed system. This project demonstrated the feasibility of a low-mass power processing architecture that replaces four of the DC-DC converters of a typical PPU with a single multifunctional converter and a low-mass Hall thruster design employing permanent magnets. In Phase II, the team developed an engineering prototype model of its low-mass BHT-600 Hall thruster system, with the primary focus on the low-mass PPU and thruster. The goal was to develop an electric propulsion thruster with the appropriate specific impulse and propellant throughput to enable radioisotope electric propulsion (REP). This is important because REP offers the benefits of nuclear electric propulsion without the need for an excessively large spacecraft and power system.

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

NASA Technical Reports Server (NTRS)

Jacobson, David (Technical Monitor); Roy, Subrata

2004-01-01

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.

Plasma Oscillation Characterization of NASA's HERMeS Hall Thruster via High Speed Imaging

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Haag, Thomas W.

2016-01-01

The performance and facility effect characterization tests of NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding had been completed. As a part of these tests, three plasma oscillation characterization studies were performed to help determine operation settings and quantify margins. The studies included the magnetic field strength variation study, background pressure effect study, and cathode flow fraction study. Separate high-speed videos of the thruster including the cathode and of only the cathode were recorded. Breathing mode at 10-15 kHz and cathode gradient-driven mode at 60-75 kHz were observed. An additional high frequency (40-70 kHz) global oscillation mode with sinusoidal probability distribution function was identified.

An evaluation of krypton propellant in Hall thrusters

NASA Astrophysics Data System (ADS)

Linnell, Jesse Allen

Due to its high specific impulse and low price, krypton has long sparked interest as an alternate Hall thruster propellant. Unfortunately at the moment, krypton's relatively poor performance precludes it as a legitimate option. This thesis presents a detailed investigation into krypton operation in Hall thrusters. These findings suggest that the performance gap can be decreased to 4% and krypton can finally become a realistic propellant option. Although krypton has demonstrated superior specific impulse, the xenon-krypton absolute efficiency gap ranges between 2 and 15%. A phenomenological performance model indicates that the main contributors to the efficiency gap are propellant utilization and beam divergence. Propellant utilization and beam divergence have relative efficiency deficits of 5 and 8%, respectively. A detailed characterization of internal phenomena is conducted to better understand the xenon-krypton efficiency gap. Krypton's large beam divergence is found to be related to a defocusing equipotential structure and a weaker magnetic field topology. Ionization processes are shown to be linked to the Hall current, the magnetic mirror topology, and the perpendicular gradient of the magnetic field. Several thruster design and operational suggestions are made to optimize krypton efficiency. Krypton performance is optimized for discharge voltages above 500 V and flow rates corresponding to an a greater than 0.015 mg/(mm-s), where alpha is a function of flow rate and discharge channel dimensions (alpha = m˙alphab/Ach). Performance can be further improved by increasing channel length or decreasing channel width for a given flow rate. Also, several magnetic field design suggestions are made to enhance ionization and beam focusing. Several findings are presented that improve the understanding of general Hall thruster physics. Excellent agreement is shown between equipotential lines and magnetic field lines. The trim coil is shown to enhance beam focusing

Study and Developement of Compact Permanent Magnet Hall Thrusters for Future Brazillian Space Missions

NASA Astrophysics Data System (ADS)

Ferreira, Jose Leonardo; Martins, Alexandre; Cerda, Rodrigo

2016-07-01

The Plasma Physics Laboratory of UnB has been developing a Permanent Magnet Hall Thruster (PHALL) for the UNIESPAÇO program, part of the Space Activities Program conducted by AEB- The Brazillian Space Agency since 2004. Electric propulsion is now a very successful method for primary and secondary propulsion systems. It is essential for several existing geostationary satellite station keeping systems and for deep space long duration solar system missions, where the thrusting system can be designed to be used on orbit transfer maneuvering and/or for satellite attitude control in long term space missions. Applications of compact versions of Permanent Magnet Hall Thrusters on future brazillian space missions are needed and foreseen for the coming years beginning with the use of small divergent cusp field (DCFH) Hall Thrusters type on CUBESATS ( 5-10 kg , 1W-5 W power consumption) and on Micro satellites ( 50- 100 kg, 10W-100W). Brazillian (AEB) and German (DLR) space agencies and research institutions are developing a new rocket dedicated to small satellite launching. The VLM- Microsatellite Launch Vehicle. The development of PHALL compact versions can also be important for the recently proposed SBG system, a future brazillian geostationary satellite system that is already been developed by an international consortium of brazillian and foreign space industries. The exploration of small bodies in the Solar System with spacecraft has been done by several countries with increasing frequency in these past twenty five years. Since their historical beginning on the sixties, most of the Solar System missions were based on gravity assisted trajectories very much depended on planet orbit positioning relative to the Sun and the Earth. The consequence was always the narrowing of the mission launch window. Today, the need for Solar System icy bodies in situ exploration requires less dependence on gravity assisted maneuvering and new high precision low thrust navigation methods

First Firing of a 100-kW Nested-Channel Hall Thruster

DTIC Science & Technology

2013-09-01

Technical Paper 3. DATES COVERED (From - To) September 2013- December 2013 4. TITLE AND SUBTITLE First Firing of a 100-kW Nested-Channel Hall Thruster 5a...STATEMENT A: Approved for public release; distribution unlimited. 1 First Firing of a 100-kW Nested-channel Hall Thruster IEPC-2013-394...converting electrical power to directed kinetic power I. Introduction ESTING the channels of Hall thrusters has proven to be a viable method to increase

Low-Cost, High-Performance Hall Thruster Support System

NASA Technical Reports Server (NTRS)

Hesterman, Bryce

2015-01-01

Colorado Power Electronics (CPE) has built an innovative modular PPU for Hall thrusters, including discharge, magnet, heater and keeper supplies, and an interface module. This high-performance PPU offers resonant circuit topologies, magnetics design, modularity, and a stable and sustained operation during severe Hall effect thruster current oscillations. Laboratory testing has demonstrated discharge module efficiency of 96 percent, which is considerably higher than current state of the art.

Ion Engine and Hall Thruster Development at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Domonkos, Matthew T.; Patterson, Michael J.; Jankovsky, Robert S.

2002-01-01

NASA's Glenn Research Center has been selected to lead development of NASA's Evolutionary Xenon Thruster (NEXT) system. The central feature of the NEXT system is an electric propulsion thruster (EPT) that inherits the knowledge gained through the NSTAR thruster that successfully propelled Deep Space 1 to asteroid Braille and comet Borrelly, while significantly increasing the thruster power level and making improvements in performance parameters associated with NSTAR. The EPT concept under development has a 40 cm beam diameter, twice the effective area of the Deep-Space 1 thruster, while maintaining a relatively-small volume. It incorporates mechanical features and operating conditions to maximize the design heritage established by the flight NSTAR 30 cm engine, while incorporating new technology where warranted to extend the power and throughput capability. The NASA Hall thruster program currently supports a number of tasks related to high power thruster development for a number of customers including the Energetics Program (formerly called the Space-based Program), the Space Solar Power Program, and the In-space Propulsion Program. In program year 2002, two tasks were central to the NASA Hall thruster program: 1.) the development of a laboratory Hall thruster capable of providing high thrust at high power; 2.) investigations into operation of Hall thrusters at high specific impulse. In addition to these two primary thruster development activities, there are a number of other on-going activities supported by the NASA Hall thruster program, These additional activities are related to issues such as thruster lifetime and spacecraft integration.

Effect of azimuthal diversion rail on an ATON-type Hall thruster

NASA Astrophysics Data System (ADS)

Xu, Zhang; Liqiu, Wei; Liang, Han; Yongjie, Ding; Daren, Yu

2017-03-01

A newly designed azimuthal diversion rail (ADR) is studied and used to enhance the ionization process in an ATON-type Hall thruster. The diversion rail efficiently reduces the neutral flow axial velocity, and hence, increases the resistance time of atoms in the discharge channel of the Hall thruster. Thrust performances, in terms of thrust, anode efficiency and ion beam divergence, are found to be improved because of the application of the diversion rail, especially at low mass flow rate conditions. Experiment results reveal that the ADR increases the mass utilization under insufficient mass flow rate operating conditions. The design of the ADR broadens the efficient operating range of Hall thrusters and has significant contribution to multi-mode Hall thruster development.

Mission and System Advantages of Iodine Hall Thrusters

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Szabo, James; Pote, Bruce; Oleson, Steve; Kamhawi, Hani

2014-01-01

The exploration of alternative propellants for Hall thrusters continues to be of interest to the community. Investments have been made and continue for the maturation of iodine based Hall thrusters. Iodine testing has shown comparable performance to xenon. However, iodine has a higher storage density and resulting higher ?V capability for volume constrained systems. Iodine's vapor pressure is low enough to permit low-pressure storage, but high enough to minimize potential adverse spacecraft-thruster interactions. The low vapor pressure also means that iodine does not condense inside the thruster at ordinary operating temperatures. Iodine is safe, it stores at sub-atmospheric pressure, and can be stored unregulated for years on end; whether on the ground or on orbit. Iodine fills a niche for both low power (<1kW) and high power (>10kW) electric propulsion regimes. A range of missions have been evaluated for direct comparison of Iodine and Xenon options. The results show advantages of iodine Hall systems for both small and microsatellite application and for very large exploration class missions.

Hall Thruster Thermal Modeling and Test Data Correlation

NASA Technical Reports Server (NTRS)

Myers, James

2016-01-01

HERMeS - Hall Effect Rocket with Magnetic Shielding. Developed through a joint effort by NASA/GRC and the Jet Propulsion Laboratory (JPL). Design goals: High power (12.5 kW) high Isp (3000 sec), high efficiency (> 60%), high throughput (10,000 kg), reduced plasma erosion and increased life (5 yrs) to support Asteroid Redirect Robotic Mission (ARRM). Further details see "Performance, Facility Pressure Effects and Stability Characterization Tests of NASAs HERMeS Thruster" by H. Kamhawi and team. Hall Thrusters (HT) inherently operate at elevated temperatures approx. 600 C (or more). Due to electric magnetic (E x B) fields used to ionize and accelerate propellant gas particles (i.e., plasma). Cooling is largely limited to radiation in vacuum environment.Thus the hardware components must withstand large start-up delta-T's. HT's are constructed of multiple materials; assorted metals, non-metals and ceramics for their required electrical and magnetic properties. To mitigate thermal stresses HT design must accommodate the differential thermal growth from a wide range of material Coef. of Thermal Expansion (CTEs). Prohibiting the use of some bolted/torqued interfaces.Commonly use spring loaded interfaces, particularly at the metal-to-ceramic interfaces to allow for slippage.However most component interfaces must also effectively conduct heat to the external surfaces for dissipation by radiation.Thus contact pressure and area are important.

Power Reduction of the Air-Breathing Hall-Effect Thruster

NASA Astrophysics Data System (ADS)

Kim, Sungrae

Electric propulsion system is spotlighted as the next generation space propulsion system due to its benefits; one of them is specific impulse. While there are a lot of types in electric propulsion system, Hall-Effect Thruster, one of electric propulsion system, has higher thrust-to-power ratio and requires fewer power supplies for operation in comparison to other electric propulsion systems, which means it is optimal for long space voyage. The usual propellant for Hall-Effect Thruster is Xenon and it is used to be stored in the tank, which may increase the weight of the thruster. Therefore, one theory that uses the ambient air as a propellant has been proposed and it is introduced as Air-Breathing Hall-Effect Thruster. Referring to the analysis on Air-Breathing Hall-Effect Thruster, the goal of this paper is to reduce the power of the thruster so that it can be applied to real mission such as satellite orbit adjustment. To reduce the power of the thruster, two assumptions are considered. First one is changing the altitude for the operation, while another one is assuming the alpha value that is electron density to ambient air density. With assumptions above, the analysis was done and the results are represented. The power could be decreased to 10s˜1000s with the assumptions. However, some parameters that do not satisfy the expectation, which would be the question for future work, and it will be introduced at the end of the thesis.

Hollow Cathode Assembly Development for the HERMeS Hall Thruster

NASA Technical Reports Server (NTRS)

Sarver-Verhey, Timothy R.; Kamhawi, Hani; Goebel, Dan M.; Polk, James E.; Peterson, Peter Y.; Robinson, Dale A.

2016-01-01

To support the operation of the HERMeS 12.5 kW Hall Thruster for NASA's Asteroid Redirect Robotic Mission, hollow cathodes using emitters based on barium oxide impregnate and lanthanum hexaboride are being evaluated through wear-testing, performance characterization, plasma modeling, and assessment of system implementation concerns. This paper will present the development approach used to assess the cathode emitter options. A 2,000-hour wear-test of development model barium-oxide-based (BaO) hollow cathode is being performed as part of the development plan. The cathode was operated with an anode that simulates the HERMeS hall thruster operating environment. Cathode discharge performance has been stable with the device accumulating 740 hours at the time of this report. Cathode operation (i.e. discharge voltage and orifice temperature) was repeatable during period variation of discharge current and flow rate. The details of the cathode assembly operation during the wear-test will be presented.

Modeling of anomalous electron mobility in Hall thrusters

DOE Office of Scientific and Technical Information (OSTI.GOV)

Koo, Justin W.; Boyd, Iain D.

Accurate modeling of the anomalous electron mobility is absolutely critical for successful simulation of Hall thrusters. In this work, existing computational models for the anomalous electron mobility are used to simulate the UM/AFRL P5 Hall thruster (a 5 kW laboratory model) in a two-dimensional axisymmetric hybrid particle-in-cell Monte Carlo collision code. Comparison to experimental results indicates that, while these computational models can be tuned to reproduce the correct thrust or discharge current, it is very difficult to match all integrated performance parameters (thrust, power, discharge current, etc.) simultaneously. Furthermore, multiple configurations of these computational models can produce reasonable integrated performancemore » parameters. A semiempirical electron mobility profile is constructed from a combination of internal experimental data and modeling assumptions. This semiempirical electron mobility profile is used in the code and results in more accurate simulation of both the integrated performance parameters and the mean potential profile of the thruster. Results indicate that the anomalous electron mobility, while absolutely necessary in the near-field region, provides a substantially smaller contribution to the total electron mobility in the high Hall current region near the thruster exit plane.« less

Iodine Plasma Species Measurements in a Hall Effect Thruster Plume

DTIC Science & Technology

2013-04-01

direction f = species fraction 0g = gravitational constant at Earth’s surface, 9.81 m/s 2 I = current, subscripts b for beam, c for cathode, d for...Hall effect thruster uses crossed electric and magnetic fields to generate and accelerate ions. The gas in the discharge is partially ionized, although...early 1960s.10 Ions are weakly magnetized and most are accelerated directly out of the channel, forming the ion beam. The bulk of the cathode

Segmented electrode hall thruster with reduced plume

DOEpatents

Fisch, Nathaniel J.; Raitses, Yevgeny

2004-08-17

An apparatus and method for thrusting plasma, utilizing a Hall thruster with segmented electrodes along the channel, which make the acceleration region as localized as possible. Also disclosed are methods of arranging the electrodes so as to minimize erosion and arcing. Also disclosed are methods of arranging the electrodes so as to produce a substantial reduction in plume divergence. The use of electrodes made of emissive material will reduce the radial potential drop within the channel, further decreasing the plume divergence. Also disclosed is a method of arranging and powering these electrodes so as to provide variable mode operation.

Hall-Effect Thruster Simulations with 2-D Electron Transport and Hydrodynamic Ions

NASA Technical Reports Server (NTRS)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard H.; Goebel, Dan M.

2009-01-01

A computational approach that has been used extensively in the last two decades for Hall thruster simulations is to solve a diffusion equation and energy conservation law for the electrons in a direction that is perpendicular to the magnetic field, and use discrete-particle methods for the heavy species. This "hybrid" approach has allowed for the capture of bulk plasma phenomena inside these thrusters within reasonable computational times. Regions of the thruster with complex magnetic field arrangements (such as those near eroded walls and magnets) and/or reduced Hall parameter (such as those near the anode and the cathode plume) challenge the validity of the quasi-one-dimensional assumption for the electrons. This paper reports on the development of a computer code that solves numerically the 2-D axisymmetric vector form of Ohm's law, with no assumptions regarding the rate of electron transport in the parallel and perpendicular directions. The numerical challenges related to the large disparity of the transport coefficients in the two directions are met by solving the equations in a computational mesh that is aligned with the magnetic field. The fully-2D approach allows for a large physical domain that extends more than five times the thruster channel length in the axial direction, and encompasses the cathode boundary. Ions are treated as an isothermal, cold (relative to the electrons) fluid, accounting for charge-exchange and multiple-ionization collisions in the momentum equations. A first series of simulations of two Hall thrusters, namely the BPT-4000 and a 6-kW laboratory thruster, quantifies the significance of ion diffusion in the anode region and the importance of the extended physical domain on studies related to the impact of the transport coefficients on the electron flow field.

Microwave Interferometry (90 GHz) for Hall Thruster Plume Density Characterization

DTIC Science & Technology

2005-06-01

Hall thruster . The interferometer has been modified to overcome initial difficulties encountered during the preliminary testing. The modifications include the ability to perform remote and automated calibrations as well as an aluminum enclosure to shield the interferometer from the Hall thruster plume. With these modifications, it will be possible to make unambiguous electron density measurements of the thruster plume as well as to rapidly and automatically calibrate the interferometer to eliminate the effects of signal drift. Due to the versatility

Low-Power Operation and Plasma Characterization of a Qualification Model SPT-140 Hall Thruster for NASA Science Missions

NASA Technical Reports Server (NTRS)

Garner, Charles E.; Jorns, Benjamin A.; van Derventer, Steven; Hofer, Richard R.; Rickard, Ryan; Liang, Raymond; Delgado, Jorge

2015-01-01

Hall thruster systems based on commercial product lines can potentially lead to lower cost electric propulsion (EP) systems for deep space science missions. A 4.5-kW SPT-140 Hall thruster presently under qualification testing by SSL leverages the substantial heritage of the SPT-100 being flown on Russian and US commercial satellites. The Jet Propulsion Laboratory is exploring the use of commercial EP systems, including the SPT-140, for deep space science missions, and initiated a program to evaluate the SPT-140 in the areas of low power operation and thruster operating life. A qualification model SPT-140 designated QM002 was evaluated for operation and plasma properties along channel centerline, from 4.5 kW to 0.8 kW. Additional testing was performed on a development model SPT-140 designated DM4 to evaluate operation with a Moog proportional flow control valve (PFCV). The PFCV was commanded by an SSL engineering model PPU-140 Power Processing Unit (PPU). Performance measurements on QM002 at 0.8 kW discharge power were 50 mN of thrust at a total specific impulse of 1250 s, a total thruster efficiency of 0.38, and discharge current oscillations of under 3% of the mean current. Steady-state operation at 0.8 kW was demonstrated during a 27 h firing. The SPT-140 DM4 was operated in closed-loop control of the discharge current with the PFCV and PPU over discharge power levels of 0.8-4.5 kW. QM002 and DM4 test data indicate that the SPT-140 design is a viable candidate for NASA missions requiring power throttling down to low thruster input power.

Plasma Potential and Langmuir Probe Measurements in the Near-field Plume of the NASA-457Mv2 Hall Thruster

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Huang, Wensheng; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2012-01-01

In order to further the design of future high-power Hall thrusters and provide experimental validation for ongoing modeling efforts, plasma potential and Langmuir probe measurements were performed on the 50-kW NASA-457Mv2. An electrostatic probe array comprised of a near-field Faraday probe, single Langmuir probe, and emissive probe was used to interrogate the near-field plume from approximately 0.1 - 2.0 mean thruster diameters downstream of the thruster exit plane at the following operating conditions: 300 V, 400 V and 500 V at 30 kW and 500 V at 50 kW. Results have shown that the acceleration zone is limited to within 0.4 mean thruster diameters of the exit plane while the high-temperature region is limited to 0.25 mean thruster diameters from the exit plane at all four operating conditions. Maximum plasma potentials in the near-field at 300 and 400 V were approximately 50 V with respect to cathode potential, while maximum electron temperatures varied from 24 - 32 eV, depending on operating condition. Isothermal lines at all operating conditions were found to strongly resemble the magnetic field topology in the high-temperature regions. This distribution was found to create regions of high temperature and low density near the magnetic poles, indicating strong, thick sheath formation along these surfaces. The data taken from this study are considered valuable for future design as well as modeling validation.

Preliminary Results of Performance Measurements on a Cylindrical Hall-Effect Thruster with Magnetic Field Generated by Permanent Magnets

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Raitses, Y.; Merino, E.; Fisch, N. J.

2008-01-01

The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic configurations. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying a higher thrust efficiency. Preliminary thruster performance measurements on this configuration were obtained over a power range of 100-250 W. The thrust levels over this power range were 3.5-6.5 mN, with anode efficiencies and specific impulses spanning 14-19% and 875- 1425 s, respectively. The magnetic field in the thruster was lower for the thrust measurements than the plasma probe measurements due to heating and weakening of the permanent magnets, reducing the maximum field strength from 2 kG to roughly 750-800 G. The discharge current levels observed during thrust stand testing were anomalously high compared to those levels measured in previous experiments with this thruster.

Interior and Exterior Laser-Induced Fluorescence and Plasma Potential Measurements on a Laboratory Hall Thruster (Postprint)

DTIC Science & Technology

1999-06-01

Hall thruster is provided by a 1 mm axial slot in the insulator outer wall. Axial ion velocity profiles for four discharge voltages (100 V, 160 V, 200 V, 250 V) are measured as are radial velocity profiles in the near field plume. Internal neutral xenon axial velocity profiles are also measured at these conditions. For comparison, the plume plasma potential profile is measured with an emissive probe. These probe based potential measurements extend from 50 mm outside the plume to the near anode region for all but the highest discharge voltage condition. For each condition,

Integrated Stirling Convertor and Hall Thruster Test Conducted

NASA Technical Reports Server (NTRS)

Mason, Lee S.

2002-01-01

An important aspect of implementing Stirling Radioisotope Generators on future NASA missions is the integration of the generator and controller with potential spacecraft loads. Some recent studies have indicated that the combination of Stirling Radioisotope Generators and electric propulsion devices offer significant trip time and payload fraction benefits for deep space missions. A test was devised to begin to understand the interactions between Stirling generators and electric thrusters. An electrically heated RG- 350 (350-W output) Stirling convertor, designed and built by Stirling Technology Company of Kennewick, Washington, under a NASA Small Business Innovation Research agreement, was coupled to a 300-W SPT-50 Hall-effect thruster built for NASA by the Moscow Aviation Institute (RIAME). The RG-350 and the SPT-50 shown, were installed in adjacent vacuum chamber ports at NASA Glenn Research Center's Electric Propulsion Laboratory, Vacuum Facility 8. The Stirling electrical controller interfaced directly with the Hall thruster power-processing unit, both of which were located outside of the vacuum chamber. The power-processing unit accepted the 48 Vdc output from the Stirling controller and distributed the power to all the loads of the SPT-50, including the magnets, keeper, heater, and discharge. On February 28, 2001, the Glenn test team successfully operated the Hall-effect thruster with the Stirling convertor. This is the world's first known test of a dynamic power source with electric propulsion. The RG-350 successfully managed the transition from the purely resistive load bank within the Stirling controller to the highly capacitive power-processing unit load. At the time of the demonstration, the Stirling convertor was operating at a hot temperature of 530 C and a cold temperature of -6 C. The linear alternator was producing approximately 250 W at 109 Vac, while the power-processing unit was drawing 175 W at 48 Vdc. The majority of power was delivered to the

Non-invasive Hall current distribution measurement in a Hall effect thruster

NASA Astrophysics Data System (ADS)

Mullins, Carl R.; Farnell, Casey C.; Farnell, Cody C.; Martinez, Rafael A.; Liu, David; Branam, Richard D.; Williams, John D.

2017-01-01

A means is presented to determine the Hall current density distribution in a closed drift thruster by remotely measuring the magnetic field and solving the inverse problem for the current density. The magnetic field was measured by employing an array of eight tunneling magnetoresistive (TMR) sensors capable of milligauss sensitivity when placed in a high background field. The array was positioned just outside the thruster channel on a 1.5 kW Hall thruster equipped with a center-mounted hollow cathode. In the sensor array location, the static magnetic field is approximately 30 G, which is within the linear operating range of the TMR sensors. Furthermore, the induced field at this distance is approximately tens of milligauss, which is within the sensitivity range of the TMR sensors. Because of the nature of the inverse problem, the induced-field measurements do not provide the Hall current density by a simple inversion; however, a Tikhonov regularization of the induced field does provide the current density distributions. These distributions are shown as a function of time in contour plots. The measured ratios between the average Hall current and the average discharge current ranged from 6.1 to 7.3 over a range of operating conditions from 1.3 kW to 2.2 kW. The temporal inverse solution at 1.5 kW exhibited a breathing mode frequency of 24 kHz, which was in agreement with temporal measurements of the discharge current.

Non-invasive Hall current distribution measurement in a Hall effect thruster.

PubMed

Mullins, Carl R; Farnell, Casey C; Farnell, Cody C; Martinez, Rafael A; Liu, David; Branam, Richard D; Williams, John D

2017-01-01

A means is presented to determine the Hall current density distribution in a closed drift thruster by remotely measuring the magnetic field and solving the inverse problem for the current density. The magnetic field was measured by employing an array of eight tunneling magnetoresistive (TMR) sensors capable of milligauss sensitivity when placed in a high background field. The array was positioned just outside the thruster channel on a 1.5 kW Hall thruster equipped with a center-mounted hollow cathode. In the sensor array location, the static magnetic field is approximately 30 G, which is within the linear operating range of the TMR sensors. Furthermore, the induced field at this distance is approximately tens of milligauss, which is within the sensitivity range of the TMR sensors. Because of the nature of the inverse problem, the induced-field measurements do not provide the Hall current density by a simple inversion; however, a Tikhonov regularization of the induced field does provide the current density distributions. These distributions are shown as a function of time in contour plots. The measured ratios between the average Hall current and the average discharge current ranged from 6.1 to 7.3 over a range of operating conditions from 1.3 kW to 2.2 kW. The temporal inverse solution at 1.5 kW exhibited a breathing mode frequency of 24 kHz, which was in agreement with temporal measurements of the discharge current.

Performance characteristics of No-Wall-Losses Hall Thruster

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Peng, Wuji; Sun, Hezhi; Wei, Liqiu; Zeng, Ming; Wang, Fufeng; Yu, Daren

2017-08-01

A 200 W No-Wall-Losses Hall Thruster (NWLHT-200 W) is designed and processed to verify the technology of pushing down magnetic field with two permanent magnetic rings. To create a magnetic field, NWLHT-200 W uses two permanent magnetic rings (inner and outer) in the absence of magnetic screen or magnetic component. The anode is at the internal magnetic separatrix position, and the thruster shell is hollow to enhance the heat dissipation of ceramics. The magnetic field strength at the channel outlet is 90% of the maximum magnetic field. In this study, the experimental results concerning the thrust, discharge current, specific impulse, and efficiency are presented and examined. Our experiments show that "no erosive discharge" of wall is achieved within the range of 120-460 W; the maximum efficiency of the anode may reach 49%. The thruster designed can work stably for a long time, without any auxiliary heat dissipation equipment (heat pipe or radiator), which significantly prolongs the life of Hall thrusters.

On channel interactions in nested Hall thrusters

NASA Astrophysics Data System (ADS)

Cusson, S. E.; Georgin, M. P.; Dragnea, H. C.; Dale, E. T.; Dhaliwal, V.; Boyd, I. D.; Gallimore, A. D.

2018-04-01

Nested Hall thrusters use multiple, concentric discharge channels to increase thrust density. They have shown enhanced performance in multi-channel operation relative to the superposition of individual channels. The X2, a two-channel nested Hall thruster, was used to investigate the mechanism behind this improved performance. It is shown that the local pressure near the thruster exit plane is an order of magnitude higher in two-channel operation. This is due to the increased neutral flow inherent to the multi-channel operation. Due to the proximity of the discharge channels in nested Hall thrusters, these local pressure effects are shown to be responsible for the enhanced production of thrust during multi-channel operation via two mechanisms. The first mechanism is the reduction of the divergence angle due to an upstream shift of the acceleration region. The displacement of the acceleration region was detected using laser induced fluorescence measurements of the ion velocity profile. Analysis of the change in beam divergence indicates that, at an operating condition of 150 V and 30 A, this effect increases the thrust by 8.7 ± 1.2 mN. The second mechanism is neutral ingestion from the adjacent channel resulting in a 2.0 + 0/-0.2 mN increase in thrust. Combined, these mechanisms are shown to explain, within uncertainty, the 17 ± 6.2 mN improvement in thrust during dual channel operation of the X2.

The Impact of Harness Impedance on Hall Thruster Discharge Oscillations

NASA Technical Reports Server (NTRS)

Pinero, Luis R.

2017-01-01

Hall thrusters exhibit characteristic discharge voltage and current oscillations during steady-state operation. The lower frequency breathing-mode current oscillations are inherent to each thruster and could impact thruster operation and power processing unit (PPU) design. The design of the discharge output filter, in particular, the output capacitor is important because it supplies the high peak current oscillations that the thruster demands. However, space-rated, high-voltage capacitors are not readily available and can have significant mass and volume. So, it is important for a PPU designer to know what is the minimum amount of capacitance required to operate a thruster. Through Simulation Program with Integrated Circuit Emphasis modeling and electrical measurements on the Hall Effect Rocket with Magnetic Shielding thruster, it was shown that the harness impedance between the power supply and the thruster is the main contributor towards generating voltage ripple at the thruster. Also, increasing the size of the discharge filter capacitor, as previously implemented during thruster tests, does not reduce the voltage oscillations. The electrical characteristics of the electrical harness between the discharge supply and the thruster is crucial to system performance and could have a negative impact on performance, life and operation.

The Impact of Harness Impedance on Hall Thruster Discharge Oscillations

NASA Technical Reports Server (NTRS)

Pinero, Luis R.

2017-01-01

Hall thrusters exhibit characteristic discharge voltage and current oscillations during steady-state operation. The lower frequency breathing-mode current oscillations are inherent to each thruster and could impact thruster operation and PPU design. The design of the discharge output filter, in particular, the output capacitor is important because it supplies the high peak current oscillations that the thruster demands. However, space-rated, high-voltage capacitors are not readily available and can have significant mass and volume. So, it is important for a PPU designer to know what is the minimum amount of capacitance required to operate a thruster. Through SPICE modeling and electrical measurements on the Hall Effect Rocket with Magnetic Shielding (HERMeS) thruster, it was shown that the harness impedance between the power supply and the thruster is the main contributor towards generating voltage ripple at the thruster. Also, increasing the size of the discharge filter capacitor, as previously implemented during thruster tests, does not reduce the voltage oscillations. The electrical characteristics of the electrical harness between the discharge supply and the thruster is crucial to system performance and could have a negative impact on performance, life and operation.

Simulation of double stage hall thruster with double-peaked magnetic field

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Li, Peng; Sun, Hezhi; Wei, Liqiu; Xu, Yu; Peng, Wuji; Su, Hongbo; Li, Hong; Yu, Daren

2017-07-01

This study adopts double permanent magnetic rings and four permanent magnetic rings to form two symmetrical magnetic peaks and two asymmetrical magnetic peaks in the channel of a Hall thruster, and uses a 2D-3V PIC-MCC model to analyze the influence of magnetic strength on the discharge characteristic and performance of Hall thrusters with an intermediate electrode and double-peaked magnetic field. As opposed to the two symmetrical magnetic peaks formed by double permanent magnetic rings, increasing the magnetic peak value deep within the channel can cause propellant ionization to occur; with the increase in the magnetic peak deep in the channel, the propellant utilization, thrust, and anode efficiency of the thruster are significantly improved. Double-peaked magnetic field can realize separate control of ionization and acceleration in a Hall thruster, and provide technical means for further improving thruster performance. Contribution to the Topical Issue "Physics of Ion Beam Sources", edited by Holger Kersten and Horst Neumann.

A Planar Hall Thruster for Investigating Electron Mobility in ExB Devices (Preprint)

DTIC Science & Technology

2007-08-24

Hall thruster that emits and collects the Hall current across a planar discharge channel is described. The planar Hall thruster (PHT) is being investigated for use as a test bed to study electron mobility in ExB devices. The planar geometry attempts to de-couple the complex electron motion found in annular thrusters by using simplified geometry. During this initial test, the PHT was operated at discharge voltages between 50-150 V to verify operability and stability of the device. Hall current was emitted by hollow cathode electron sources and

Thrust and Performance Study of Micro Pulsed Plasma Thrusters

DTIC Science & Technology

2010-03-01

Due to the high- voltage potential, numerous electrons are able to collect in a small area. As the collection of the electrons grows, the ...quasi- neutral plasma removes the need to have a second emitter of free electrons to neutralize the plasma like in the Hall thrusters. PPTs and µPPTs...surface of the cathode. The micro-protrusions

1000 Hours of Testing Completed on 10-kW Hall Thruster

NASA Technical Reports Server (NTRS)

Mason, Lee S.

2001-01-01

Between the months of April and August 2000, a 10-kW Hall effect thruster, designated T- 220, was subjected to a 1000-hr life test evaluation. Hall effect thrusters are propulsion devices that electrostatically accelerate xenon ions to produce thrust. Hall effect propulsion has been in development for many years, and low-power devices (1.35 kW) have been used in space for satellite orbit maintenance. The T-220, shown in the photo, produces sufficient thrust to enable efficient orbital transfers, saving hundreds of kilograms in propellant over conventional chemical propulsion systems. This test is the longest operation ever achieved on a high-power Hall thruster (greater than 4.5 kW) and is a key milestone leading to the use of this technology for future NASA, commercial, and military missions.

In-Situ Measurement of Hall Thruster Erosion Using a Fiber Optic Regression Probe

NASA Technical Reports Server (NTRS)

Polzink, Kurt A.; Korman, Valentin

2008-01-01

One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster's operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor was tested using a linear Hall thruster geometry, which served as a means of producing plasma erosion of a ceramic discharge chamber. The mass flow rate, discharge voltage, and applied magnetic field strength could be varied, allowing for erosion measurements over a broad thruster operating envelope. Results are presented demonstrating the ability of the REAST sensor to capture not only the insulator erosion rates but also changes in these rates as a function of the discharge parameters.

Study of electron transport in a Hall thruster by axial–radial fully kinetic particle simulation

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cho, Shinatora, E-mail: choh.shinatora@jaxa.jp; Kubota, Kenichi; Funaki, Ikkoh

2015-10-15

Electron transport across a magnetic field in a magnetic-layer-type Hall thruster was numerically investigated for the future predictive modeling of Hall thrusters. The discharge of a 1-kW-class magnetic-layer-type Hall thruster designed for high-specific-impulse operation was modeled using an r-z two-dimensional fully kinetic particle code with and without artificial electron-diffusion models. The thruster performance results showed that both electron transport models captured the experimental result within discrepancies less than 20% in thrust and discharge current for all the simulated operation conditions. The electron cross-field transport mechanism of the so-called anomalous diffusion was self-consistently observed in the simulation without artificial diffusion models;more » the effective electron mobility was two orders of magnitude higher than the value obtained using the classical diffusion theory. To account for the self-consistently observed anomalous transport, the oscillation of plasma properties was speculated. It was suggested that the enhanced random-walk diffusion due to the velocity oscillation of low-frequency electron flow could explain the observed anomalous diffusion within an order of magnitude. The dominant oscillation mode of the electron flow velocity was found to be 20 kHz, which was coupled to electrostatic oscillation excited by global ionization instability.« less

Carbon Back Sputter Modeling for Hall Thruster Testing

NASA Technical Reports Server (NTRS)

Gilland, James H.; Williams, George J.; Burt, Jonathan M.; Yim, John Tamin

2016-01-01

Lifetime requirements for electric propulsion devices, including Hall Effect thrusters, are continually increasing, driven in part by NASA's inclusion of this technology in it's exploration architecture. NASA will demonstrate high-power electric propulsion system on the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). The Asteroid Redirect Robotic mission is one candidate SEP TDM, which is projected to require tens of thousands of thruster life. As thruster life is increased, for example through the use of improved magnetic field designs, the relative influence of facility effects increases. One such effect is the sputtering and redeposition, or back sputter, of facility materials by the high energy thruster plumes. In support of wear testing for the Hall Effect Rocket with Magnetic Shielding (HERMeS) project, the back sputter from a Hall effect thruster plume has been modeled for the NASA Glenn Research Center's Vacuum Facility 5. The predicted wear at a near-worst case condition of 600 V, 12.5 kW was found to be on the order of 1 micron/kh in a fully carbon-lined chamber. A more detailed numerical Monte Carlo code was also modified to estimate back sputter for a detailed facility and pumping configuration. This code demonstrated similar back sputter rate distributions, but is not yet accurately modeling the magnitudes. The modeling has been benchmarked to recent HERMeS wear testing, using multiple microbalance measurements. These recent measurements have yielded values on the order of 1.5 - 2 micron/kh at 600 V and 12.5 kW.

ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS IN HALL EFFECT THRUSTERS

DTIC Science & Technology

2017-06-30

17394 4 / 13 HALL EFFECT THRUSTERS Hall Effect Thrusters (HET): Traditionally Modeled in R-Z Named for Hall Current in θ Uses Quasi -1D Electron Fluid...HET): Traditionally Modeled in R-Z Named for Hall Current in θ Uses Quasi -1D Electron Fluid Solve Ohm’s Law→ No e−-momentum Zθ Unrolled to YZ...Current in θ Uses Quasi -1D Electron Fluid Solve Ohm’s Law→ No e−-momentum Zθ Unrolled to YZ Electron ExB Drift Unmagnetized Ions Results in Hall Current

Effects of an Internally-Mounted Cathode on Hall Thruster Plume Properties

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Johnson, Lee K.; Goebel, Dan M.; Fitzgerald, Dennis J.

2006-01-01

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.

In-Situ Measurement of Hall Thruster Erosion Using a Fiber Optic Regression Probe

NASA Technical Reports Server (NTRS)

Polzin, Kurt; Korman, Valentin

2009-01-01

the thruster is in operation (i.e. none yield a continuous channel erosion measurement). A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor was tested using a linear Hall thruster geometry (see Fig. 1), which served as a means of producing plasma erosion of a ceramic discharge chamber. The mass flow rate, discharge voltage, and applied magnetic field strength could be varied, allowing for erosion measurements over a broad thruster operating envelope. Results are presented demonstrating the ability of the REAST sensor to capture not only the insulator erosion rates but also changes in these rates as a function of the discharge parameters.

A non-invasive Hall current distribution measurement system for Hall Effect thrusters

NASA Astrophysics Data System (ADS)

Mullins, Carl Raymond

A direct, accurate method to measure thrust produced by a Hall Effect thruster on orbit does not currently exist. The ability to calculate produced thrust will enable timely and precise maneuvering of spacecraft---a capability particularly important to satellite formation flying. The means to determine thrust directly is achievable by remotely measuring the magnetic field of the thruster and solving the inverse magnetostatic problem for the Hall current density distribution. For this thesis, the magnetic field was measured by employing an array of eight tunneling magnetoresistive (TMR) sensors capable of milligauss sensitivity when placed in a high background field. The array was positioned outside the channel of a 1.5 kW Colorado State University Hall thruster equipped with a center-mounted electride cathode. In this location, the static magnetic field is approximately 30 Gauss, which is within the linear operating range of the TMR sensors. Furthermore, the induced field at this distance is greater than tens of milligauss, which is within the sensitivity range of the TMR sensors. Due to the nature of the inverse problem, the induced-field measurements do not provide the Hall current density by a simple inversion; however, a Tikhonov regularization of the induced field along with a non-negativity constraint and a zero boundary condition provides current density distributions. Our system measures the sensor outputs at 2 MHz allowing the determination of the Hall current density distribution as a function of time. These data are shown in contour plots in sequential frames. The measured ratios between the average Hall current and the discharge current ranged from 0.1 to 10 over a range of operating conditions from 1.3 kW to 2.2 kW. The temporal inverse solution at 2.0 kW exhibited a breathing mode of 37 kHz, which was in agreement with temporal measurements of the discharge current.

Performance of a Cylindrical Hall-Effect Thruster Using Permanent Magnets

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Raitses, Y.; Merino, E.; Fisch, N. J.

2009-01-01

While annular Hall thrusters can operate at high efficiency at kW power levels, it is difficult to construct one that operates over a broad envelope from 1 kW down to 100 W while maintaining an efficiency of 45-55%. Scaling to low power while holding the main dimensionless parameters constant requires a decrease in the thruster channel size and an increase in the magnetic field strength. Increasing the magnetic field becomes technically challenging since the field can saturate the miniaturized inner components of the magnetic circuit and scaling down the magnetic circuit leaves very little room for magnetic pole pieces and heat shields. In addition, the central magnetic pole piece defining the interior wall of the annular channel can experience excessive heat loads in a miniaturized Hall thruster, with the temperature eventually exceeding the Curie temperature of the material and in extreme circumstances leading to accelerated erosion of the channel wall. An alternative approach is to employ a cylindrical Hall thruster (CHT) geometry. Laboratory model CHTs have operated at power levels ranging from 50 W up to 1 kW. These thrusters exhibit performance characteristics that are comparable to conventional, annular Hall thrusters of similar size. Compared to the annular Hall thruster, the CHTs insulator surface area to discharge chamber volume ratio is lower. Consequently, there is the potential for reduced wall losses in the channel of a CHT, and any reduction in wall losses should translate into lower channel heating rates and reduced erosion, making the CHT geometry promising for low-power applications. This potential for high performance in the low-power regime has served as the impetus for research and development efforts aimed at understanding and improving CHT performance. Recently, a 2.6 cm channel diameter permanent magnet CHT (shown in Fig. 1) was tested. This thruster has the promise of reduced power consumption over previous CHT iterations that employed

Iodine Hall Thruster Propellant Feed System for a CubeSat

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.

2014-01-01

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ?I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of

Investigations of an Environmentally Induced Long Duration Hall Thruster Start Transient (PREPRINT)

DTIC Science & Technology

2006-02-06

Hall thruster start transient is produced by exposure of the thruster to ambient laboratory atmosphere. This behavior was first observed during operation of a cluster of four 200 W BHT-200 Hall effect thrusters where large anode discharge fluctuations, visible as increased anode current and a diffuse plume structure, occurred in an apparently random manner. During operation of a single thruster, the start transient appears as a quickly rising and later smoothly decaying elevated anode current with a diffuse plume that persists for less than 500 seconds. The start transient

Modeling an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT)

NASA Technical Reports Server (NTRS)

Choi, Maria

2016-01-01

An iodine-operated 200-W Hall thruster plume has been simulated using a hybrid-PIC model to predict the spacecraft surface-plume interaction for spacecraft integration purposes. For validation of the model, the plasma potential, electron temperature, ion current flux, and ion number density of xenon propellant were compared with available measurement data at the nominal operating condition. To simulate iodine plasma, various collision cross sections were found and used in the model. While time-varying atomic iodine species (i.e., I, I+, I2+) information is provided by HP Hall simulation at the discharge channel exit, the molecular iodine species (i.e., I2, I2+) are introduced as Maxwellian particles at the channel exit. Simulation results show that xenon and iodine plasma plumes appear to be very similar under the assumptions of the model. Assuming a sticking coefficient of unity, iodine deposition rate is estimated.

Comparisons in Performance of Electromagnet and Permanent-Magnet Cylindrical Hall-Effect Thrusters

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Raitses, Y.; Gayoso, J. C.; Fisch, N. J.

2010-01-01

Three different low-power cylindrical Hall thrusters, which more readily lend themselves to miniaturization and low-power operation than a conventional (annular) Hall thruster, are compared to evaluate the propulsive performance of each. One thruster uses electromagnet coils to produce the magnetic field within the discharge channel while the others use permanent magnets, promising power reduction relative to the electromagnet thruster. A magnetic screen is added to the permanent magnet thruster to improve performance by keeping the magnetic field from expanding into space beyond the exit of the thruster. The combined dataset spans a power range from 50-350 W. The thrust levels over this range were 1.3-7.3 mN, with thruster efficiencies and specific impulses spanning 3.5-28.7% and 400-1940 s, respectively. The efficiency is generally higher for the permanent magnet thruster with the magnetic screen, while That thruster s specific impulse as a function of discharge voltage is comparable to the electromagnet thruster.

Ion Voltage Diagnostics in the Far-Field Plume of a High-Specific Impulse Hall Thruster

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Haas, James M.; Gallimore, Alec D.

2003-01-01

The effects of the magnetic field and discharge voltage on the far-field plume of the NASA 173Mv2 laboratory-model Hall thruster were investigated. A cylindrical Langmuir probe was used to measure the plasma potential and a retarding potential analyzer was employed to measure the ion voltage distribution. The plasma potential was affected by relatively small changes in the external magnetic field, which suggested a means to control the plasma surrounding the thruster. As the discharge voltage increased, the ion voltage distribution showed that the acceleration efficiency increased and the dispersion efficiency decreased. This implied that the ionization zone was growing axially and moving closer to the anode, which could have affected thruster efficiency and lifetime due to higher wall losses. However, wall losses may have been reduced by improved focusing efficiency since the total efficiency increased and the plume divergence decreased with discharge voltage.

The Effects of Insulator Wall Material on Hall Thruster Discharges: A Numerical Study

DTIC Science & Technology

2001-01-03

An investigation was undertaken to determine how the choice of insulator wall material inside a Hall thruster discharge channel might affect thruster operation. In order to study this, an evolved hybrid particle-in-cell (PIC) numerical Hall thruster model, HPHall, was used. HPHall solves a set of quasi-one-dimensional fluid equations for electrons and tracks heavy particles using a PIC method.

Laser characterization of electric field oscillations in the Hall thruster breathing mode

NASA Astrophysics Data System (ADS)

Young, Christopher; Lucca Fabris, Andrea; MacDonald-Tenenbaum, Natalia; Hargus, William, Jr.; Cappelli, Mark

2016-10-01

Hall thrusters are a mature technology for space propulsion applications that exhibit a wide array of dynamic behavior, including plasma waves, instabilities and turbulence. One common low frequency (10-50 kHz) discharge current oscillation is the breathing mode, a cycle of neutral propellant injection, strong ionization, and ion acceleration by a steep potential gradient. A time-resolved laser-induced fluorescence diagnostic non-intrusively captures this propagating ionization front in the channel of a commercial BHT-600 Hall thruster manufactured by Busek Co. Measurements of ion velocity and relative ion density (using the 5 d[ 4 ] 7 / 2 - 6 p[ 3 ] 5 / 2 Xe II transition at 834.95 nm, vacuum) reveal a dynamic electric field structure traversing the channel throughout the breathing mode cycle. This work is sponsored by the U.S. Air Force Office of Scientific Research, with Dr. M. Birkan as program manager. C.Y. acknowledges support from the DOE NSSA Stewardship Science Graduate Fellowship under contract DE-FC52-08NA28752.

Design and Testing of a Hall Effect Thruster with Additively Manufactured Components

NASA Astrophysics Data System (ADS)

Hopping, Ethan

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville to study the application of low-cost additive manufacturing in the design and fabrication of Hall thrusters. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. The thruster features channel walls and a propellant distributor that were manufactured using 3D printing with a variety of materials including ABS, ULTEM, and glazed ceramic. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. The design of the thruster and the transient performance measurements are presented here. Measured thrust ranged from 17.2 mN to 30.4 mN over a discharge power of 280 W to 520 W with an anode Isp range of 870 s to 1450 s. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state. While the current thruster design is not yet ready for continuous operation, revisions to the device that could enable longer duration tests are discussed.

50 KW Class Krypton Hall Thruster Performance

NASA Technical Reports Server (NTRS)

Jacobson, David T.; Manzella, David H.

2003-01-01

The performance of a 50-kilowatt-class Hall thruster designed for operation on xenon propellant was measured using kryton propellant. The thruster was operated at discharge power levels ranging from 6.4 to 72.5 kilowatts. The device produced thrust ranging from 0.3 to 2.5 newtons. The thruster was operated at discharge voltages between 250 and 1000 volts. At the highest anode mass flow rate and discharge voltage and assuming a 100 percent singly charged condition, the discharge specific impulse approached the theoretical value. Discharge specific impulse of 4500 seconds was demonstrated at a discharge voltage of 1000 volts. The peak discharge efficiency was 64 percent at 650 volts.

Nonlinear excitation of long-wavelength modes in Hall plasmas

NASA Astrophysics Data System (ADS)

Lakhin, V. P.; Ilgisonis, V. I.; Smolyakov, A. I.; Sorokina, E. A.

2016-10-01

Hall plasmas with magnetized electrons and unmagnetized ions exhibit a wide range of small scale fluctuations in the lower-hybrid frequency range as well as low-frequency large scale modes. Modulational instability of lower-hybrid frequency modes is investigated in this work for typical conditions in Hall plasma devices such as magnetrons and Hall thrusters. In these conditions, the dispersion of the waves in the lower-hybrid frequency range propagating perpendicular to the external magnetic field is due to the gradients of the magnetic field and the plasma density. It is shown that such lower-hybrid modes are unstable with respect to the secondary instability of the large scale quasimode perturbations. It is suggested that the large scale slow coherent modes observed in a number of Hall plasma devices may be explained as a result of such secondary instabilities.

Assessment of Pole Erosion in a Magnetically Shielded Hall Thruster

NASA Technical Reports Server (NTRS)

Mikellides, Ioannis G.; Ortega, Alejandro L.

2014-01-01

Numerical simulations of a 6-kW laboratory Hall thruster called H6 have been performed to quantify the erosion rate at the inner pole. The assessments have been made in two versions of the thruster, namely the unshielded (H6US) and magnetically shielded (H6MS) configurations. The simulations have been performed with the 2-D axisymmetric code Hall2De which employs a new multi-fluid ion algorithm to capture the presence of low-energy ions in the vicinity of the poles. It is found that the maximum computed erosion rate at the inner pole of the H6MS exceeds the measured rate of back-sputtered deposits by 4.5 times. This explains only part of the surface roughening that was observed after a 150-h wear test, which covered most of the pole area exposed to the plasma. For the majority of the pole surface the computed erosion rates are found to be below the back-sputter rate and comparable to those in the H6US which exhibited little to no sputtering in previous tests. Possible explanations for the discrepancy are discussed.

An Inversion Method for Reconstructing Hall Thruster Plume Parameters from the Line Integrated Measurements (Preprint)

DTIC Science & Technology

2007-06-05

From - To) 05-06-2007 Technical Paper 4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER An Inversion Method for Reconstructing Hall Thruster Plume...239.18 An Inversion Method for Reconstructing Hall Thruster Plume Parameters from Line Integrated Measurements (Preprint) Taylor S. Matlock∗ Jackson...dimensional estimate of the plume electron temperature using a published xenon collisional radiative model. I. Introduction The Hall thruster is a high

Overview of Iodine Propellant Hall Thruster Development Activities at NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Haag, Thomas; Benavides, Gabriel; Hickman, Tyler; Smith, Timothy; Williams, George; Myers, James; Polzin, Kurt; Dankanich, John; Byrne, Larry;

Overview of Iodine Propellant Hall Thruster Development Activities at NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Benavides, Gabriel; Haag, Thomas; Hickman, Tyler; Smith, Timothy; Williams, George; Myers, James; Polzin, Kurt; Dankanich, John; Byrne, Larry;

A cavity ring-down spectroscopy sensor for real-time Hall thruster erosion measurements.

PubMed

Lee, B C; Huang, W; Tao, L; Yamamoto, N; Gallimore, A D; Yalin, A P

2014-05-01

A continuous-wave cavity ring-down spectroscopy sensor for real-time measurements of sputtered boron from Hall thrusters has been developed. The sensor uses a continuous-wave frequency-quadrupled diode laser at 250 nm to probe ground state atomic boron sputtered from the boron nitride insulating channel. Validation results from a controlled setup using an ion beam and target showed good agreement with a simple finite-element model. Application of the sensor for measurements of two Hall thrusters, the H6 and SPT-70, is described. The H6 was tested at power levels ranging from 1.5 to 10 kW. Peak boron densities of 10 ± 2 × 10(14) m(-3) were measured in the thruster plume, and the estimated eroded channel volume agreed within a factor of 2 of profilometry. The SPT-70 was tested at 600 and 660 W, yielding peak boron densities of 7.2 ± 1.1 × 10(14) m(-3), and the estimated erosion rate agreed within ~20% of profilometry. Technical challenges associated with operating a high-finesse cavity in the presence of energetic plasma are also discussed.

Propulsion Instruments for Small Hall Thruster Integration

NASA Technical Reports Server (NTRS)

Johnson, Lee K.; Conroy, David G.; Spanjers, Greg G.; Bromaghim, Daron R.

2001-01-01

Planning and development are underway for the propulsion instrumentation necessary for the next AFRL electric propulsion flight project, which includes both a small Hall thruster and a micro-PPT. These instruments characterize the environment induced by the thruster and the associated data constitute part of a 'user's manual' for these thrusters. Several instruments probe the back-flow region of the thruster plume, and the data are intended for comparison with detailed numerical models in this region. Specifically, an ion probe is under development to determine the energy and species distributions, and a Langmuir probe will be employed to characterize the electron density and temperature. Other instruments directly measure the effects of thruster operation on spacecraft thermal control surfaces, optical surfaces, and solar arrays. Specifically, radiometric, photometric, and solar-cell-based sensors are under development. Prototype test data for most sensors should be available, together with details of the instrumentation subsystem and spacecraft interface.

Performance of Solar Electric Powered Deep Space Missions Using Hall Thruster Propulsion

NASA Technical Reports Server (NTRS)

Witzberger, Kevin E.; Manzella, David

2006-01-01

Power limited, low-thrust trajectories were assessed for missions to Jupiter, Saturn, and Neptune utilizing a single Venus Gravity Assist (VGA) and a primary propulsion system based on either a 3-kW high voltage Hall thruster, of the type being developed by the NASA In-Space Propulsion Technology Program, or an 8-kW variant of this thruster. These Hall thrusters operate with specific impulses below 3,000 seconds. A trade study was conducted to examine mission parameters that include: net delivered mass (NDM), beginning-of-life (BOL) solar array power, heliocentric transfer time, required launch vehicle, number of operating thrusters, and throttle profile. The top performing spacecraft configuration was defined to be the one that delivered the highest mass for a range of transfer times. In order to evaluate the potential future benefit of using next generation Hall thrusters as the primary propulsion system, comparisons were made with the advanced state-of-the-art (ASOA), 7-kW, 4,100 second NASA's Evolutionary Xenon Thruster (NEXT) for the same mission scenarios. For the BOL array powers considered in this study (less than 30 kW), the results show that the performance of the Hall thrusters, relative to NEXT, is largely dependant on the performance capability of the launch vehicle, and that at least a 10 percent performance gain, equating to at least an additional 200 kg dry mass at each target planet, is achieved over the higher specific impulse NEXT when launched on an Atlas 551.

Non-Contact Thermal Characterization of NASA's HERMeS Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Meyers, James L.; Yim, John T.; Neff, Gregory

2015-01-01

The Thermal Characterization Test of NASAs 12.5-kW Hall thruster is being completed. This thruster is being developed to support of a number of potential Solar Electric Propulsion Technology Demonstration Mission concepts, including the Asteroid Redirect Robotic Mission concept. As a part of this test, an infrared-based, non-contact thermal imaging system was developed to measure Hall thruster surfaces that are exposed to high voltage or harsh environment. To increase the accuracy of the measurement, a calibration array was implemented, and a pilot test was performed to determine key design parameters for the calibration array. The raw data is analyzed in conjunction with a simplified thermal model of the channel to account for reflection. The reduced data will be used to refine the thruster thermal model, which is critical to the verification of the thruster thermal specifications. The present paper will give an overview of the decision process that led to identification of the need for a non-contact temperature diagnostic, the development of said diagnostic, the measurement results, and the simplified thermal model of the channel.

Characteristics and transport effects of the electron drift instability in Hall-effect thrusters

NASA Astrophysics Data System (ADS)

Lafleur, T.; Baalrud, S. D.; Chabert, P.

2017-02-01

The large electron {E}× {B} drift (relative to the ions) in the azimuthal direction of Hall-effect thrusters is well known to excite a strong instability. In a recent paper (Lafleur et al 2016 Phys. Plasmas 23 053503) we demonstrated that this instability leads to an enhanced electron-ion friction force that increases the electron cross-field mobility to levels similar to those seen experimentally. Here we extend this work by considering in detail the onset criteria for the formation of this instability (both in xenon, and other propellants of interest), and identify a number of important characteristics that it displays within Hall-effect thrusters (HETs): including the appearance of an additional non-dimensionalized scaling parameter (the instability growth-to-convection ratio), which controls the instability evolution and amplitude. We also investigate the effect that the instability has on electron and ion heating in HETs, and show that it leads to an ion rotation in the azimuthal direction that is in agreement with that seen experimentally.

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Van Noord, Jonathan

2012-01-01

NASA's Office of the Chief Technologist In-Space Propulsion project is sponsoring the testing and development of high power Hall thrusters for implementation in NASA missions. As part of the project, NASA Glenn Research Center is developing and testing new high current hollow cathode assemblies that can meet and exceed the required discharge current and life-time requirements of high power Hall thrusters. This paper presents test results of three high current hollow cathode configurations. Test results indicated that two novel emitter configurations were able to attain lower peak emitter temperatures compared to state-of-the-art emitter configurations. One hollow cathode configuration attained a cathode orifice plate tip temperature of 1132 degC at a discharge current of 100 A. More specifically, test and analysis results indicated that a novel emitter configuration had minimal temperature gradient along its length. Future work will include cathode wear tests, and internal emitter temperature and plasma properties measurements along with detailed physics based modeling.

The Development of Plasma Thrusters and Its Importance for Space Technology and Science Education at University of Brasilia

NASA Astrophysics Data System (ADS)

Ferreira, Jose Leonardo; Calvoso, Lui; Gessini, Paolo; Ferreira, Ivan

Since 2004 The Plasma Physics Laboratory of University of Brasilia (Brazil) is developing Hall Plasma Thurusters for Satellite station keeping and orbit control. The project is supported by CNPq, CAPES, FAP DF and from The Brazillian Space Agency-AEB. The project is part of The UNIESPAÇO Program for Space Activities Development in Brazillian Universities. In this work we are going to present the highlights of this project together with its vital contribution to include University of Brasilia in the Brazillian Space Program. Electric propulsion has already shown, over the years, its great advantages in being used as main and secondary thruster system of several space mission types. Between the many thruster concepts, one that has more tradition in flying real spacecraft is the Hall Effect Thruster (HET). These thrusters, first developed by the USSR in the 1960s, uses, in the traditional design, the radial magnetic field and axial electric field to trap electrons, ionize the gas and accelerate the plasma to therefore generate thrust. In contrast to the usual solution of using electromagnets to generate the magnetic field, the research group of the Plasma Physics Laboratory of University of Brasília has been working to develop new models of HETs that uses combined permanent magnets to generate the necessary magnetic field, with the main objective of saving electric power in the final system design. Since the beginning of this research line it was developed and implemented two prototypes of the Permanent Magnet Hall Thruster (PMHT). The first prototype, called P-HALL1, was successfully tested with the using of many diagnostics instruments, including, RF probe, Langmuir probe, Ion collector and Ion energy analyzer. The second prototype, P-HALL2, is currently under testing, and it’s planned the increasing of the plasma diagnostics and technology analysis, with the inclusion of a thrust balance, mass spectroscopy and Doppler broadening. We are also developing an

Effect of dust on tilted electrostatic resistive instability in a Hall thruster

NASA Astrophysics Data System (ADS)

Tyagi, Jasvendra; Singh, Sukhmander; Malik, Hitendra K.

2018-03-01

Effect of negatively charged dust on resistive instability corresponding to the electrostatic wave is investigated in a Hall thruster plasma when this purely azimuthal wave is tilted and strong axial component of wave vector is developed. Analytical calculations are done to obtain the relevant dispersion equation, which is solved numerically to investigate the growth rate of the instability. The magnitude of the growth rate in the plasma having dust particles is found to be much smaller than the case of pure plasma. However, the instability grows faster for the increasing dust density and the higher charge on the dust particles. The higher magnetic field is also found to support the instability.

Observation of a high-energy tail in ion energy distribution in the cylindrical Hall thruster plasma

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lim, Youbong; Kim, Holak; Choe, Wonho, E-mail: wchoe@kaist.ac.kr

2014-10-15

A novel method is presented to determine populations and ion energy distribution functions (IEDFs) of individual ion species having different charge states in an ion beam from the measured spectrum of an E × B probe. The inversion of the problem is performed by adopting the iterative Tikhonov regularization method with the characteristic matrices obtained from the calculated ion trajectories. In a cylindrical Hall thruster plasma, an excellent agreement is observed between the IEDFs by an E × B probe and those by a retarding potential analyzer. The existence of a high-energy tail in the IEDF is found to be mainly due to singlymore » charged Xe ions, and is interpreted in terms of non-linear ion acceleration.« less

Fundamental Studies of the Electrode Regions in Arcjet Thrusters

DTIC Science & Technology

1998-03-01

Hall thruster . This contributed to a comprehensive study of the near exit region of our Hall discharge device. To compliment the LIF diagnostics on our Hall thrusters, we have made extensive measurements of the transient and time average plasma properties using conventional electrostatic

Computation of Neutral Gas Flow from a Hall Thruster into a Vacuum Chamber

DTIC Science & Technology

2002-10-18

try to quantify these effects, the direct simulation Monte Carlo method is applied to model a cold flow of xenon gas expanding from a Hall thruster into...a vacuum chamber. The simulations are performed for the P5 Hall thruster operating in a large vacuum tank at the University of Michigan. Comparison

Laboratory Model 50 kW Hall Thruster

NASA Technical Reports Server (NTRS)

Manzella, David; Jankovsky, Robert; Hofer, Richard

2002-01-01

A 0.46 meter diameter Hall thruster was fabricated and performance tested at powers up to 72 kilowatts. Thrusts up to 2.9 Newtons were measured. Discharge specific impulses ranged from 1750 to 3250 seconds with discharge efficiencies between 46 and 65 percent. Overall specific impulses ranged from 1550 to 3050 seconds with overall efficiencies between 40 and 57 percent. Performance data indicated significant fraction of multiple-charged ions during operation at elevated power levels. Cathode mass flow rate was shown to be a significant parameter with regard to thruster efficiency.

Effect of Background Pressure on the Plasma Oscillation Characteristics of the HiVHAc Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Lobbia, Robert B.; Brown, Daniel L.

2014-01-01

During a component compatibility test of the NASA HiVHAc Hall thruster, a high-speed camera and a set of high-speed Langmuir probes were implemented to study the effect of varying facility background pressure on thruster operation. The results show a rise in the oscillation frequency of the breathing mode with rising background pressure, which is hypothesized to be due to a shortening accelerationionization zone. An attempt is made to apply a simplified ingestion model to the data. The combined results are used to estimate the maximum acceptable background pressure for performance and wear testing.

Optical Emission Characterization of High-Power Hall Thruster Wear

NASA Technical Reports Server (NTRS)

WIlliams, George J.; Kamhawi, Hani

2013-01-01

Optical emission spectroscopy is employed to correlate BN insulator erosion with high-power operation of the NASA 300M Hall-effect thruster. Actinometry leveraging excited xenon states is used to normalize the emission spectra of ground state boron as a function of thruster operating condition. Trends in the strength of the boron signal are correlated with thruster power, discharge voltage, discharge current and magnetic field strength. The boron signals are shown to trend with discharge current and show weak dependence on discharge voltage. The trends are consistent with data previously collected on the NASA 300M and NASA 457M thrusters but are different from conventional wisdom.

Effect of magnetic field configuration on the multiply charged ion and plume characteristics in Hall thruster plasmas

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kim, Holak; Lim, Youbong; Choe, Wonho, E-mail: wchoe@kaist.ac.kr

2015-04-13

Multiply charged ions and plume characteristics in Hall thruster plasmas are investigated with regard to magnetic field configuration. Differences in the plume shape and the fraction of ions with different charge states are demonstrated by the counter-current and co-current magnetic field configurations, respectively. The significantly larger number of multiply charged and higher charge state ions including Xe{sup 4+} are observed in the co-current configuration than in the counter-current configuration. The large fraction of multiply charged ions and high ion currents in this experiment may be related to the strong electron confinement, which is due to the strong magnetic mirror effectmore » in the co-current magnetic field configuration.« less

High Throughput 600 Watt Hall Effect Thruster for Space Exploration

NASA Technical Reports Server (NTRS)

Szabo, James; Pote, Bruce; Tedrake, Rachel; Paintal, Surjeet; Byrne, Lawrence; Hruby, Vlad; Kamhawi, Hani; Smith, Tim

2016-01-01

A nominal 600-Watt Hall Effect Thruster was developed to propel unmanned space vehicles. Both xenon and iodine compatible versions were demonstrated. With xenon, peak measured thruster efficiency is 46-48% at 600-W, with specific impulse from 1400 s to 1700 s. Evolution of the thruster channel due to ion erosion was predicted through numerical models and calibrated with experimental measurements. Estimated xenon throughput is greater than 100 kg. The thruster is well sized for satellite station keeping and orbit maneuvering, either by itself or within a cluster.

Faraday Probe Analysis, Part 2: Evaluation of Facility Effects on Ion Migration in a Hall Thruster Plume (Preprint)

DTIC Science & Technology

2010-02-24

A nested Faraday probe was designed and fabricated to assess facility effects in a systematic study of ion migration in a Hall thruster plume...Current density distributions were studied at 8, 12, 16, and 20 thruster diameters downstream of the Hall thruster exit plane with four probe configurations...measurements are a significant improvement for comparisons with numerical simulations and investigations of Hall thruster performance.

High-Energy Two-Stage Pulsed Plasma Thruster

NASA Technical Reports Server (NTRS)

Markusic, Tom

2003-01-01

A high-energy (28 kJ per pulse) two-stage pulsed plasma thruster (MSFC PPT-1) has been constructed and tested. The motivation of this project is to develop a high power (approximately 500 kW), high specific impulse (approximately 10000 s), highly efficient (greater than 50%) thruster for use as primary propulsion in a high power nuclear electric propulsion system. PPT-1 was designed to overcome four negative characteristics which have detracted from the utility of pulsed plasma thrusters: poor electrical efficiency, poor propellant utilization efficiency, electrode erosion, and reliability issues associated with the use of high speed gas valves and high current switches. Traditional PPTs have been plagued with poor efficiency because they have not been operated in a plasma regime that fully exploits the potential benefits of pulsed plasma acceleration by electromagnetic forces. PPTs have generally been used to accelerate low-density plasmas with long current pulses. Operation of thrusters in this plasma regime allows for the development of certain undesirable particle-kinetic effects, such as Hall effect-induced current sheet canting. PPT-1 was designed to propel a highly collisional, dense plasma that has more fluid-like properties and, hence, is more effectively pushed by a magnetic field. The high-density plasma loading into the second stage of the accelerator is achieved through the use of a dense plasma injector (first stage). The injector produces a thermal plasma, derived from a molten lithium propellant feed system, which is subsequently accelerated by the second stage using mega-amp level currents, which eject the plasma at a speed on the order of 100 kilometers per second. Traditional PPTs also suffer from dynamic efficiency losses associated with snowplow loading of distributed neutral propellant. The twostage scheme used in PPT-I allows the propellant to be loaded in a manner which more closely approximates the optimal slug loading. Lithium propellant

Efficiency Analysis of a High-Specific Impulse Hall Thruster

NASA Technical Reports Server (NTRS)

Jacobson, David (Technical Monitor); Hofer, Richard R.; Gallimore, Alec D.

2004-01-01

Performance and plasma measurements of the high-specific impulse NASA-173Mv2 Hall thruster were analyzed using a phenomenological performance model that accounts for a partially-ionized plasma containing multiply-charged ions. Between discharge voltages of 300 to 900 V, the results showed that although the net decrease of efficiency due to multiply-charged ions was only 1.5 to 3.0 percent, the effects of multiply-charged ions on the ion and electron currents could not be neglected. Between 300 to 900 V, the increase of the discharge current was attributed to the increasing fraction of multiply-charged ions, while the maximum deviation of the electron current from its average value was only +5/-14 percent. These findings revealed how efficient operation at high-specific impulse was enabled through the regulation of the electron current with the applied magnetic field. Between 300 to 900 V, the voltage utilization ranged from 89 to 97 percent, the mass utilization from 86 to 90 percent, and the current utilization from 77 to 81 percent. Therefore, the anode efficiency was largely determined by the current utilization. The electron Hall parameter was nearly constant with voltage, decreasing from an average of 210 at 300 V to an average of 160 between 400 to 900 V. These results confirmed our claim that efficient operation can be achieved only over a limited range of Hall parameters.

Non-Contact Thermal Characterization of NASA's HERMeS Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Myers, James L.; Yim, John T.; Neff, Gregory

2015-01-01

The thermal characterization test of NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding has been completed. This thruster was developed to support a number of potential Solar Electric Propulsion Technology Demonstration Mission concepts, including the Asteroid Redirect Robotic Mission concept. As a part of the preparation for this characterization test, an infrared-based, non-contact thermal imaging system was developed to measure the temperature of various thruster surfaces that are exposed to high voltage or plasma. An in-situ calibration array was incorporated into the setup to improve the accuracy of the temperature measurement. The key design parameters for the calibration array were determined in a separate pilot test. The raw data from the characterization test was analyzed though further work is needed to obtain accurate anode temperatures. Examination of the front pole and discharge channel temperatures showed that the thruster temperature was driven more by discharge voltage than by discharge power. Operation at lower discharge voltages also yielded more uniform temperature distributions than at higher discharge voltages. When operating at high discharge voltage, increasing the magnetic field strength appeared to have made the thermal loading azimuthally more uniform.

Magnetic Shielding of the Acceleration Channel Walls in a Long-Life Hall Thruster

NASA Technical Reports Server (NTRS)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.; Goebel, Dan M.; de Grys, Kristi; Mathers, Alex

2010-01-01

In a Qualification Life Test (QLT) of the BPT-4000 Hall thruster that recently accumulated greater than 10,000 h it was found that the erosion of the acceleration channel practically stopped after approximately 5,600 h. Numerical simulations of this thruster using a 2-D axisymmetric, magnetic field-aligned-mesh (MFAM) plasma solver reveal that the process that led to this significant reduction of the erosion was multifaceted. It is found that when the channel receded from its early-in-life geometry to its steady-state configuration several changes in the near-wall plasma and sheath were induced by the magnetic field that, collectively, constituted an effective shielding of the walls from any significant ion bombardment. Because all such changes in the behavior of the ionized gas near the eroding surfaces were caused by the topology of the magnetic field there, we term this process "magnetic shielding."

Optical Characterization of Component Wear and Near-Field Plasma of the Hermes Thruster

NASA Technical Reports Server (NTRS)

Williams, George J., Jr.; Kamhawi, Hani

2015-01-01

Optical emission spectral (OES) data are presented which correlate trends in sputtered species and the near-field plasma with the Hall-Effect Rocket with Magnetic Shielding (HERMeS) thruster operating condition. The relative density of singly-ionized xenon (Xe II) is estimated using a collisional-radiative model. OES data were collected at three radial and several axial locations downstream of the thruster's exit plane. These data were deconvolved to show the structure for the near-field plasma as a function of thruster operating condition. The magnetic field is shown to have a much greater affect on plasma structure than the discharge voltage with the primary ionization/acceleration zone boundary being similar for all nominal operating voltages at constant power. OES measurement of sputtered boron shows that the HERMeS thruster is magnetically shielded across its operating envelope. Preliminary assessment of carbon sputtered from the keeper face suggest it increases significantly with operating voltage, but the uncertainty associated with these measurements is very high.

Numerical Modeling of the Hall Thruster Discharge

DTIC Science & Technology

2005-04-01

This collection of seven previously published papers performed under Grant No. FA8655-04-1-3003 provide the background for the development of a new version of the HPHall hybrid code (HPHallv.2) for the numerical modeling of Hall Thruster discharge and new insights on discharge physics obtained during the development.

ExB Measurements of a 200 W Xenon Hall Thruster (Preprint)

DTIC Science & Technology

2007-08-28

Hall thruster Busek BHT-200 plume were measured using an ExB probe under a variety of thruster operating conditions and background pressures. The thruster was operated at several operating conditions by varying the anode potential of the thruster from 200 V to 325 V in 25 V increments. Measurements of the ion species fractions were made 90 from thruster centerline 60 cm downstream of the exit plane. At reduced discharge voltages, the species fractions of multiply-charged xenon ions were lower, while at increased discharge voltages, Xe+2 and Xe+3 showed an increase in their

Hall Thruster With an External Acceleration Zone

DTIC Science & Technology

2005-09-14

Hall Thruster in a high vacuum environment. The ionized propellant velocities were measured using laser induced fluorescence of the excited state xenon ionic transition at 834.7 nm. Ion velocities were interrogated from the channel exit plane to a distance 30 mm from it. Both axial and cross-field (along the electron Hall current direction) velocities were measured. The results presented here, combined with those of previous work, highlight the high sensitivity of electron mobility inside and outside the channel, depending on the background gas density, type of wall

Magnetically Filtered Faraday Probe for Measuring the Ion Current Density Profile of a Hall Thruster

DTIC Science & Technology

2006-01-01

Hall thruster is investigated. The MFFP is designed to eliminate the collection of low-energy, charge-exchange (CEX) ions by using a variable magnetic field as an ion filter. In this study, a MFFP, Faraday probe with a reduced acceptance angle (BFP), and nude Faraday probe are used to measure the ion current density profile of a 5 kW Hall thruster operating over the range of 300-500 V and 5-10 mg/s. The probes are evaluated on a xenon propellant Hall thruster in the University of Michigan Large Vacuum Test Facility at operating

Study of the Accelerating Channel Wall Property Influence on the Hall Thruster Discharge Characteristics

DTIC Science & Technology

2004-11-01

Hall thruster characteristics there was prepared Hall thruster model of the SPT-100 type for these experiments and there were manufactured the required discharge chamber parts (rings) made of the Russian BN-SiO2 (borosil) ceramics and of the Russian AIN-BN (ABN) and Western ABN ceramics having secondary electron emission yield (SEEY) different from that one for borosil. These parts were replaceable during experiments. Thruster model was equipped by set of the near wall probes mounted at external discharge chamber wall. There was made characterization

Performance of a Cylindrical Hall-Effect Thruster with Magnetic Field Generated by Permanent Magnets

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Raitses, Yevgeny; Fisch, Nathaniel J.

2008-01-01

While Hall thrusters can operate at high efficiency at kW power levels, it is difficult to construct one that operates over a broad envelope down to 100W while maintaining an efficiency of 45- 55%. Scaling to low power while holding the main dimensionless parameters constant requires a decrease in the thruster channel size and an increase in the magnetic field strength. Increasing the magnetic field becomes technically challenging since the field can saturate the miniaturized inner components of the magnetic circuit and scaling down the magnetic circuit leaves very little room for magnetic pole pieces and heat shields. An alternative approach is to employ a cylindrical Hall thruster (CHT) geometry. Laboratory model CHTs have operated at power levels ranging from the order of 50 Watts up to 1 kW. These thrusters exhibit performance characteristics which are comparable to conventional, annular Hall thrusters of similar size. Compared to the annular Hall thruster, the CHT has a lower insulator surface area to discharge chamber volume ratio. Consequently, there is the potential for reduced wall losses in the channel of a CHT, and any reduction in wall losses should translate into lower channel heating rates and reduced erosion. This makes the CHT geometry promising for low-power applications. Recently, a CHT that uses permanent magnets to produce the magnetic field topology was tested. This thruster has the promise of reduced power consumption over previous CHT iterations that employed electromagnets. Data are presented for two purposes: to expose the effect different controllable parameters have on the discharge and to summarize performance measurements (thrust, Isp, efficiency) obtained using a thrust stand. These data are used to gain insight into the thruster's operation and to allow for quantitative comparisons between the permanent magnet CHT and the electromagnet CHT.

Effect of Background Pressure on the Performance and Plume of the HiVHAc Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Haag, Thomas

2013-01-01

During the Single String Integration Test of the NASA HiVHAc Hall thruster, a number of plasma diagnostics were implemented to study the effect of varying facility background pressure on thruster operation. These diagnostics include thrust stand, Faraday probe, ExB probe, and retarding potential analyzer. The test results indicated a rise in thrust and discharge current with background pressure. There was also a decrease in ion energy per charge, an increase in multiply-charged species production, a decrease in plume divergence, and a decrease in ion beam current with increasing background pressure. A simplified ingestion model was applied to determine the maximum acceptable background pressure for thrust measurement. The maximum acceptable ingestion percentage was found to be around 1%. Examination of the diagnostics results suggest the ionization and acceleration zones of the thruster were shifting upstream with increasing background pressure.

Optical Diagnostic Characterization of High-Power Hall Thruster Wear and Operation

NASA Technical Reports Server (NTRS)

Williams, George J., Jr.; Soulas, George C.; Kamhawi, Hani

2012-01-01

Optical emission spectroscopy is employed to correlate BN insulator erosion with high-power Hall thruster operation. Specifically, actinometry leveraging excited xenon states is used to normalize the emission spectra of ground state boron as a function of thruster operating condition. Trends in the strength of the boron signal are correlated with thruster power, discharge voltage, and discharge current. In addition, the technique is demonstrated on metallic coupons embedded in the walls of the HiVHAc EM thruster. The OES technique captured the overall trend in the erosion of the coupons which boosts credibility in the method since there are no data to which to calibrate the erosion rates of high-power Hall thrusters. The boron signals are shown to trend linearly with discharge voltage for a fixed discharge current as expected. However, the boron signals of the higher-power NASA 300M and NASA 457Mv2 trend with discharge current and show an unexpectedly weak to inverse dependence on discharge voltage. Electron temperatures measured optically in the near-field plume of the thruster agree well with Langmuir probe data. However, the optical technique used to determine Te showed unacceptable sensitivity to the emission intensities. Near-field, single-frequency imaging of the xenon neutrals is also presented as a function of operating condition for the NASA 457 Mv2.

The effect of segmented anodes on the performance and plume of a Hall thruster

NASA Astrophysics Data System (ADS)

Kieckhafer, Alexander W.

Development of alternative propellants for Hall thruster operation is an active area of research. Xenon is the current propellant of choice for Hall thrusters, but can be costly in large thrusters and for extended test periods. Condensible propellants may offer an alternative to xenon, as they will not require costly active pumping to remove from a test facility, and may be less expensive to purchase. A method has been developed which uses segmented electrodes in the discharge channel of a Hall thruster to divert discharge current to and from the main anode and thus control the anode temperature. By placing a propellant reservoir in the anode, the evaporation rate, and hence, mass flow of propellant can be controlled. Segmented electrodes for thermal control of a Hall thruster represent a unique strategy of thruster design, and thus the performance of the thruster must be measured to determine the effect the electrodes have on the thruster. Furthermore, the source of any changes in thruster performance due to the adjustment of discharge current between the shims and the main anode must be characterized. A Hall thruster was designed and constructed with segmented electrodes. It was then tested at anode voltages between 300 and 400 V and mass flows between 4 and 6 mg/s, as well as 100%, 75%, 50%, 25%, and <5% of the discharge current on the shim electrodes. The level of current on the shims was adjusted by changing the shim voltage. At each operating point, the thruster performance, plume divergence, ion energy, and multiply charged ion fraction were measured. Thruster performance exhibited a small change with the level of discharge current on the shim electrodes. Thrust and specific impulse increased by as much as 6% and 7.7%, respectively, as discharge current was shifted from the main anode to the shims at constant anode voltage. Thruster efficiency did not change. Plume divergence was reduced by approximately 4 degrees of half-angle at high levels of current on

Time-Synchronized Continuous Wave Laser Induced Fluorescence Velocity Measurements of a 600 Watt Hall Thruster

DTIC Science & Technology

2015-07-01

channel and near- field plume region of a 600 W Hall thruster operating on xenon. Results show significant fluctuations in LIF signal intensity... LIF signal intensity (corre- lated with the density of the probed excited metastable state) in time during the discharge current cycle, with the peak...fluorescence ( LIF ).1 LIF provides the opportunity to investigate plasma sources non-intrusively with higher spatial resolution (typically < 1 mm) than

Iodine Hall Thruster Propellant Feed System for a CubeSat

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Steven

2014-01-01

The components required for an in-space iodine vapor-fed Hall effect thruster propellant management system are described. A laboratory apparatus was assembled and used to produce iodine vapor and control the flow through the application of heating to the propellant reservoir and through the adjustment of the opening in a proportional flow control valve. Changing of the reservoir temperature altered the flowrate on the timescale of minutes while adjustment of the proportional flow control valve changed the flowrate immediately without an overshoot or undershoot in flowrate with the requisite recovery time associated with thermal control systems. The flowrates tested spanned a range from 0-1.5 mg/s of iodine, which is sufficient to feed a 200-W Hall effect thruster.

Hall thruster microturbulence under conditions of modified electron wall emission

NASA Astrophysics Data System (ADS)

Tsikata, S.; Héron, A.; Honoré, C.

2017-05-01

In recent numerical, theoretical, and experimental papers, the short-scale electron cyclotron drift instability (ECDI) has been studied as a possible contributor to the anomalous electron current observed in Hall thrusters. In this work, features of the instability, in the presence of a zero-electron emission material at the thruster exit plane, are analyzed using coherent Thomson scattering. Limiting the electron emission at the exit plane alters the localization of the accelerating electric field and the expected drift velocity profile, which in turn modifies the amplitude and localization of the ECDI. The resulting changes to the standard thruster operation are expected to favor an increased contribution by the ECDI to electron current. Such an operation is associated with a degradation of thruster performance and stability.

Eight Kilowatt Hall Thruster System Characterization

DTIC Science & Technology

2013-08-01

divergence13,14 and maximize thrust efficiency. Electrical isolation of the anode from the propellant line is achieved with a custom high voltage ceramic ... Electric Co., Inc. 55 A, 730 V (40 kW) (+) (-) Hall Effect Thruster Hollow Cathode b h k Figure 4. T8 test facility and nominal power supply diagram...may be 700-V or higher. Electrical isolation of the anode from the propellant line is provided by a custom ceramic break developed by Busek. The

High-Power Krypton Hall Thruster Technology Being Developed for Nuclear-Powered Applications

NASA Technical Reports Server (NTRS)

Jacobson, David T.; Manzella, David H.

2004-01-01

The NASA Glenn Research Center has been performing research and development of moderate specific impulse, xenon-fueled, high-power Hall thrusters for potential solar electric propulsion applications. These applications include Mars missions, reusable tugs for low-Earth-orbit to geosynchronous-Earth-orbit transportation, and missions that require transportation to libration points. This research and development effort resulted in the design and fabrication of the NASA-457M Hall thruster that has been tested at input powers up to 95 kW. During project year 2003, NASA established Project Prometheus to develop technology in the areas of nuclear power and propulsion, which are enabling for deep-space science missions. One of the Project-Prometheus-sponsored Nuclear Propulsion Research tasks is to investigate alternate propellants for high-power Hall thruster electric propulsion. The motivation for alternate propellants includes the disadvantageous cost and availability of xenon propellant for extremely large scale, xenon-fueled propulsion systems and the potential system performance benefits of using alternate propellants. The alternate propellant krypton was investigated because of its low cost relative to xenon. Krypton propellant also has potential performance benefits for deep-space missions because the theoretical specific impulse for a given voltage is 20 percent higher than for xenon because of krypton's lower molecular weight. During project year 2003, the performance of the high-power NASA-457M Hall thruster was measured using krypton as the propellant at power levels ranging from 6.4 to 72.5 kW. The thrust produced ranged from 0.3 to 2.5 N at a discharge specific impulse up to 4500 sec.

Integration Tests of the 4 kW-Class High Voltage Hall Accelerator Power Processing Unit with the HiVHAc and the SPT-140 Hall Effect Thrusters

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Pinero, Luis; Haag, Thomas; Huang, Wensheng; Ahern, Drew; Liang, Ray; Shilo, Vlad

2016-01-01

NASA's Science Mission Directorate is sponsoring the development of a 4 kW-class Hall propulsion system for implementation in NASA science and exploration missions. The main components of the system include the High Voltage Hall Accelerator (HiVHAc), an engineering model power processing unit (PPU) developed by Colorado Power Electronics, and a xenon flow control module (XFCM) developed by VACCO Industries. NASA Glenn Research Center is performing integrated tests of the Hall thruster propulsion system. This paper presents results from integrated tests of the PPU and XFCM with the HiVHAc engineering development thruster and a SPT-140 thruster provided by Space System Loral. The results presented in this paper demonstrate thruster discharge initiation along with open-loop and closed-loop control of the discharge current with anode flow for both the HiVHAc and the SPT-140 thrusters. Integrated tests with the SPT-140 thruster indicated that the PPU was able to repeatedly initiate the thruster's discharge, achieve steady state operation, and successfully throttle the thruster between 1.5 and 4.5 kW. The measured SPT-140 performance was identical to levels reported by Space Systems Loral.

Integration Tests of the 4 kW-class High Voltage Hall Accelerator Power Processing Unit with the HiVHAc and the SPT-140 Hall Effect Thrusters

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Pinero, Luis; Haag, Thomas; Huang, Wensheng; Ahern, Drew; Liang, Ray; Shilo, Vlad

2016-01-01

NASAs Science Mission Directorate is sponsoring the development of a 4 kW-class Hall propulsion system for implementation in NASA science and exploration missions. The main components of the system include the High Voltage Hall Accelerator (HiVHAc), an engineering model power processing unit (PPU) developed by Colorado Power Electronics, and a xenon flow control module (XFCM) developed by VACCO Industries. NASA Glenn Research Center is performing integrated tests of the Hall thruster propulsion system. This presentation presents results from integrated tests of the PPU and XFCM with the HiVHAc engineering development thruster and a SPT-140 thruster provided by Space System Loral. The results presented in this paper demonstrate thruster discharge initiation, open-loop and closed-loop control of the discharge current with anode flow for both the HiVHAc and the SPT-140 thrusters. Integrated tests with the SPT-140 thruster indicated that the PPU was able to repeatedly initiate the thrusters discharge, achieve steady state operation, and successfully throttle the thruster between 1.5 and 4.5 kW. The measured SPT-140 performance was identical to levels reported by Space Systems Loral.

Performance Evaluation of a 50kW Hall Thruster

NASA Technical Reports Server (NTRS)

Jacobson, David T.; Jankovsky, Robert S.

1999-01-01

An experimental investigation was conducted on a laboratory model Hall thruster designed to operate at power levels up to 50 kW. During this investigation the engine's performance was characterized over a range of discharge currents from 10 to 36 A and a range of discharge voltages from 200 to 800 V Operating on the Russian cathode a maximum thrust of 966 mN was measured at 35.6 A and 713.0 V. This corresponded to a specific impulse of 3325 s and an efficiency of 62%. The maximum power the engine was operated at was 25 kW. Additional testing was conducted using a NASA cathode designed for higher current operation. During this testing, thrust over 1 N was measured at 40.2 A and 548.9 V. Several issues related to operation of Hall thrusters at these high powers were encountered.

Hybrid-PIC simulation of sputtering product distribution in a Hall thruster

NASA Astrophysics Data System (ADS)

Cao, Xifeng; Hang, Guanrong; Liu, Hui; Meng, Yingchao; Luo, Xiaoming; Yu, Daren

2017-10-01

Hall thrusters have been widely used in orbit correction and the station-keeping of geostationary satellites due to their high specific impulse, long life, and high reliability. During the operating life of a Hall thruster, high-energy ions will bombard the discharge channel and cause serious erosion. As time passes, this sputtering process will change the macroscopic surface morphology of the discharge channel, especially near the exit, thus affecting the performance of the thruster. Therefore, it is necessary to carry out research on the motion of the sputtering products and erosion process of the discharge wall. To better understand the moving characteristics of sputtering products, based on the hybrid particle-in-cell (PIC) numerical method, this paper simulates the different erosion states of the thruster discharge channel in different moments and analyzes the moving process of different particles, such as B atoms and B+ ions. In this paper, the main conclusion is that B atoms are mainly produced on both sides of the channel exit, and B+ ions are mainly produced in the middle of the channel exit. The ionization rate of B atoms is approximately 1%.

Modeling an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT)

NASA Technical Reports Server (NTRS)

Choi, Maria

2016-01-01

An iodine-operated 200-W Hall thruster plume has been simulated using a hybrid-PIC model to predict the spacecraft surface-plume interaction for spacecraft integration purposes. For validation of the model, the plasma potential, electron temperature, ion current flux, and ion number density of xenon propellant were compared with available measurement data at the nominal operating condition. To simulate iodine plasma, various collision cross sections were found and used in the model. While time-varying atomic iodine species (i.e., I, I+, I2+) information is provided by HPHall simulation at the discharge channel exit, the molecular iodine species (i.e., I2, I2+) are introduced as Maxwellian particles at the channel exit. Simulation results show that xenon and iodine plasma plumes appear to be very similar under the assumptions of the model. Assuming a sticking coefficient of unity, iodine deposition rate is estimated.

Investigation of discharge channel wall material influence on lifetime of hall effect thruster with high specific impulse

NASA Astrophysics Data System (ADS)

Abashkin, V. V.; Belikov, M. B.; Gorshkov, O. A.; Lovtsov, A. S.; Khrapach, I. N.

2011-10-01

Results of 500-hour life tests of the 900-watt Hall-thruster laboratory model with the specific impulse of 2000 s are presented. The thruster discharge channel walls were manufactured from 60% BN + 40% SiO2 and >90% BN hot-pressed ceramics. The predicted total lifetime was ˜3000 h for both wall materials in spite of greater erosion resistance of pure BN in comparison with BN-SiO2 mixture. To clarify the accompanying phenomena, the following diagnostics were carried out. The surface microstructure and composition insulators were investigated by means of electron microscopy and X-ray fluorescence analysis and nearwall plasma parameters were measured with flat Langmuir probes. The obtained distributions of plasma parameters were compared with the results of stationary one-dimensional (1D) hydrodynamic modeling of discharge channel.

Direct Drive Hall Thruster System Development

NASA Technical Reports Server (NTRS)

Hoskins, W. Andrew; Homiak, Daniel; Cassady, R. Joseph; Kerslake, Tom; Peterson, Todd; Ferguson, Dale; Snyder, Dave; Mikellides, Ioannis; Jongeward, Gary; Schneider, Todd

2003-01-01

The sta:us of development of a Direct Drive Ha!! Thruster System is presented. 13 the first part. a s:udy of the impacts to spacecraft systems and mass benefits of a direct-drive architecture is reviewed. The study initially examines four cases of SPT-100 and BPT-4000 Hall thrusters used for north-south station keeping on an EXPRESS-like geosynchronous spacecraft and for primary propulsion for a Deep Space- 1 based science spacecraft. The study is also extended the impact of direct drive on orbit raising for higher power geosynchronous spacecraft and on other deep space missions as a function of power and delta velocity. The major system considerations for accommodating a direct drive Hall thruster are discussed, including array regulation, system grounding, distribution of power to the spacecraft bus, and interactions between current-voltage characteristics for the arrays and thrusters. The mass benefit analysis shows that, for the initial cases, up to 42 kg of dry mass savings is attributable directly to changes in the propulsion hardware. When projected mass impacts of operating the arrays and the electric power system at 300V are included, up to 63 kg is saved for the four initial cases. Adoption of high voltage lithium ion battery technology is projected to further improve these savings. Orbit raising of higher powered geosynchronous spacecraft, is the mission for which direct drive provides the most benefit, allowing higher efficiency electric orbit raising to be accomplished in a limited period of time, as well as nearly eliminating significant power processing heat rejection mass. The total increase in useful payload to orbit ranges up to 278 kg for a 25 kW spacecraft, launched from an Atlas IIA. For deep space missions, direct drive is found to be most applicable to higher power missions with delta velocities up to several km/s , typical of several Discovery-class missions. In the second part, the status of development of direct drive propulsion power

Plasma Interactions with High Voltage Solar Arrays for a Direct Drive Hall Effect Thruster System

NASA Technical Reports Server (NTRS)

Schneider, T.; Horvater, M. A.; Vaughn, J.; Carruth, M. R.; Jongeward, G. A.; Mikellides, I. G.

2003-01-01

The Environmental Effects Group of NASA s Marshall Space Flight Center (MSFC) is conducting research into the effects of plasma interaction with high voltage solar arrays. These high voltage solar arrays are being developed for a direct drive Hall Effect Thruster propulsion system. A direct drive system configuration will reduce power system mass by eliminating a conventional power-processing unit. The Environmental Effects Group has configured two large vacuum chambers to test different high-voltage array concepts in a plasma environment. Three types of solar arrays have so far been tested, an International Space Station (ISS) planar array, a Tecstar planar array, and a Tecstar solar concentrator array. The plasma environment was generated using a hollow cathode plasma source, which yielded densities between 10(exp 6) - 10(exp 7) per cubic centimeter and electron temperatures of 0.5-1 eV. Each array was positioned in this plasma and biased in the -500 to + 500 volt range. The current collection was monitored continuously. In addition, the characteristics of arcing, snap over, and other features, were recorded. Analysis of the array performance indicates a time dependence associated with the current collection as well as a tendency for "conditioning" over a large number of runs. Mitigation strategies, to reduce parasitic current collection, as well as arcing, include changing cover-glass geometry and layout as well as shielding the solar cell edges. High voltage performance data for each of the solar array types tested will be presented. In addition, data will be provided to indicate the effectiveness of the mitigation techniques.

Modulating action of low frequency oscillations on high frequency instabilities in Hall thrusters

DOE Office of Scientific and Technical Information (OSTI.GOV)

Liqiu, Wei, E-mail: weiliqiu@gmail.com, E-mail: weiliqiu@hit.edu.cn; Liang, Han; Ziyi, Yang

2015-02-07

It is found that the low frequency oscillations have modulating action on high frequency instabilities in Hall thrusters. The physical mechanism of this modulation is discussed and verified by numerical simulations. Theoretical analyses indicate that the wide-range fluctuations of plasma density and electric field associated with the low frequency oscillations affect the electron drift velocity and anomalous electron transport across the magnetic field. The amplitude and frequency of high frequency oscillations are modulated by low frequency oscillations, which show the periodic variation in the time scale of low frequency oscillations.

Integration Testing of a Modular Discharge Supply for NASA's High Voltage Hall Accelerator Thruster

NASA Technical Reports Server (NTRS)

Pinero, Luis R.; Kamhawi, hani; Drummond, Geoff

2010-01-01

NASA s In-Space Propulsion Technology Program is developing a high performance Hall thruster that can fulfill the needs of future Discovery-class missions. The result of this effort is the High Voltage Hall Accelerator thruster that can operate over a power range from 0.3 to 3.5 kW and a specific impulse from 1,000 to 2,800 sec, and process 300 kg of xenon propellant. Simultaneously, a 4.0 kW discharge power supply comprised of two parallel modules was developed. These power modules use an innovative three-phase resonant topology that can efficiently supply full power to the thruster at an output voltage range of 200 to 700 V at an input voltage range of 80 to 160 V. Efficiencies as high as 95.9 percent were measured during an integration test with the NASA103M.XL thruster. The accuracy of the master/slave current sharing circuit and various thruster ignition techniques were evaluated.

Ion Species Fractions in the Far-Field Plume of a High-Specific Impulse Hall Thruster

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Gallimore, Alec D.

2003-01-01

An ExB probe was used to measure the ion species fractions of Xe(+), Xe(2+), and Xe(3+) in the far-field plume of the NASA-173Mv2 laboratory-model Hall thruster. The thruster was operated at a constant xenon flow rate of 10 milligrams per second and discharge voltages of 300 to 900 V. The ExB probe was placed two meters downstream of the thruster exit plane on the thruster centerline. At a discharge voltage of 300 V, the species fractions of Xe(2+) and Xe(3+) were lower, but still consistent with, previous Hall thruster studies using other mass analyzers. Over discharge voltages of 300 to 900 V, the Xe(2+) species fractions increased from 0.04 to 0.12 and the Xe(3+) species fraction increased from 0.01 to 0.02.

Laser-Induced Fluorescence Velocity Measurements of a Low Power Cylindrical Hall Thruster

DTIC Science & Technology

2009-08-25

Hall thruster . Xenon ion velocities for the thruster are derived from laser-induced fluorescence measurements of the 5d[4]7/2-6p[3]5/2 xenon ion excited state transition. Three operating conditions are considered with variations to the magnetic field strength and chamber background pressure in an effort to capture their effects on ion acceleration and centerline ion energy distributions. Under nominal conditions, xenon ions are accelerated to an energy of 25 eV within the thruster with an additional 188 eV gain in the thruster plume. At a position 40 mm into the plume,

Confinement effect of cylindrical-separatrix-type magnetic field on the plume of magnetic focusing type Hall thruster

NASA Astrophysics Data System (ADS)

Yu, Daren; Meng, Tianhang; Ning, Zhongxi; Liu, Hui

2017-04-01

A magnetic focusing type Hall thruster was designed with a cylindrical magnetic seperatrix. During the process of a hollow cathode crossing the separatrix, the variance of plume parameter distribution was monitored. Results show that the ion flux on the large spatial angle is significantly lower when the hollow cathode is located in the inner magnetic field. This convergence effect is preserved even in a distant area. A mechanism was proposed for plume divergence from the perspective of cathode-to-plume potential difference, through which the confinement effect of cylindrical-separatrix-type magnetic field on thruster plume was confirmed and proposed as a means of plume protection for plasma propulsion devices.

High-Efficiency Hall Thruster Discharge Power Converter

NASA Technical Reports Server (NTRS)

Jaquish, Thomas

2015-01-01

Busek Company, Inc., is designing, building, and testing a new printed circuit board converter. The new converter consists of two series or parallel boards (slices) intended to power a high-voltage Hall accelerator (HiVHAC) thruster or other similarly sized electric propulsion devices. The converter accepts 80- to 160-V input and generates 200- to 700-V isolated output while delivering continually adjustable 300-W to 3.5-kW power. Busek built and demonstrated one board that achieved nearly 94 percent efficiency the first time it was turned on, with projected efficiency exceeding 97 percent following timing software optimization. The board has a projected specific mass of 1.2 kg/kW, achieved through high-frequency switching. In Phase II, Busek optimized to exceed 97 percent efficiency and built a second prototype in a form factor more appropriate for flight. This converter then was integrated with a set of upgraded existing boards for powering magnets and the cathode. The program culminated with integrating the entire power processing unit and testing it on a Busek thruster and on NASA's HiVHAC thruster.

Evaluating the accuracy of recent electron transport models at predicting Hall thruster plasma dynamics

NASA Astrophysics Data System (ADS)

Cappelli, Mark; Young, Christopher

2016-10-01

We present continued efforts towards introducing physical models for cross-magnetic field electron transport into Hall thruster discharge simulations. In particular, we seek to evaluate whether such models accurately capture ion dynamics, both averaged and resolved in time, through comparisons with measured ion velocity distributions which are now becoming available for several devices. Here, we describe a turbulent electron transport model that is integrated into 2-D hybrid fluid/PIC simulations of a 72 mm diameter laboratory thruster operating at 400 W. We also compare this model's predictions with one recently proposed by Lafluer et al.. Introducing these models into 2-D hybrid simulations is relatively straightforward and leverages the existing framework for solving the electron fluid equations. The models are tested for their ability to capture the time-averaged experimental discharge current and its fluctuations due to ionization instabilities. Model predictions are also more rigorously evaluated against recent laser-induced fluorescence measurements of time-resolved ion velocity distributions.

Ion Current Density Study of the NASA-300M and NASA-457Mv2 Hall Thrusters

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2012-01-01

NASA Glenn Research Center is developing a Hall thruster in the 15-50 kW range to support future NASA missions. As a part of the process, the performance and plume characteristics of the NASA-300M, a 20-kW Hall thruster, and the NASA-457Mv2, a 50-kW Hall thruster, were evaluated. The collected data will be used to improve the fidelity of the JPL modeling tool, Hall2De, which will then be used to aid the design of the 15-50 kW Hall thruster. This paper gives a detailed overview of the Faraday probe portion of the plume characterization study. The Faraday probe in this study is a near-field probe swept radially at many axial locations downstream of the thruster exit plane. Threshold-based integration limits with threshold values of 1/e, 1/e2, and 1/e3 times the local peak current density are tried for the purpose of ion current integration and divergence angle calculation. The NASA-300M is operated at 7 conditions and the NASA-457Mv2 at 14 conditions. These conditions span discharge voltages of 200 to 500 V and discharge power of 10 to 50 kW. The ion current density profiles of the near-field plume originating from the discharge channel are discovered to strongly resemble Gaussian distributions. A novel analysis approach involving a form of ray tracing is used to determine an effective point of origin for the near-field plume. In the process of performing this analysis, definitive evidence is discovered that showed the near-field plume is bending towards the thruster centerline.

Ion Current Density Study of the NASA-300M and NASA-457Mv2 Hall Thrusters

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2012-01-01

NASA Glenn Research Center is developing a Hall thruster in the 15-50 kW range to support future NASA missions. As a part of the process, the performance and plume characteristics of the NASA-300M, a 20-kW Hall thruster, and the NASA-457Mv2, a 50-kW Hall thruster, were evaluated. The collected data will be used to improve the fidelity of the JPL modeling tool, Hall2De, which will then be used to aid the design of the 15-50 kW Hall thruster. This paper gives a detailed overview of the Faraday probe portion of the plume characterization study. The Faraday probe in this study is a near-field probe swept radially at many axial locations downstream of the thruster exit plane. Threshold-based integration limits with threshold values of 1/e, 1/e(sup 2), and 1/e(sup 3) times the local peak current density are tried for the purpose of ion current integration and divergence angle calculation. The NASA-300M is operated at 7 conditions and the NASA-457Mv2 at 14 conditions. These conditions span discharge voltages of 200 to 500 V and discharge power of 10 to 50 kW. The ion current density profiles of the near-field plume originating from the discharge channel are discovered to strongly resemble Gaussian distributions. A novel analysis approach involving a form of ray tracing is used to determine an effective point of origin for the near-field plume. In the process of performing this analysis, definitive evidence is discovered that showed the near-field plume is bending towards the thruster centerline.

Non-Maxwellian electron energy probability functions in the plume of a SPT-100 Hall thruster

NASA Astrophysics Data System (ADS)

Giono, G.; Gudmundsson, J. T.; Ivchenko, N.; Mazouffre, S.; Dannenmayer, K.; Loubère, D.; Popelier, L.; Merino, M.; Olentšenko, G.

2018-01-01

We present measurements of the electron density, the effective electron temperature, the plasma potential, and the electron energy probability function (EEPF) in the plume of a 1.5 kW-class SPT-100 Hall thruster, derived from cylindrical Langmuir probe measurements. The measurements were taken on the plume axis at distances between 550 and 1550 mm from the thruster exit plane, and at different angles from the plume axis at 550 mm for three operating points of the thruster, characterized by different discharge voltages and mass flow rates. The bulk of the electron population can be approximated as a Maxwellian distribution, but the measured distributions were seen to decline faster at higher energy. The measured EEPFs were best modelled with a general EEPF with an exponent α between 1.2 and 1.5, and their axial and angular characteristics were studied for the different operating points of the thruster. As a result, the exponent α from the fitted distribution was seen to be almost constant as a function of the axial distance along the plume, as well as across the angles. However, the exponent α was seen to be affected by the mass flow rate, suggesting a possible relationship with the collision rate, especially close to the thruster exit. The ratio of the specific heats, the γ factor, between the measured plasma parameters was found to be lower than the adiabatic value of 5/3 for each of the thruster settings, indicating the existence of non-trivial kinetic heat fluxes in the near collisionless plume. These results are intended to be used as input and/or testing properties for plume expansion models in further work.

Multi-Kilowatt Power Module for High-Power Hall Thrusters

NASA Technical Reports Server (NTRS)

Pinero, Luis R.; Bowers, Glen E.

2005-01-01

Future NASA missions will require high-performance electric propulsion systems. Hall thrusters are being developed at NASA Glenn for high-power, high-specific impulse operation. These thrusters operate at power levels up to 50 kW of power and discharge voltages in excess of 600 V. A parallel effort is being conducted to develop power electronics for these thrusters that push the technology beyond the 5kW state-of-the-art power level. A 10 kW power module was designed to produce an output of 500 V and 20 A from a nominal 100 V input. Resistive load tests revealed efficiencies in excess of 96 percent. Load current share and phase synchronization circuits were designed and tested that will allow connecting multiple modules in parallel to process higher power.

Hall Effect Thruster Plume Contamination and Erosion Study

NASA Technical Reports Server (NTRS)

Jaworske, Donald A.

2000-01-01

The objective of the Hall effect thruster plume contamination and erosion study was to evaluate the impact of a xenon ion plume on various samples placed in the vicinity of a Hall effect thruster for a continuous 100 hour exposure. NASA Glenn Research Center was responsible for the pre- and post-test evaluation of three sample types placed around the thruster: solar cell cover glass, RTV silicone, and Kapton(R). Mass and profilometer), were used to identify the degree of deposition and/or erosion on the solar cell cover glass, RTV silicone, and Kapton@ samples. Transmittance, reflectance, solar absorptance, and room temperature emittance were used to identify the degree of performance degradation of the solar cell cover glass samples alone. Auger spectroscopy was used to identify the chemical constituents found on the surface of the exposed solar cell cover glass samples. Chemical analysis indicated some boron nitride contamination on the samples, from boron nitride insulators used in the body of the thruster. However, erosion outweighted contamination. All samples exhibited some degree of erosion. with the most erosion occurring near the centerline of the plume and the least occurring at the +/- 90 deg positions. For the solar cell cover glass samples, erosion progressed through the antireflective coating and into the microsheet glass itself. Erosion occurred in the solar cell cover glass, RTV silicone and Kapton(R) at different rates. All optical properties changed with the degree of erosion, with solar absorptance and room temperature emittance increasing with erosion. The transmittance of some samples decreased while the reflectance of some samples increased and others decreased. All results are consistent with an energetic plume of xenon ions serving as a source for erosion.

Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Chang, Li; Clayman, Lauren; Herman, Daniel; Shastry, Rohit; Thomas, Robert; Verhey, Timothy;

Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Chang, Li; Clayman, Lauren; Herman, Daniel; Shastry, Rohit; Thomas, Robert; Verhey, Timothy;

Effect of low-frequency oscillation on performance of Hall thrusters

NASA Astrophysics Data System (ADS)

Liqiu, WEI; Wenbo, LI; Yongjie, DING; Daren, YU

2018-07-01

In this paper, a direct connection between the discharge current amplitude and the thruster performance is established by varying solely the capacitance of the filter unit of the Hall thrusters. To be precise, the variation characteristics of ion current, propellant utilization efficiency, and divergence angle of plume at different low-frequency oscillation amplitudes are measured. The findings demonstrate that in the case of the propellant in the discharge channel just meets or falls below the full ionization condition, the increase of low-frequency oscillation amplitude can significantly enhance the ionization degree of the neutral gas in the channel and increase the thrust and anode efficiency of thruster. On the contrary, the increase in the amplitude of low-frequency oscillation will lead to increase the loss of plume divergence, therefore the thrust and anode efficiency of thruster decrease.

Sputtering Erosion Measurement on Boron Nitride as a Hall Thruster Material

NASA Technical Reports Server (NTRS)

Britton, Melissa; Waters, Deborah; Messer, Russell; Sechkar, Edward; Banks, Bruce

2002-01-01

The durability of a high-powered Hall thruster may be limited by the sputter erosion resistance of its components. During normal operation, a small fraction of the accelerated ions will impact the interior of the main discharge channel, causing its gradual erosion. A laboratory experiment was conducted to simulate the sputter erosion of a Hall thruster. Tests of sputter etch rate were carried out using 300 to 1000 eV Xenon ions impinging on boron nitride substrates with angles of attack ranging from 30 to 75 degrees from horizontal. The erosion rates varied from 3.41 to 14.37 Angstroms/[sec(mA/sq cm)] and were found to depend on the ion energy and angle of attack, which is consistent with the behavior of other materials.

Effect of segmented electrode length on the performances of Hall thruster

NASA Astrophysics Data System (ADS)

Duan, Ping; Chen, Long; Liu, Guangrui; Bian, Xingyu; Yin, Yan

2016-09-01

The influences of the low-emissive graphite segmented electrode placed near the channel exit on the discharge characteristics of Hall thruster are studied using the particle-in-cell method. A two-dimensional physical model is established according to the Hall thruster discharge channel configuration. The effects of electrode length on potential, ion density, electron temperature, ionization rate and discharge current are investigated. It is found that, with the increasing of segmented electrode length, the equipotential lines bend towards the channel exit, and approximately parallel to the wall at the channel surface, radial velocity and radial flow of ions are increased, and the electron temperature is also enhanced. Due to the conductive characteristic of electrodes, the radial electric field and the axial electron conductivity near the wall are enhanced, and the probability of the electron-atom ionization is reduced, which leads to the degradation of ionization rate in discharge channel. However, the interaction between electrons and the wall enhances the near wall conductivity, therefore the discharge current grows along with the segmented electrode length, and the performance of the thruster is also affected.

Modeling Electron Transport within the Framework of Hydrodynamic Description of Hall Thrusters (Preprint)

DTIC Science & Technology

2008-06-16

Framework of Hydrodynamic Description of Hall Thrusters (Preprint) 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) M . keidar (The George...within the framework of hydrodynamic description of Hall thrusters (PREPRINT) M . Keidar 1 and L. Brieda 2 1) Department of Mechanical and...the electron current density: y yw y m ew y w z w ew dV V y kT mV kT e kT e kT m B E nj y )sin() 2 exp()exp()exp( 2 2 2 2 2/1 0 (2) In this case, a

Ion Velocity Measurements in a Linear Hall Thruster (Postprint)

DTIC Science & Technology

2005-06-14

Hall Thruster in a high vacuum environment. The ionized propellant velocities were measured using laser induced fluorescence of the excited state xenon ionic transition at 834.7 nm. Ion velocities were interrogated from the channel exit plane to a distance 30 mm from it. Both axial and cross-field (along the electron Hall current direction) velocities were measured. The results presented here, combined with those of previous work, highlight the high sensitivity of electron mobility inside and outside the channel, depending on the background gas density, type of wall

4.5-kW Hall Effect Thruster Evaluated

NASA Technical Reports Server (NTRS)

Mason, Lee S.

2000-01-01

As part of an Interagency Agreement with the Air Force Research Lab (AFRL), a space simulation test of a Russian SPT 140 Hall Effect Thruster was completed in September 1999 at Vacuum Facility 6 at the NASA Glenn Research Center at Lewis Field. The thruster was subjected to a three-part test sequence that included thrust and performance characterization, electromagnetic interference, and plume contamination. SPT 140 is a 4.5-kW thruster developed under a joint agreement between AFRL, Atlantic Research Corp, and Space Systems/Loral, and was manufactured by the Fakal Experimental Design Bureau of Russia. All objectives were satisfied, and the thruster performed exceptionally well during the 120-hr test program, which comprised 33 engine firings. The Glenn testing provided a critical contribution to the thruster development effort, and the large volume and high pumping speed of this vacuum facility was key to the test s success. The low background pressure (1 10 6 torr) provided a more accurate representation of space vacuum than is possible in most vacuum chambers. The facility had been upgraded recently with new cryogenic pumps and sputter shielding to support the active electric propulsion program at Glenn. The Glenn test team was responsible for all test support equipment, including the thrust stand, power supplies, data acquisition, electromagnetic interference measurement equipment, and the contamination measurement system.

Stationary plasma thruster evaluation in Russia

NASA Technical Reports Server (NTRS)

Brophy, John R.

1992-01-01

A team of electric propulsion specialists from U.S. government laboratories experimentally evaluated the performance of a 1.35-kW Stationary Plasma Thruster (SPT) at the Scientific Research Institute of Thermal Processes in Moscow and at 'Fakel' Enterprise in Kaliningrad, Russia. The evaluation was performed using a combination of U.S. and Russian instrumentation and indicated that the actual performance of the thruster appears to be close to the claimed performance. The claimed performance was a specific impulse of 16,000 m/s, an overall efficiency of 50 percent, and an input power of 1.35 kW, and is superior to the performance of western electric thrusters at this specific impulse. The unique performance capabilities of the stationary plasma thruster, along with claims that more than fifty of the 660-W thrusters have been flown in space on Russian spacecraft, attracted the interest of western spacecraft propulsion specialists. A two-phase program was initiated to evaluate the stationary plasma thruster performance and technology. The first phase of this program, to experimentally evaluate the performance of the thruster with U.S. instrumentation in Russia, is described in this report. The second phase objective is to determine the suitability of the stationary plasma thruster technology for use on western spacecraft. This will be accomplished by bringing stationary plasma thrusters to the U.S. for quantification of thruster erosion rates, measurements of the performance variation as a function of long-duration operation, quantification of the exhaust beam divergence angle, and determination of the non-propellant efflux from the thruster. These issues require quantification in order to maximize the probability for user application of the SPT technology and significantly increase the propulsion capabilities of U.S. spacecraft.

Effect of vortex inlet mode on low-power cylindrical Hall thruster

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Jia, Boyang; Xu, Yu; Wei, Liqiu; Su, Hongbo; Li, Peng; Sun, Hezhi; Peng, Wuji; Cao, Yong; Yu, Daren

2017-08-01

This paper examines a new propellant inlet mode for a low-power cylindrical Hall thruster called the vortex inlet mode. This new mode makes propellant gas diffuse in the form of a circumferential vortex in the discharge channel of the thruster. Simulation and experimental results show that the neutral gas density in the discharge channel increases upon the application of the vortex inlet mode, effectively extending the dwell time of the propellant gas in the channel. According to the experimental results, the vortex inlet increases the propellant utilization of the thruster by 3.12%-8.81%, thrust by 1.1%-53.5%, specific impulse by 1.1%-53.5%, thrust-to-power ratio by 10%-63%, and anode efficiency by 1.6%-7.3%, greatly improving the thruster performance.

Magnetic circuit for hall effect plasma accelerator

NASA Technical Reports Server (NTRS)

Manzella, David H. (Inventor); Jacobson, David T. (Inventor); Hofer, Richard (Inventor); Peterson, Peter (Inventor); Jankovsky, Robert S. (Inventor)

2009-01-01

A Hall effect plasma accelerator includes inner and outer electromagnets, circumferentially surrounding the inner electromagnet along a thruster centerline axis and separated therefrom, inner and outer magnetic conductors, in physical connection with their respective inner and outer electromagnets, with the inner magnetic conductor having a mostly circular shape and the outer magnetic conductor having a mostly annular shape, a discharge chamber, located between the inner and outer magnetic conductors, a magnetically conducting back plate, in magnetic contact with the inner and outer magnetic conductors, and a combined anode electrode/gaseous propellant distributor, located at a bottom portion of the discharge chamber. The inner and outer electromagnets, the inner and outer magnetic conductors and the magnetically conducting back plate form a magnetic circuit that produces a magnetic field that is largely axial and radially symmetric with respect to the thruster centerline.

Evaluation of a Magnetically-Filtered Faraday Probe for Measuring the ion Current Density Profile of a Hall Thruster

DTIC Science & Technology

2004-07-01

The ability of a magnetically-filtered Faraday probe (MFFP) to obtain the ion current density profile of a Hall thruster is investigated. The MFFP is...MFFP, boxed Faraday probe (BFP), and nude Faraday probe are used to measure the ion current density profile of a 5 kW Hall thruster operated over the

Development of a Miniature Low Power Cylindrical Hall Thruster for Microsatellites

NASA Astrophysics Data System (ADS)

Pigeon, Carl

To enable more advanced commercial microsatellite missions, a low power electric propulsion system was designed by the University of Toronto Space Flight Laboratory. A prototype cylindrical Hall thruster was first developed using electromagnets. The thruster's performance was evaluated over a range of 20-300 W. At the nominal 200 W operation, 6.2 mN of thrust with a specific impulse of 1139 s was measured with xenon propellant. Significant erosion of the thruster's discharge chamber wall was observed which limited its lifetime to 100 hours. Subsequently, a flight representative version of the thruster was developed. Permanent magnets were used to reduce the size, mass, and power consumption. Changes to the design were implemented to improve lifetime. Performance characterization and literature suggest that a reduction in performance is expected with the use of permanent magnets. Lastly, thermal vacuum and vibration tests were performed to bring the thruster to Technology Readiness Level 6.

Performance Test Results of the NASA-457M v2 Hall Thruster

NASA Technical Reports Server (NTRS)

Soulas, George C.; Haag, Thomas W.; Herman, Daniel A.; Huang, Wensheng; Kamhawi, Hani; Shastry, Rohit

2012-01-01

Performance testing of a second generation, 50 kW-class Hall thruster labeled NASA-457M v2 was conducted at the NASA Glenn Research Center. This NASA-designed thruster is an excellent candidate for a solar electric propulsion system that supports human exploration missions. Thruster discharge power was varied from 5 to 50 kW over discharge voltage and current ranges of 200 to 500 V and 15 to 100 A, respectively. Anode efficiencies varied from 0.56 to 0.71. The peak efficiency was similar to that of other state-of-the-art high power Hall thrusters, but outperformed these thrusters at lower discharge voltages. The 0.05 to 0.18 higher anode efficiencies of this thruster compared to its predecessor were primarily due to which of two stable discharge modes the thruster was operated. One stable mode was at low magnetic field strengths, which produced high anode efficiencies, and the other at high magnetic fields where its predecessor was operated. Cathode keeper voltages were always within 2.1 to 6.2 V and cathode voltages were within 13 V of tank ground during high anode efficiency operation. However, during operation at high magnetic fields, cathode-to-ground voltage magnitudes increased dramatically, exceeding 30 V, due to the high axial magnetic field strengths in the immediate vicinity of the centrally-mounted cathode. The peak thrust was 2.3 N and this occurred at a total thruster input power of 50.0 kW at a 500 V discharge voltage. The thruster demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 46.1 mN/kW at full power, and a specific impulse range of 1420 to 2740 s. For a discharge voltage of 300 V, where specific impulses would be about 2000 s, thrust efficiencies varied from 0.57 to 0.63.

Application of hollow anodes in a Hall thruster with double-peak magnetic fields

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Sun, Hezhi; Li, Peng; Wei, Liqiu; Su, Hongbo; Peng, Wuji; Li, Hong; Yu, Daren

2017-08-01

A low-power Hall thruster was designed with two permanent magnet rings. Unlike conventional Hall thrusters, this one has a symmetrical double-peak magnetic field with a larger gradient. Moreover, the highest magnetic field strength appears in the plume region; hence, the distance from the zero-magnetic region to the channel outlet is shorter than that of other Hall thrusters. This paper presents the law and mechanism of the effect of a U-shaped hollow anode with the front end in the zero-magnetic region and anodes at the first magnetic peak and zero-magnetic point (corresponding to the front and rear end faces of the U-shaped anode, respectively) on the discharge characteristics of the thruster. The study shows that the overall performance of the hollow anode under the same operating conditions is the highest. For the anode at the magnetic peak, although the ionization rate is the highest, most of the ions generated by ionization collide with the walls, causing greater energy loss and minimizing its performance. For the anode at the zero-magnetic point, although its maximum ionization rate is higher than that of the hollow anode, and the power deposition on the walls is slightly smaller, its propellant utilization and voltage utilization are lower than those of the hollow anode; furthermore, its overall performance is poorer than that of the hollow anode because of the short channel and shorter ionization region.

Design and Testing of a Hall Effect Thruster with 3D Printed Channel and Propellant Distributor

NASA Technical Reports Server (NTRS)

Hopping, Ethan P.; Xu, Kunning G.

2017-01-01

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville with channel walls and a propellant distributor manufactured using 3D printing. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. An overview of the thruster design and transient performance measurements are presented here. Measured thrust ranged from 17.2 millinewtons to 30.4 millinewtons over a discharge power of 280 watts to 520 watts with an anode I (sub SP)(Specific Impulse) range of 870 seconds to 1450 seconds. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state.

A Comprehensive Investigation of Facility Effects on the Testing of High-Power Monolithic and Clustered Hall Thruster Systems

DTIC Science & Technology

2004-09-02

path for developing high-power EP systems is somewhat certain given NASA’s recent success with its 70+ kW NASA-457M Hall thruster , it is clear that...current density distribution, and summarize findings from cold- and hot-flow pressure map data of our vacuum chamber for a number of Hall thruster mass flow rates.

High Voltage Solar Array ARC Testing for a Direct Drive Hall Effect Thruster System

NASA Technical Reports Server (NTRS)

Schneider, T.; Vaughn, J.; Carruth, M. R.; Mikellides, I. G.; Jongeward, G. A.; Peterson, T.; Kerslake, T. W.; Snyder, D.; Ferguson, D.; Hoskins, A.

2003-01-01

The deleterious effects of spacecraft charging are well known, particularly when the charging leads to arc events. The damage that results from arcing can severely reduce system lifetime and even cause critical system failures. On a primary spacecraft system such as a solar array, there is very little tolerance for arcing. Motivated by these concerns, an experimental investigation was undertaken to determine arc thresholds for a high voltage (200-500 V) solar array in a plasma environment. The investigation was in support of a NASA program to develop a Direct Drive Hall-Effect Thruster (112HET) system. By directly coupling the solar array to a Hall-effect thruster, the D2HET program seeks to reduce mass, cost and complexity commonly associated with the power processing in conventional power systems. In the investigation, multiple solar array technologies and configurations were tested. The cell samples were biased to a negative voltage, with an applied potential difference between them, to imitate possible scenarios in solar array strings that could lead to damaging arcs. The samples were tested in an environment that emulated a low-energy, HET-induced plasma. Short duration "trigger" arcs as well as long duration "sustained" arcs were generated. Typical current and voltage waveforms associated with the arc events are presented. Arc thresholds are also defined in terms of vo!tage, (current and power. The data will be used to propose a new, high-voltage (>300 V) solar array design for which the likelihood of damage from arcing is minimal.

High Voltage Solar Array Arc Testing for a Direct Drive Hall Effect Thruster System

NASA Technical Reports Server (NTRS)

Schneider, Todd; Carruth, M. R., Jr.; Vaughn, J. A.; Jongeward, G. A.; Mikellides, I. G.; Ferguson, D.; Kerslake, T. W.; Peterson, T.; Snyder, D.; Hoskins, A.

2004-01-01

The deleterious effects of spacecraft charging are well known, particularly when the charging leads to arc events. The damage that results from arcing can severely reduce system lifetime and even cause critical system failures. On a primary spacecraft system such as a solar array, there is very little tolerance for arcing. Motivated by these concerns, an experimental investigation was undertaken to determine arc thresholds for a high voltage (200-500 V) solar array in a plasma environment. The investigation was in support of a NASA program to develop a Direct Drive Hall-Effect Thruster (D2HET) system. By directly coupling the solar array to a Hall-effect thruster, the D2HET program seeks to reduce mass, cost and complexity commonly associated with the power processing in conventional power systems. In the investigation, multiple solar array technologies and configurations were tested. The cell samples were biased to a negative voltage, with an applied potential difference between them, to imitate possible scenarios in solar array strings that could lead to damaging arcs. The samples were tested in an environment that emulated a low-energy, HET-induced plasma. Short duration trigger arcs as well as long duration sustained arcs were generated. Typical current and voltage waveforms associated with the arc events are presented. Arc thresholds are also defined in terms of voltage, current and power. The data will be used to propose a new, high-voltage (greater than 300 V) solar array design for which the likelihood of damage from arcing is minimal.

Role of ionization and electron drift velocity profile to Rayleigh instability in a Hall thruster plasma

DOE Office of Scientific and Technical Information (OSTI.GOV)

Singh, Sukhmander; Malik, Hitendra K.

Role of ionization to Rayleigh instability is clarified in a Hall thruster plasma under the variety of profiles of electron drift velocity, namely, step-like profile (SLP) and two different super-Gaussian profiles (SGP1 and SGP2). For this, a relevant Rayleigh equation is derived and solved numerically using fourth-order Runge-Kutta method. Interestingly, an upper cutoff frequency of oscillations {omega}{sub max} is realized for the occurrence of the instability that shows dependence on the ionization rate {alpha}, electron drift velocity u{sub 0}, electron cyclotron frequency {Omega}, azimuthal wave number k{sub y}, plasma density n{sub 0}, density gradient {partial_derivative}n{sub 0}/{partial_derivative}x, ion (electron) thermal speedmore » V{sub thI}(V{sub thE}), and ion (electron) plasma frequency {omega}{sub pi}({omega}{sub pe}). The frequency {omega}{sub max} follows the trend {omega}{sub max} (for SGP2) >{omega}{sub max} (for SLP) >{omega}{sub max} (for SGP1) and shows a similar behaviour with ionization for all types of the velocity profiles. The instability is found to grow faster for the higher {alpha} and the ion temperature but it acquires lower rate under the effect of the higher electron temperature; the perturbed potential also varies in accordance with the growth rate. The electron temperature influences the growth rate and cutoff frequency less significantly in comparison with the ion temperature.« less

Helicon plasma thruster discharge model

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lafleur, T., E-mail: trevor.lafleur@lpp.polytechnique.fr

2014-04-15

By considering particle, momentum, and energy balance equations, we develop a semi-empirical quasi one-dimensional analytical discharge model of radio-frequency and helicon plasma thrusters. The model, which includes both the upstream plasma source region as well as the downstream diverging magnetic nozzle region, is compared with experimental measurements and confirms current performance levels. Analysis of the discharge model identifies plasma power losses on the radial and back wall of the thruster as the major performance reduction factors. These losses serve as sinks for the input power which do not contribute to the thrust, and which reduce the maximum plasma density andmore » hence propellant utilization. With significant radial plasma losses eliminated, the discharge model (with argon) predicts specific impulses in excess of 3000 s, propellant utilizations above 90%, and thruster efficiencies of about 30%.« less

E × B electron drift instability in Hall thrusters: Particle-in-cell simulations vs. theory

NASA Astrophysics Data System (ADS)

Boeuf, J. P.; Garrigues, L.

2018-06-01

The E × B Electron Drift Instability (E × B EDI), also called Electron Cyclotron Drift Instability, has been observed in recent particle simulations of Hall thrusters and is a possible candidate to explain anomalous electron transport across the magnetic field in these devices. This instability is characterized by the development of an azimuthal wave with wavelength in the mm range and velocity on the order of the ion acoustic velocity, which enhances electron transport across the magnetic field. In this paper, we study the development and convection of the E × B EDI in the acceleration and near plume regions of a Hall thruster using a simplified 2D axial-azimuthal Particle-In-Cell simulation. The simulation is collisionless and the ionization profile is not-self-consistent but rather is given as an input parameter of the model. The aim is to study the development and properties of the instability for different values of the ionization rate (i.e., of the total ion production rate or current) and to compare the results with the theory. An important result is that the wavelength of the simulated azimuthal wave scales as the electron Debye length and that its frequency is on the order of the ion plasma frequency. This is consistent with the theory predicting destruction of electron cyclotron resonance of the E × B EDI in the non-linear regime resulting in the transition to an ion acoustic instability. The simulations also show that for plasma densities smaller than under nominal conditions of Hall thrusters the field fluctuations induced by the E × B EDI are no longer sufficient to significantly enhance electron transport across the magnetic field, and transit time instabilities develop in the axial direction. The conditions and results of the simulations are described in detail in this paper and they can serve as benchmarks for comparisons between different simulation codes. Such benchmarks would be very useful to study the role of numerical noise (numerical

Method for analyzing E x B probe spectra from Hall thruster plumes.

PubMed

Shastry, Rohit; Hofer, Richard R; Reid, Bryan M; Gallimore, Alec D

2009-06-01

Various methods for accurately determining ion species' current fractions using E x B probes in Hall thruster plumes are investigated. The effects of peak broadening and charge exchange on the calculated values of current fractions are quantified in order to determine the importance of accounting for them in the analysis. It is shown that both peak broadening and charge exchange have a significant effect on the calculated current fractions over a variety of operating conditions, especially at operating pressures exceeding 10(-5) torr. However, these effects can be accounted for using a simple approximation for the velocity distribution function and a one-dimensional charge exchange correction model. In order to keep plume attenuation from charge exchange below 30%, it is recommended that pz < or = 2, where p is the measured facility pressure in units of 10(-5) torr and z is the distance from the thruster exit plane to the probe inlet in meters. The spatial variation of the current fractions in the plume of a Hall thruster and the error induced from taking a single-point measurement are also briefly discussed.

Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase III: Comparison of Theory with Experiment

NASA Technical Reports Server (NTRS)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.; Goebel, Dan M.

2012-01-01

A proof-of-principle effort to demonstrate a technique by which erosion of the acceleration channel in Hall thrusters of the magnetic-layer type can be eliminated has been completed. The first principles of the technique, now known as "magnetic shielding," were derived based on the findings of numerical simulations in 2-D axisymmetric geometry. The simulations, in turn, guided the modification of an existing 6-kW laboratory Hall thruster. This magnetically shielded (MS) thruster was then built and tested. Because neither theory nor experiment alone can validate fully the first principles of the technique, the objective of the 2-yr effort was twofold: (1) to demonstrate in the laboratory that the erosion rates can be reduced by >order of magnitude, and (2) to demonstrate that the near-wall plasma properties can be altered according to the theoretical predictions. This paper concludes the demonstration of magnetic shielding by reporting on a wide range of comparisons between results from numerical simulations and laboratory diagnostics. Collectively, we find that the comparisons validate the theory. Near the walls of the MS thruster, theory and experiment agree: (1) the plasma potential has been sustained at values near the discharge voltage, and (2) the electron temperature has been lowered by at least 2.5-3 times compared to the unshielded (US) thruster. Also, based on carbon deposition measurements, the erosion rates at the inner and outer walls of the MS thruster are found to be lower by at least 2300 and 1875 times, respectively. Erosion was so low along these walls that the rates were below the resolution of the profilometer. Using a sputtering yield model with an energy threshold of 25 V, the simulations predict a reduction of 600 at the MS inner wall. At the outer wall ion energies are computed to be below 25 V, for which case we set the erosion to zero in the simulations. When a 50-V threshold is used the computed ion energies are below the threshold at both

Experimental Investigations with a 5-kW-Class Laboratory Model Closed-Drifted Hall Thruster

DTIC Science & Technology

2001-01-01

Hall thruster (CDT). The project was composed of the following segments: 1) a 5-kW-class CDT (P5) was built and characterized in terms of performance and plume divergence; 2) the molecular-beam mass spectrometer (MBMS) was used to measure the ion energy distribution finction and charge state throughout the PS plume; 3) laser-induced fluorescence was used to measure the ion velocity and temperature in the near-field plume; 4) a 35 GHz microwave interferometer was developed to measure plasma oscillations and electron density in the plume; and 5) the near-field and internal

Current Driven Instabilities and Anomalous Mobility in Hall-effect Thrusters

NASA Astrophysics Data System (ADS)

Tran, Jonathan; Eckhardt, Daniel; Martin, Robert

2017-10-01

Due to the extreme cost of fully resolving the Debye length and plasma frequency, hybrid plasma simulations utilizing kinetic ions and quasi-steady state fluid electrons have long been the principle workhorse methodology for Hall-effect thruster (HET) modeling. Plasma turbulence and the resulting anomalous electron transport in HETs is a promising candidate for developing predictive models for the observed anomalous transport. In this work, we investigate the implementation of an anomalous electron cross field transport model for hybrid HET simulations such a HPHall. A theory for anomalous transport in HETs and current driven instabilities has been recently studied by Lafleur et al. This work has shown collective electron-wave scattering due to large amplitude azimuthal fluctuations of the electric field. We will further adapt the previous results for related current driven instabilities to electric propulsion relevant mass ratios and conduct a preliminary study of resolving this instability with a modified hybrid (fluid electron and kinetic ion) simulation with the hope of integration with established hybrid HET simulations. This work is supported by the Air Force Office of Scientific Research award FA9950-17RQCOR465.

Numerical study of low-frequency discharge oscillations in a 5 kW Hall thruster

NASA Astrophysics Data System (ADS)

Le, YANG; Tianping, ZHANG; Juanjuan, CHEN; Yanhui, JIA

2018-07-01

A two-dimensional particle-in-cell plasma model is built in the R–Z plane to investigate the low-frequency plasma oscillations in the discharge channel of a 5 kW LHT-140 Hall thruster. In addition to the elastic, excitation, and ionization collisions between neutral atoms and electrons, the Coulomb collisions between electrons and electrons and between electrons and ions are analyzed. The sheath characteristic distortion is also corrected. Simulation results indicate the capability of the built model to reproduce the low-frequency oscillation with high accuracy. The oscillations of the discharge current and ion density produced by the model are consistent with the existing conclusions. The model predicts a frequency that is consistent with that calculated by the zero-dimensional theoretical model.

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

NASA Technical Reports Server (NTRS)

Lopez Ortega, Alejandro; Jorns, Benjamin A.; Mikellides, Ioannis G.; Hofer, Richard R.

2015-01-01

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.

Overview of NASA Iodine Hall Thruster Propulsion System Development

NASA Technical Reports Server (NTRS)

Smith, Timothy D.; Kamhawi, Hani; Hickman, Tyler; Haag, Thomas; Dankanich, John; Polzin, Kurt; Byrne, Lawrence; Szabo, James

2016-01-01

NASA is continuing to invest in advancing Hall thruster technologies for implementation in commercial and government missions. The most recent focus has been on increasing the power level for large-scale exploration applications. However, there has also been a similar push to examine applications of electric propulsion for small spacecraft in the range of 300 kg or less. There have been several recent iodine Hall propulsion system development activities performed by the team of the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and Busek Co. Inc. In particular, the work focused on qualification of the Busek 200-W BHT-200-I and development of the 600-W BHT-600-I systems. This paper discusses the current status of iodine Hall propulsion system developments along with supporting technology development efforts.

Helical plasma thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Beklemishev, A. D., E-mail: bekl@bk.ru

2015-10-15

A new scheme of plasma thruster is proposed. It is based on axial acceleration of rotating magnetized plasmas in magnetic field with helical corrugation. The idea is that the propellant ionization zone can be placed into the local magnetic well, so that initially the ions are trapped. The E × B rotation is provided by an applied radial electric field that makes the setup similar to a magnetron discharge. Then, from the rotating plasma viewpoint, the magnetic wells of the helically corrugated field look like axially moving mirror traps. Specific shaping of the corrugation can allow continuous acceleration of trapped plasma ionsmore » along the magnetic field by diamagnetic forces. The accelerated propellant is expelled through the expanding field of magnetic nozzle. By features of the acceleration principle, the helical plasma thruster may operate at high energy densities but requires a rather high axial magnetic field, which places it in the same class as the VASIMR{sup ®} rocket engine.« less

Magnetic Shielding of the Channel Walls in a Hall Plasma Accelerator

NASA Technical Reports Server (NTRS)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.; Goebel, Dan M.; deGrys, Kristi; Mathers, Alex

2011-01-01

In a qualification life test of a Hall thruster it was found that the erosion of the acceleration channel practically stopped after approx 5,600 h. Numerical simulations using a two-dimensional axisymmetric plasma solver with a magnetic field-aligned mesh reveal that when the channel receded from its early-in-life to its steady-state configuration the following changes occurred near the wall: (1) reduction of the electric field parallel to the wall that prohibited ions from acquiring significant impact kinetic energy before entering the sheath, (2) reduction of the potential fall in the sheath that further diminished the total energy ions gained before striking the material, and (3) reduction of the ion number density that decreased the flux of ions to the wall. All these changes, found to have been induced by the magnetic field, constituted collectively an effective shielding of the walls from any significant ion bombardment. Thus, we term this process in Hall thrusters "magnetic shielding."

Langmuir Probe Measurements Within the Discharge Channel of the 20-kW NASA-300M and NASA-300MS Hall Thrusters

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Huang, Wensheng; Haag, Thomas W.; Kamhawi, Hani

2013-01-01

NASA is presently developing a high-power, high-efficiency, long-lifetime Hall thruster for the Solar Electric Propulsion Technology Demonstration Mission. In support of this task, studies have been performed on the 20-kW NASA-300M Hall thruster to aid in the overall design process. The ability to incorporate magnetic shielding into a high-power Hall thruster was also investigated with the NASA- 300MS, a modified version of the NASA-300M. The inclusion of magnetic shielding would allow the thruster to push existing state-of-the-art technology in regards to service lifetime, one of the goals of the Technology Demonstration Mission. Langmuir probe measurements were taken within the discharge channels of both thrusters in order to characterize differences at higher power levels, as well as validate ongoing modeling efforts using the axisymmetric code Hall2De. Flush-mounted Langmuir probes were also used within the channel of the NASA-300MS to verify that magnetic shielding was successfully applied. Measurements taken from 300 V, 10 kW to 600 V, 20 kW have shown plasma potentials near anode potential and electron temperatures of 4 to 12 eV at the walls near the thruster exit plane of the NASA-300MS, verifying magnetic shielding and validating the design process at this power level. Channel centerline measurements on the NASA-300M from 300 V, 10 kW to 500 V, 20 kW show the electron temperature peak at approximately 0.1 to 0.2 channel lengths upstream of the exit plane, with magnitudes increasing with discharge voltage. The acceleration profiles appear to be centered about the exit plane with a width of approximately 0.3 to 0.4 channel lengths. Channel centerline measurements on the NASA-300MS were found to be more challenging due to additional probe heating. Ionization and acceleration zones appeared to move downstream on the NASA-300MS compared to the NASA-300M, as expected based on the shift in peak radial magnetic field. Additional measurements or alternative

Electron energy balance and ionization in the channel of a stationary plasma thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Veselovzorov, A. N., E-mail: Veselovzorov-AN@nrcki.ru; Pogorelov, A. A.; Svirskiy, E. B.

2016-03-15

The paper presents results of numerical simulations of the electron dynamics in the field of the azimuthal and longitudinal waves excited in the channel of a stationary plasma thruster (SPT). The simulations are based on the experimentally determined wave characteristics. The simulation results show that the azimuthal wave displayed as ionization instability enhances electron transport along the thruster channel. It is established that the electron transport rate in the azimuthal wave increases as compared to the rate of diffusion caused by electron scattering from neutral atoms in proportion to the ratio between the times of electron− neutral collisions responsible formore » ionization and elastic electron scattering, respectively. An expression governing the plasma conductivity is derived with allowance for electron interaction with the azimuthal wave. The Hall parameter, the electron component of the discharge current, and the electron heating power in the thruster channel are calculated for two model SPTs operating with krypton and xenon. The simulation results agree well with the results of experimental studies of these two SPTs.« less

Facility Effect Characterization Test of NASA's HERMeS Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Haag, Thomas W.; Ortega, Alejandro Lopez; Mikellides, Ioannis G.

2016-01-01

A test to characterize the effect of varying background pressure on NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding had being completed. This thruster is the baseline propulsion system for the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). Potential differences in thruster performance and oscillation characteristics when in ground facilities versus on-orbit are considered a primary risk for the propulsion system of the Asteroid Redirect Robotic Mission, which is a candidate for SEP TDM. The first primary objective of this test was to demonstrate that the tools being developed to predict the zero-background-pressure behavior of the thruster can provide self-consistent results. The second primary objective of this test was to provide data for refining a physics-based model of the thruster plume that will be used in spacecraft interaction studies. Diagnostics deployed included a thrust stand, Faraday probe, Langmuir probe, retarding potential analyzer, Wien filter spectrometer, and high-speed camera. From the data, a physics-based plume model was refined. Comparisons of empirical data to modeling results are shown.

Recent Results From Internal and Very-Near-Field Plasma Diagnostics of a High Specific Impulse Hall Thruster

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Gallimore, Alec D.; Jacobson, David (Technical Monitor)

2003-01-01

Floating potential and ion current density measurements were taken on the laboratory model NASA-173Mv2 in order to improve understanding of the physical processes affecting Hall thruster performance at high specific impulse. Floating potential was measured on discharge chamber centerline over axial positions spanning 10 mm from the anode to 100 mm downstream of the exit plane. Ion current density was mapped radially up to 300 mm from thruster centerline over axial positions in the very-near-field (10 to 250 mm from the exit plane). All data were collected using a planar probe in conjunction with a high-speed translation stage to minimize probe-induced thruster perturbations. Measurements of floating potential at a xenon flow rate of 10 mg/s have shown that the acceleration layer moved upstream 3 1 mm when the voltage increased from 300 to 600 V. The length of the acceleration layer was 14 2 mm and was approximately constant with voltage and magnetic field. Ion current density measurements indicated the annular ion beam crossed the thruster centerline 163 mm downstream of the exit plane. Radial integration of the ion current density at the cathode plane provided an estimate of the ion current fraction. At 500 V and 5 mg/s, the ion current fraction was calculated as 0.77.

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Shastry, Rohit; Thomas, Robert; Yim, John; Herman, Daniel; Williams, George; Myers, James; Hofer, Richard;

Coaxial plasma thrusters for high specific impulse propulsion

NASA Technical Reports Server (NTRS)

Schoenberg, Kurt F.; Gerwin, Richard A.; Barnes, Cris W.; Henins, Ivars; Mayo, Robert; Moses, Ronald, Jr.; Scarberry, Richard; Wurden, Glen

1991-01-01

A fundamental basis for coaxial plasma thruster performance is presented and the steady-state, ideal MHD properties of a coaxial thruster using an annular magnetic nozzle are discussed. Formulas for power usage, thrust, mass flow rate, and specific impulse are acquired and employed to assess thruster performance. The performance estimates are compared with the observed properties of an unoptimized coaxial plasma gun. These comparisons support the hypothesis that ideal MHD has an important role in coaxial plasma thruster dynamics.

Visual evidence of suppressing the ion and electron energy loss on the wall in Hall thrusters

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Peng, Wuji; Sun, Hezhi; Wei, Liqiu; Zeng, Ming; Wang, Fufeng; Yu, Daren

2017-03-01

A method of pushing down magnetic field with two permanent magnetic rings is proposed in this paper. It can realize ionization in a channel and acceleration outside the channel. The wall will only suffer from the bombardment of low-energy ions and electrons, which can effectively reduce channel erosion and extend the operational lifetime of thrusters. Furthermore, there is no additional power consumption of coils, which improves the efficiency of systems. We show here the newly developed 200 W no wall-loss Hall thruster (NWLHT-200) that applies the method of pushing down magnetic field with two permanent magnetic rings; the visual evidence we obtained preliminarily confirms the feasibility that the proposed method can realize discharge without wall energy loss or erosion of Hall thrusters.

A zero-equation turbulence model for two-dimensional hybrid Hall thruster simulations

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cappelli, Mark A., E-mail: cap@stanford.edu; Young, Christopher V.; Cha, Eunsun

2015-11-15

We present a model for electron transport across the magnetic field of a Hall thruster and integrate this model into 2-D hybrid particle-in-cell simulations. The model is based on a simple scaling of the turbulent electron energy dissipation rate and the assumption that this dissipation results in Ohmic heating. Implementing the model into 2-D hybrid simulations is straightforward and leverages the existing framework for solving the electron fluid equations. The model recovers the axial variation in the mobility seen in experiments, predicting the generation of a transport barrier which anchors the region of plasma acceleration. The predicted xenon neutral andmore » ion velocities are found to be in good agreement with laser-induced fluorescence measurements.« less

Computed versus measured ion velocity distribution functions in a Hall effect thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Garrigues, L.; CNRS, LAPLACE, F-31062 Toulouse; Mazouffre, S.

2012-06-01

We compare time-averaged and time-varying measured and computed ion velocity distribution functions in a Hall effect thruster for typical operating conditions. The ion properties are measured by means of laser induced fluorescence spectroscopy. Simulations of the plasma properties are performed with a two-dimensional hybrid model. In the electron fluid description of the hybrid model, the anomalous transport responsible for the electron diffusion across the magnetic field barrier is deduced from the experimental profile of the time-averaged electric field. The use of a steady state anomalous mobility profile allows the hybrid model to capture some properties like the time-averaged ion meanmore » velocity. Yet, the model fails at reproducing the time evolution of the ion velocity. This fact reveals a complex underlying physics that necessitates to account for the electron dynamics over a short time-scale. This study also shows the necessity for electron temperature measurements. Moreover, the strength of the self-magnetic field due to the rotating Hall current is found negligible.« less

Plasma properties in electron-bombardment ion thrusters

NASA Technical Reports Server (NTRS)

Matossian, J. N.; Beattie, J. R.

1987-01-01

The paper describes a technique for computing volume-averaged plasma properties within electron-bombardment ion thrusters, using spatially varying Langmuir-probe measurements. Average values of the electron densities are defined by integrating the spatially varying Maxwellian and primary electron densities over the ionization volume, and then dividing by the volume. Plasma properties obtained in the 30-cm-diameter J-series and ring-cusp thrusters are analyzed by the volume-averaging technique. The superior performance exhibited by the ring-cusp thruster is correlated with a higher average Maxwellian electron temperature. The ring-cusp thruster maintains the same fraction of primary electrons as does the J-series thruster, but at a much lower ion production cost. The volume-averaged predictions for both thrusters are compared with those of a detailed thruster performance model.

The X3: A 200 kW Class Nested Channel Hall Thruster

NASA Astrophysics Data System (ADS)

Sheehan, J. P.

2016-10-01

Electric propulsion has seen rapid adoption in recent years for commercial, scientific, and exploratory space missions. The X3 is a three channel nested channel Hall thruster, designed to push the boundaries of high power electric propulsion for cargo transfer to Mars and large military assets. It has been operated at thermal steady state up to 30 kW of power. Thrust measurements were made on an inverted pendulum thrust stand, indicating over 2000 s specific impulse and 65 mN/kW thrust to power ratio. Detailed plume measurements were made with Faraday and Langmuir probes. The multiple concentric channels provide better performance than the sum of the individual channel operations due to superior propellant utilization from its compact design. Using a high speed camera, the breathing and spoke mode instabilities were captured in all three channels. Spoke and breathing instabilities couple between the channels, indicating that complex plasma and neutral interactions are at play. Electron transport, both cross field and in the cathode plume, are well suited to be explored in a thruster of this size. Supported under NASA contract No. NNH16CP17C.

Time-resolved ion velocity distribution in a cylindrical Hall thruster: heterodyne-based experiment and modeling.

PubMed

Diallo, A; Keller, S; Shi, Y; Raitses, Y; Mazouffre, S

2015-03-01

Time-resolved variations of the ion velocity distribution function (IVDF) are measured in the cylindrical Hall thruster using a novel heterodyne method based on the laser-induced fluorescence technique. This method consists in inducing modulations of the discharge plasma at frequencies that enable the coupling to the breathing mode. Using a harmonic decomposition of the IVDF, one can extract each harmonic component of the IVDF from which the time-resolved IVDF is reconstructed. In addition, simulations have been performed assuming a sloshing of the IVDF during the modulation that show agreement between the simulated and measured first order perturbation of the IVDF.

Experimental Analysis of Dampened Breathing Mode Oscillation on Hall Thruster Performance

DTIC Science & Technology

2013-03-01

38 4.5 Analysis of Discharge RMS Effect on Breathing Mode Amplitude...20 xii EXPERIMENTAL ANALYSIS OF DAMPENED BREATHING MODE OSCILLATION ON HALL EFFECT THRUSTER...the large error in the data presented above prevents many conclusions from being drawn. 4.5 Analysis of Discharge RMS Effect on Breathing Mode

Pseudospectral Model for Hybrid PIC Hall-effect Thruster Simulation

DTIC Science & Technology

2015-07-01

and Fernandez6 (hybrid- PIC ). This work follows the example of Lam and Fernandez but substitutes a spectral description in the azimuthal direction to...Paper 3. DATES COVERED (From - To) July 2015-July 2015 4. TITLE AND SUBTITLE Pseudospectral model for hybrid PIC Hall-effect thruster simulationect...of a pseudospectral azimuthal-axial hybrid- PIC HET code which is designed to explicitly resolve and filter azimuthal fluctuations in the

Anode power deposition in a MPD thruster with a magnetically annulled Hall parameter anode

NASA Technical Reports Server (NTRS)

Gallimore, Alec D.; Kelly, Arnold J.; Jahn, Robert G.

1992-01-01

Results from previous studies indicate that the anode fall increases monotonically with the electron Hall parameter. In an attempt to reduce the anode fall by decreasing the local electron Hall parameter, a proof-of-concept test was performed in which an array of 36 permanent magnets were imbedded within the anode of a high power quasi-steady MPD thruster to decrease the local azimuthal component of the induced magnetic field. The modified thruster was operated at power levels between 150 kW and 4 MW with Ar and He propellants. Terminal voltage, triple probe, floating probe, and magnetic probe measurements were made to characterize the performance of the thruster with new anode. Incorporation of the modified anode resulted in a reduction of the anode fall by up to 15 V with Ar and 20 V with He, which corresponded to decreased anode power fractions of 40 and 45 percent with Ar and He, respectively.

Mission Benefits of Gridded Ion and Hall Thruster Hybrid Propulsion Systems

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Polsgrove, Tara

2006-01-01

The NASA In-Space Propulsion Technology (ISPT) Project Office has been developing the NEXT gridded ion thruster system and is planning to procure a low power Hall system. The new ion propulsion systems will join NSTAR as NASA's primary electric propulsion system options. Studies have been performed to show mission benefits of each of the stand alone systems. A hybrid ion propulsion system (IPS) can have the advantage of reduced cost, decreased flight time and greater science payload delivery over comparable homogeneous systems. This paper explores possible advantages of combining various thruster options for a single mission.

Computer simulations of Hall thrusters without wall losses designed using two permanent magnetic rings

NASA Astrophysics Data System (ADS)

Yongjie, Ding; Wuji, Peng; Liqiu, Wei; Guoshun, Sun; Hong, Li; Daren, Yu

2016-11-01

A type of Hall thruster without wall losses is designed by adding two permanent magnet rings in the magnetic circuit. The maximum strength of the magnetic field is set outside the channel. Discharge without wall losses is achieved by pushing down the magnetic field and adjusting the channel accordingly. The feasibility of the Hall thrusters without wall losses is verified via a numerical simulation. The simulation results show that the ionization region is located in the discharge channel and the acceleration region is outside the channel, which decreases the energy and flux of ions and electrons spattering on the wall. The power deposition on the channel walls can be reduced by approximately 30 times.

Charge-exchange plasma generated by an ion thruster

NASA Technical Reports Server (NTRS)

Kaufman, H. R.

1977-01-01

The charge exchange plasma generated by an ion thruster was investigated experimentally using both 5 cm and 15 cm thrusters. Results are shown for wide ranges of radial distance from the thruster and angle from the beam direction. Considerations of test environment, as well as distance from the thruster, indicate that a valid simulation of a thruster on a spacecraft was obtained. A calculation procedure and a sample calculation of charge exchange plasma density and saturation electron current density are included.

Stationary Plasma Thruster Plume Emissions

NASA Technical Reports Server (NTRS)

Manzella, David H.

1994-01-01

The emission spectrum from a xenon plasma produced by a Stationary Plasma Thruster provided by the Ballistic Missile Defense Organization (BMDO) was measured. Approximately 270 individual Xe I, Xe II, and XE III transitions were identified. A total of 250 mW of radiated optical emission was estimated from measurements taken at the thruster exit plane. There was no evidence of erosion products in the emission signature. Ingestion and ionization of background gas at elevated background pressure was detected. The distribution of excited states could be described by temperatures ranging from fractions of 1 eV to 4 eV with a high degree of uncertainty due to the nonequilibrium nature of this plasma. The plasma was over 95 percent ionized at the thruster exit plane. Between 10 and 20 percent of the ions were doubly charged. Two modes of operation were identified. The intensity of plasma emission increased by a factor of two during operation in an oscillatory mode. The transfer between the two modes of operation was likely related to unidentified phenomena occurring on a time scale of minutes.

Experimental Investigation of a Hall-Current Accelerator. M.S. Thesis

NASA Technical Reports Server (NTRS)

Plank, G. M.

1983-01-01

The Hall-current accelerator is being investigated for use in the 1000-2000 sec. range of specific impulse. Three models of this thruster were tested. The first two models had three permanent magnets to supply the magnetic field and the third model had six magnets to supply the field. The third model thus had approximately twice the magnetic field of the first two. The first and second models differ only in the shape of the magnetic field. All other factors remained the same for the three models except for the anode-cathode distance, which was changed to allow for the three thrusters to have the same magnetic field integral between the anode and the cathode. These Hall thrusters were tested to determine the plasma properties, the beam characteristics, and the thruster characteristics. The thruster operated in three modes: (1) main cathode only, (2) main cathode with neutralizer cathode, and (3) neutralizer cathode only. The plasma properties were measured along an axial line, 1 mm inside the cathode radius, at a distance of 0.2 to 6.2 cm from the anode. Results show that the current used to heat the cathode produced nonuniformities in the magnetic field, hence also in the plasma properties. In a Hall thruster this general design appears to provide the most thrust when operated at a magnetic field less than the maximum value studied.

Analysis of Wien filter spectra from Hall thruster plumes.

PubMed

Huang, Wensheng; Shastry, Rohit

2015-07-01

A method for analyzing the Wien filter spectra obtained from the plumes of Hall thrusters is derived and presented. The new method extends upon prior work by deriving the integration equations for the current and species fractions. Wien filter spectra from the plume of the NASA-300M Hall thruster are analyzed with the presented method and the results are used to examine key trends. The new integration method is found to produce results slightly different from the traditional area-under-the-curve method. The use of different velocity distribution forms when performing curve-fits to the peaks in the spectra is compared. Additional comparison is made with the scenario where the current fractions are assumed to be proportional to the heights of peaks. The comparison suggests that the calculated current fractions are not sensitive to the choice of form as long as both the height and width of the peaks are accounted for. Conversely, forms that only account for the height of the peaks produce inaccurate results. Also presented are the equations for estimating the uncertainty associated with applying curve fits and charge-exchange corrections. These uncertainty equations can be used to plan the geometry of the experimental setup.

The Importance of the Cathode Plume and Its Interactions with the Ion Beam in Numerical Simulations of Hall Thrusters

NASA Technical Reports Server (NTRS)

Lopez Ortega, Alejandro; Mikellides, Ioannis G.

2015-01-01

Hall2De is a first-principles, 2-D axisymmetric code that solves the equations of motion for ions, electrons, and neutrals on a magnetic-field-aligned grid. The computational domain downstream of the acceleration channel exit plane is large enough to include self-consistently the cathode boundary. In this paper, we present results from numerical simulations of the H6 laboratory thruster with an internally mounted cathode, with the aim of highlighting the importance of properly accounting for the interactions between the ion beam and cathode plume. The anomalous transport of electrons across magnetic field lines in Hall2De is modelled using an anomalous collision frequency, ?anom, yielding ?anom approximately equal to omega ce (i.e., the electron cyclotron frequency) in the plume. We first show that restricting the anomalous collision frequency to only regions where the current density of ions is large does not alter the plasma discharge in the Hall thruster as long as the interaction between the ion beam and the cathode plume is captured properly in the computational domain. This implies that the boundary conditions must be placed sufficiently far as to not interfere with the electron transport in this region. These simulation results suggest that electron transport across magnetic field lines occurs largely inside the beam and may be driven by the interactions between beam ions and electrons. A second finding that puts in relevance the importance of including the cathode plume in numerical simulations is on the significance of accounting for the ion acoustic turbulence (IAT), now known to occur in the vicinity of the cathode exit. We have included in the Hall2De simulations a model of the IAT-driven anomalous collision frequency based on Sagdeev's model for saturation of the ion-acoustic instability. This implementation has allowed us to achieve excellent agreement with experimental measurements in the near plume obtained during the operation of the H6 thruster at

Effect of oblique channel on discharge characteristics of 200-W Hall thruster

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Peng, Wuji; Sun, Hezhi; Xu, Yu; Wei, Liqiu; Li, Hong; Zeng, Ming; Wang, Fufeng; Yu, Daren

2017-02-01

In an experiment involving a 200-W Hall thruster, partial ionization occurs in the plume area because of the extrapolation of the magnetic field. To improve the thruster performance, the concept of an oblique channel is proposed for improving the ionization degree in the plume area. Calculations performed using a Particle-in-cell (PIC) simulator and the experimental results both show that an oblique channel structure can reduce the wall loss. Compared with a straight channel under similar conditions of the discharge voltage and current, the ionization degree in the plume area, thrust, specific impulse, propellant utilization, and anode efficiency are improved by ˜20%. The oblique channel is an important design consideration for improving the partial ionization of the plume area in the thruster.

Carbon Back Sputter Modeling for Hall Thruster Testing

NASA Technical Reports Server (NTRS)

Gilland, James H.; Williams, George J.; Burt, Jonathan M.; Yim, John T.

2016-01-01

In support of wear testing for the Hall Effect Rocket with Magnetic Shielding (HERMeS) program, the back sputter from a Hall effect thruster plume has been modeled for the NASA Glenn Research Centers Vacuum Facility 5. The predicted wear at a near-worst case condition of 600 V, 12.5 kW was found to be on the order of 3 4 mkhour in a fully carbon-lined chamber. A more detailed numerical monte carlo code was also modified to estimate back sputter for a detailed facility and pumping configuration. This code demonstrated similar back sputter rate distributions, but is not yet accurately modeling the magnitudes. The modeling has been benchmarked to recent HERMeS wear testing, using multiple microbalance measurements. These recent measurements have yielded values, on the order of 1.5- 2 microns/khour.

Performance, Facility Pressure Effects, and Stability Characterization Tests of NASA's Hall Effect Rocket with Magnetic Shielding Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Peterson, Peter; Hofer, Richard; Mikellides, Ioannis

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for flight system development. Part of the technology maturation effort included experimental evaluation of the TDU-1 thruster with conducting and dielectric front pole cover materials in two different electrical configurations. A graphite front pole cover thruster configuration with the thruster body electrically tied to cathode and an alumina front pole cover thruster configuration with the thruster body floating were evaluated. Both configurations were also evaluated at different facility background pressure conditions to evaluate background pressure effects on thruster operation. Performance characterization tests found that higher thruster performance was attained with the graphite front pole cover configuration with the thruster electrically tied to cathode. A total thrust efficiency of 68 and a total specific impulse of 2,820 s was demonstrated at a discharge voltage of 600 V and a discharge power of 12.5 kW. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations and with maps of the current-voltage-magnetic field (IVB). Analysis of TDU-1 discharge current waveforms found that lower normalized discharge current peak-to-peak and root mean square magnitudes were attained when the thruster was electrically floated with alumina front pole covers. Background pressure effects characterization tests indicated that the thruster performance and stability was mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 110-5 Torr-Xe (for thruster flow rate above 8 mgs). Power spectral density analysis of the discharge current waveform showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant frequency. Finally the IVB maps of the TDU-1

Where is the breathing mode? High voltage Hall effect thruster studies with EMD method

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kurzyna, J.; Makowski, K.; Mazouffre, S.

2008-03-19

Discharge current and local plasma oscillations are studied in a high voltage Hall effect thruster PPS registered -X000. Characteristic time scales that appear in different operating conditions are resolved with the use of Hilbert-Huang spectra (HHS) which display time dependenc of instantaneous frequency and power. Sets of intrinsic mode functions (imfs) that are used for HHS calculation result due to application of empirical mode decomposition method (EMD) to nonstationary multicomponent signals. In the experiment the signals are captured in the electric circuit of the thruster as well locally, in the vicinity of the thruster exhaust region. Classical electric probes spacedmore » along the azimuth and/or thruster axis let us study correlations of signals which were captured in different locations. In this way azimuthal and axial propagation of disturbances is inspected. The discharge voltage is varied in the range of 200 divide 900 V while xenon mas flow rate of 5 divide 9 mg/s. LF, MF, and HF characteristic bands that are known from previous studies of PPS registered -100 thruster have been also detected here. However, expanding discharge current onto a set of intrinsic modes we can resolve MF mode more reliably than before. Moreover, for higher discharge voltages, this irregular mode turns into more regular waveform and tends to dominate in the discharge current masking almost completely the breathing mode (LF oscillations of the discharge current). In such a case triggering of HF oscillations is correlated with the phase of MF mode while in the case of PPS registered -100 thruster it was correlated with the appropriate phase of the breathing mode (LF band). Regular HF emission that can be unambiguously interpreted as azimuthal electrostatic wave now is observed only in the specific operating conditions of the thruster. However, even if irregular HF emission is observed the time delay of cross-correlated signals which are captured in different azimuthal

Laser characterization of the unsteady 2-D ion flow field in a Hall thruster with breathing mode oscillations

NASA Astrophysics Data System (ADS)

Lucca Fabris, Andrea; Young, Christopher; MacDonald-Tenenbaum, Natalia; Hargus, William, Jr.; Cappelli, Mark

2016-10-01

Hall thrusters are a mature form of electric propulsion for spacecraft. One commonly observed low frequency (10-50 kHz) discharge current oscillation in these E × B devices is the breathing mode, linked to a propagating ionization front traversing the channel. The complex time histories of ion production and acceleration in the discharge channel and near-field plume lead to interesting dynamics and interactions in the central plasma jet and downstream plume regions. A time-resolved laser-induced fluorescence (LIF) diagnostic non-intrusively measures 2-D ion velocity and relative ion density throughout the plume of a commercial BHT-600 Hall thruster manufactured by Busek Co. Low velocity classes of ions observed in addition to the main accelerated population are linked to propellant ionization outside of the device. Effects of breathing mode dynamics are shown to persist far downstream where modulations in ion velocity and LIF intensity are correlated with discharge current oscillations. This work is sponsored by the U.S. Air Force Office of Scientific Research with Dr. M. Birkan as program manager. C.Y. acknowledges support from the DOE NSSA Stewardship Science Graduate Fellowship under contract DE-FC52-08NA28752.

Preliminary Evaluation of a 10 kW Hall Thruster

NASA Technical Reports Server (NTRS)

Jankovsky, Robert S.; McLean, Chris; McVey, John

1999-01-01

A 10 kW Hall thruster was characterized over a range of discharge voltages from 300-500 V and a range of discharge currents from 15-23 A. This corresponds to power levels from a low of 4.6 kW to a high of 10.7 kW. Over this range of discharge powers, thrust varied from 278 mN to 524 mN, specific impulse ranged from 1644 to 2392 seconds, and efficiency peaked at approximately 59%. A continuous 40 hour test was also undertaken in an attempt to gain insight with regard to long term operation of the engine. For this portion of the testing the thruster was operated at a discharge voltage of 500 V and a discharge current of 20 A. Steady-state temperatures were achieved after 3-5 hrs and very little variation in performance was detected.

Development of a two-dimensional dual pendulum thrust stand for Hall thrusters.

PubMed

Nagao, N; Yokota, S; Komurasaki, K; Arakawa, Y

2007-11-01

A two-dimensional dual pendulum thrust stand was developed to measure thrust vectors [axial and horizontal (transverse) direction thrusts] of a Hall thruster. A thruster with a steering mechanism is mounted on the inner pendulum, and thrust is measured from the displacement between inner and outer pendulums, by which a thermal drift effect is canceled out. Two crossover knife-edges support each pendulum arm: one is set on the other at a right angle. They enable the pendulums to swing in two directions. Thrust calibration using a pulley and weight system showed that the measurement errors were less than 0.25 mN (1.4%) in the main thrust direction and 0.09 mN (1.4%) in its transverse direction. The thrust angle of the thrust vector was measured with the stand using the thruster. Consequently, a vector deviation from the main thrust direction of +/-2.3 degrees was measured with the error of +/-0.2 degrees under the typical operating conditions for the thruster.

Electron energy distribution function in a low-power Hall thruster discharge and near-field plume

NASA Astrophysics Data System (ADS)

Tichý, M.; Pétin, A.; Kudrna, P.; Horký, M.; Mazouffre, S.

2018-06-01

Electron temperature and plasma density, as well as the electron energy distribution function (EEDF), have been obtained inside and outside the dielectric channel of a 200 W permanent magnet Hall thruster. Measurements were carried out by means of a cylindrical Langmuir probe mounted onto a compact fast moving translation stage. The 3D particle-in cell numerical simulations complement experiments. The model accounts for the crossed electric and magnetic field configuration in a weakly collisional regime where only electrons are magnetized. Since only the electron dynamics is of interest in this study, an artificial mass of ions corresponding to mi = 30 000me was used to ensure ions could be assumed at rest. The simulation domain is located at the thruster exit plane and does not include the cathode. The measured EEDF evidences a high-energy electron population that is superimposed onto the low energy bulk population outside the channel. Inside the channel, the EEDF is close to Maxwellian. Both the experimental and numerical EEDF depart from an equilibrium distribution at the channel exit plane, a region of high magnetic field. We therefore conclude that the fast electron group found in the experiment corresponds to the electrons emitted by the external cathode that reach the thruster discharge without experiencing collision events.

Magnetic field deformation due to electron drift in a Hall thruster

NASA Astrophysics Data System (ADS)

Liang, Han; Yongjie, Ding; Xu, Zhang; Liqiu, Wei; Daren, Yu

2017-01-01

The strength and shape of the magnetic field are the core factors in the design of the Hall thruster. However, Hall current can affect the distribution of static magnetic field. In this paper, the Particle-In-Cell (PIC) method is used to obtain the distribution of Hall current in the discharge channel. The Hall current is separated into a direct and an alternating part to calculate the induced magnetic field using Finite Element Method Magnetics (FEMM). The results show that the direct Hall current decreases the magnetic field strength in the acceleration region and also changes the shape of the magnetic field. The maximum reduction in radial magnetic field strength in the exit plane is 10.8 G for an anode flow rate of 15 mg/s and the maximum angle change of the magnetic field line is close to 3° in the acceleration region. The alternating Hall current induces an oscillating magnetic field in the whole discharge channel. The actual magnetic deformation is shown to contain these two parts.

Acceleration and focusing of plasma flows

DOE Office of Scientific and Technical Information (OSTI.GOV)

Griswold, Martin Elias

The acceleration of flowing plasmas is a fundamental problem that is useful in a wide variety of technological applications. We consider the problem from the perspective of plasma propulsion. Gridded ion thrusters and Hall thrusters are the most commonly used devices to create flowing plasma for space propulsion, but both suffer from fundamental limitations. Gridded ion sources create good quality beams in terms of energy spread and spatial divergence, but the Child-Langmuir law in the non-neutral acceleration region limits the maximum achievable current density. Hall thrusters avoid this limitation by accelerating ions in quasi-neutral plasma but, as a result, producemore » plumes with high spatial divergence and large energy spread. In addition the more complicated magnetized plasma in the Hall Thruster produces oscillations that can reduce the efficiency of the thruster by increasing electron transport to the anode. We present investigations of three techniques to address the fundamental limitations on the performance of each thruster. First, we propose a method to increase the time-averaged current density (and thus thrust density) produced by a gridded ion source above the Child-Langmuir limit by introducing time-varying boundary conditions. Next, we use an electrostatic plasma lens to focus the Hall thruster plume, and finally we develop a technique to suppress a prominent oscillation that degrades the performance of Hall thrusters. The technique to loosen the constraints on current density from gridded ion thrusters actually applies much more broadly to any space charge limited flow. We investigate the technique with a numerical simulation and by proving a theoretical upper bound. While we ultimately conclude that the approach is not suitable for space propulsion, our results proved useful in another area, providing a benchmark for research into the spontaneously time-dependent current that arises in microdiodes. Next, we experimentally demonstrate a novel

The Effects of Magnetic Nozzle Configurations on Plasma Thrusters

NASA Technical Reports Server (NTRS)

Turchi, P. J.

1997-01-01

Over the course of eight years, the Ohio State University has performed research in support of electric propulsion development efforts at the NASA Lewis Research Center, Cleveland, OH. This research has been largely devoted to plasma propulsion systems including MagnetoPlasmaDynamic (MPD) thrusters with externally-applied, solenoidal magnetic fields, hollow cathodes, and Pulsed Plasma Microthrusters (PPT's). Both experimental and theoretical work has been performed, as documented in four master's theses, two doctoral dissertations, and numerous technical papers. The present document is the final report for the grant period 5 December 1987 to 31 December 1995, and summarizes all activities. Detailed discussions of each area of activity are provided in appendices: Appendix 1 - Experimental studies of magnetic nozzle effects on plasma thrusters; Appendix 2 - Numerical modeling of applied-field MPD thrusters; Appendix 3 - Theoretical and experimental studies of hollow cathodes; and Appendix 4 -Theoretical, numerical and experimental studies of pulsed plasma thrusters. Especially notable results include the efficacy of using a solenoidal magnetic field downstream of a plasma thruster to collimate the exhaust flow, the development of a new understanding of applied-field MPD thrusters (based on experimentally-validated results from state-of-the art, numerical simulation) leading to predictions of improved performance, an experimentally-validated, first-principles model for orificed, hollow-cathode behavior, and the first time-dependent, two-dimensional calculations of ablation-fed, pulsed plasma thrusters.

Los Alamos NEP research in advanced plasma thrusters

NASA Technical Reports Server (NTRS)

Schoenberg, Kurt; Gerwin, Richard

1991-01-01

Research was initiated in advanced plasma thrusters that capitalizes on lab capabilities in plasma science and technology. The goal of the program was to examine the scaling issues of magnetoplasmadynamic (MPD) thruster performance in support of NASA's MPD thruster development program. The objective was to address multi-megawatt, large scale, quasi-steady state MPD thruster performance. Results to date include a new quasi-steady state operating regime which was obtained at space exploration initiative relevant power levels, that enables direct coaxial gun-MPD comparisons of thruster physics and performance. The radiative losses are neglible. Operation with an applied axial magnetic field shows the same operational stability and exhaust plume uniformity benefits seen in MPD thrusters. Observed gun impedance is in close agreement with the magnetic Bernoulli model predictions. Spatial and temporal measurements of magnetic field, electric field, plasma density, electron temperature, and ion/neutral energy distribution are underway. Model applications to advanced mission logistics are also underway.

Comparison of Hall Thruster Plume Expansion Model with Experimental Data

DTIC Science & Technology

2006-05-23

focus of this study, is a hybrid particle- in-cell ( PIC ) model that tracks particles along an unstructured tetrahedral mesh. * Research Engineer...measurements of the ion current density profile, ion energy distributions, and ion species fraction distributions using a nude Faraday probe, retarding...Vol.37 No.1. 6 Oh, D. and Hastings, D., “Three Dimensional PIC -DSMC Simulations of Hall Thruster Plumes and Analysis for Realistic Spacecraft

Compact High Current Rare-Earth Emitter Hollow Cathode for Hall Effect Thrusters

NASA Technical Reports Server (NTRS)

Goebel, Dan M. (Inventor); Watkins, Ronnie M. (Inventor); Hofer, Richard R. (Inventor)

2012-01-01

An apparatus and method for achieving an efficient central cathode in a Hall effect thruster is disclosed. A hollow insert disposed inside the end of a hollow conductive cathode comprises a rare-earth element and energized to emit electrons from an inner surface. The cathode employs an end opening having an area at least as large as the internal cross sectional area of the rare earth insert to enhance throughput from the cathode end. In addition, the cathode employs a high aspect ratio geometry based on the cathode length to width which mitigates heat transfer from the end. A gas flow through the cathode and insert may be impinged by the emitted electrons to yield a plasma. One or more optional auxiliary gas feeds may also be employed between the cathode and keeper wall and external to the keeper near the outlet.

Development of a two-dimensional dual pendulum thrust stand for Hall thrusters

DOE Office of Scientific and Technical Information (OSTI.GOV)

Nagao, N.; Yokota, S.; Komurasaki, K.

A two-dimensional dual pendulum thrust stand was developed to measure thrust vectors (axial and horizontal (transverse) direction thrusts) of a Hall thruster. A thruster with a steering mechanism is mounted on the inner pendulum, and thrust is measured from the displacement between inner and outer pendulums, by which a thermal drift effect is canceled out. Two crossover knife-edges support each pendulum arm: one is set on the other at a right angle. They enable the pendulums to swing in two directions. Thrust calibration using a pulley and weight system showed that the measurement errors were less than 0.25 mN (1.4%)more » in the main thrust direction and 0.09 mN (1.4%) in its transverse direction. The thrust angle of the thrust vector was measured with the stand using the thruster. Consequently, a vector deviation from the main thrust direction of {+-}2.3 deg. was measured with the error of {+-}0.2 deg. under the typical operating conditions for the thruster.« less

Pulsed Plasma Thruster Contamination

NASA Technical Reports Server (NTRS)

Myers, Roger M.; Arrington, Lynn A.; Pencil, Eric J.; Carter, Justin; Heminger, Jason; Gatsonis, Nicolas

1996-01-01

Pulsed Plasma Thrusters (PPT's) are currently baselined for the Air Force Mightysat II.1 flight in 1999 and are under consideration for a number of other missions for primary propulsion, precision positioning, and attitude control functions. In this work, PPT plumes were characterized to assess their contamination characteristics. Diagnostics included planar and cylindrical Langmuir probes and a large number of collimated quartz contamination sensors. Measurements were made using a LES 8/9 flight PPT at 0.24, 0.39, 0.55, and 1.2 m from the thruster, as well as in the backflow region behind the thruster. Plasma measurements revealed a peak centerline ion density and velocity of approx. 6 x 10(exp 12) cm(exp -3) and 42,000 m/s, respectively. Optical transmittance measurements of the quartz sensors after 2 x 10(exp 5) pulses showed a rapid decrease in plume contamination with increasing angle from the plume axis, with a barely measurable transmittance decrease in the ultraviolet at 90 deg. No change in optical properties was detected for sensors in the backflow region.

A New Method for Analyzing Near-Field Faraday Probe Data in Hall Thrusters

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2013-01-01

This paper presents a new method for analyzing near-field Faraday probe data obtained from Hall thrusters. Traditional methods spawned from far-field Faraday probe analysis rely on assumptions that are not applicable to near-field Faraday probe data. In particular, arbitrary choices for the point of origin and limits of integration have made interpretation of the results difficult. The new method, called iterative pathfinding, uses the evolution of the near-field plume with distance to provide feedback for determining the location of the point of origin. Although still susceptible to the choice of integration limits, this method presents a systematic approach to determining the origin point for calculating the divergence angle. The iterative pathfinding method is applied to near-field Faraday probe data taken in a previous study from the NASA-300M and NASA-457Mv2 Hall thrusters. Since these two thrusters use centrally mounted cathodes the current density associated with the cathode plume is removed before applying iterative pathfinding. A procedure is presented for removing the cathode plume. The results of the analysis are compared to far-field probe analysis results. This paper ends with checks on the validity of the new method and discussions on the implications of the results.

A New Method for Analyzing Near-Field Faraday Probe Data in Hall Thrusters

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2013-01-01

This paper presents a new method for analyzing near-field Faraday probe data obtained from Hall thrusters. Traditional methods spawned from far-field Faraday probe analysis rely on assumptions that are not applicable to near-field Faraday probe data. In particular, arbitrary choices for the point of origin and limits of integration have made interpretation of the results difficult. The new method, called iterative pathfinding, uses the evolution of the near-field plume with distance to provide feedback for determining the location of the point of origin. Although still susceptible to the choice of integration limits, this method presents a systematic approach to determining the origin point for calculating the divergence angle. The iterative pathfinding method is applied to near-field Faraday probe data taken in a previous study from the NASA-300M and NASA-457Mv2 Hall thrusters. Since these two thrusters use centrally mounted cathodes, the current density associated with the cathode plume is removed before applying iterative pathfinding. A procedure is presented for removing the cathode plume. The results of the analysis are compared to far-field probe analysis results. This paper ends with checks on the validity of the new method and discussions on the implications of the results.

High Input Voltage Discharge Supply for High Power Hall Thrusters Using Silicon Carbide Devices

NASA Technical Reports Server (NTRS)

Pinero, Luis R.; Scheidegger, Robert J.; Aulsio, Michael V.; Birchenough, Arthur G.

2014-01-01

A power processing unit for a 15 kW Hall thruster is under development at NASA Glenn Research Center. The unit produces up to 400 VDC with two parallel 7.5 kW discharge modules that operate from a 300 VDC nominal input voltage. Silicon carbide MOSFETs and diodes were used in this design because they were the best choice to handle the high voltage stress while delivering high efficiency and low specific mass. Efficiencies in excess of 97 percent were demonstrated during integration testing with the NASA-300M 20 kW Hall thruster. Electromagnet, cathode keeper, and heater supplies were also developed and will be integrated with the discharge supply into a vacuum-rated brassboard power processing unit with full flight functionality. This design could be evolved into a flight unit for future missions that requires high power electric propulsion.

Development, Demonstration, and Analysis of an Integrated Iodine Hall Thruster Feed System

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Steven R.; Burt, Adam O.; Martin, Adam K.; Martinez, Armando; Seixal, Joao F.; Mauro, Stephanie

2016-01-01

The design of an in-space iodine-vapor-fed Hall effect thruster propellant management system is described. The solid-iodine propellant tank has unique issues associated with the microgravity environment, requiring a solution where the iodine is maintained in intimate thermal contact with the heated tank walls. The flow control valves required alterations from earlier iterations to survive for extended periods of time in the corrosive iodine-vapor environment. Materials have been selected for the entire feed system that can chemically resist the iodine vapor, with the design now featuring Hastelloy or Inconel for almost all the wetted components. An integrated iodine feed system/Hall thruster demonstration unit was fabricated and tested, with all control being handled by an onboard electronics card specifically designed to operate the feed system. Structural analysis shows that the feed system can survive launch loads after the implementation of some minor reinforcement. Flow modeling, while still requiring significant additional validation, is presented to show its potential in capturing the behavior of components in this low-flow, low-pressure system.

a Permanent Magnet Hall Thruster for Orbit Control of Lunar Polar Satellites

NASA Astrophysics Data System (ADS)

Ferreira, Jose Leonardo; Silva Moraes, Bruno; Soares Ferreira, Ivan; Cardozo Mour, Decio; Winter, Othon

Future moon missions devoted to lunar surface remote sensing and to many others scientific exploration topics will require more fine and higher precision orbit control. It is well known that, lunar satellites in polar orbits will suffer a high increase on the eccentricity due to the gravitational perturbation of the Earth. Without proper orbit correction the satellite life time will decrease and end up in a collision with the moon surface. It is pointed out by many authors that this effect is a natural consequence of the Lidov-Kozai resonance. In the present work, we propose a precise method of orbit eccentricity control based on the use of a low thrust Hall plasma thruster. The proposed method is based on an approach intended to keep the orbital eccentricity of the satellite at low values. A previous work on this subject was made using numerical integration considering two systems: the 3-body problem, Moon-Earth-satellite and the 4-body problem, Moon-Earth-Sun-satellite (??). In such simulation it is possible to follow the evolution of the satellite's eccentricity and find empirical expressions for the length of time needed to occur the collision with the moon. In this work, a satellite orbit eccentricity control maneuvering is proposed. It is based on working parameters of a low thrust propulsion permanent magnet Hall plasma thruster (PMHT), which is been developed at University of Brasilia, Brazil. We studied different arcs of active lunar satellite propulsion in order to be able to introduce a correction of the eccentricity at each cycle. The calculations were made considering a set of different thrust values, from 0.1N up to 0.4N which can be obtained by using the PMHT. In each calculation procedure we measured the length of eccentricity correction provided by active propulsion. From these results we obtained empirical expressions of the time needed for the corrections as a function of the initial altitude and as a function of the thrust value. 1. Winter, O. C

Test Results of a 200 W Class Hall Thruster

NASA Technical Reports Server (NTRS)

Jacobson, David; Jankovsky, Robert S.

1999-01-01

The performance of a 200 W class Hall thruster was evaluated. Performance measurements were taken at power levels between 90 W and 250 W. At the nominal 200 W design point, the measured thrust was 11.3 mN. and the specific impulse was 1170 s excluding cathode flow in the calculation. A laboratory model 3 mm diameter hollow cathode was used for all testing. The engine was operated on laboratory power supplies in addition to a breadboard power processing unit fabricated from commercially available DC to DC converters.

Development of a Methodology for Conducting Hall Thruster EMI Tests in Metal Vacuum Chambers of Arbitrary Shape and Size

NASA Technical Reports Server (NTRS)

Gallimore, Alec D.

2000-01-01

While the closed-drift Hall thruster (CDT) offers significant improvement in performance over conventional chemical rockets and other advanced propulsion systems such as the arcjet, its potential impact on spacecraft communication signals must be carefully assessed before widespread use of this device can take place. To this end, many of the potentially unique issues that are associated with these thrusters center on its plume plasma characteristics and the its interaction with electromagnetic waves. Although a great deal of experiments have been made in characterizing the electromagnetic interference (EMI) potential of these thrusters, the interpretation of the resulting data is difficult because most of these measurements have been made in vacuum chambers with metal walls which reflect radio waves emanating from the thruster. This project developed a means of assessing the impact of metal vacuum chambers of arbitrary size or shape on EMI experiments, thereby allowing for test results to be interpreted properly. Chamber calibration techniques were developed and initially tested at RIAME using their vacuum chamber. Calibration experiments were to have been made at Tank 5 of NASA GRC and the 6 m by 9 m vacuum chamber at the University of Michigan to test the new procedure, however the subcontract to RIAME was cancelled by NASA memorandum on Feb. 26. 1999.

Mode Transitions in Magnetically Shielded Hall Effect Thrusters

NASA Technical Reports Server (NTRS)

Sekerak, Michael J.; Longmier, Benjamin W.; Gallimore, Alec D.; Huang, Wensheng; Kamhawi, Hani; Hofer, Richard R.; Jorns, Benjamin A.; Polk, James E.

2014-01-01

A mode transition study is conducted in magnetically shielded thrusters where the magnetic field magnitude is varied to induce mode transitions. Three different oscillatory modes are identified with the 20-kW NASA-300MS-2 and the 6-kW H6MS: Mode 1) global mode similar to unshielded thrusters at low magnetic fields, Mode 2) cathode oscillations at nominal magnetic fields, and Mode 3) combined spoke, cathode and breathing mode oscillations at high magnetic fields. Mode 1 exhibits large amplitude, low frequency (1-10 kHz), breathing mode type oscillations where discharge current mean value and oscillation amplitude peak. The mean discharge current is minimized while thrust-to-power and anode efficiency are maximized in Mode 2, where higher frequency (50-90 kHz), low amplitude, cathode oscillations dominate. Thrust is maximized in Mode 3 and decreases by 5-6% with decreasing magnetic field strength. The presence or absence of spokes and strong cathode oscillations do not affect each other or discharge current. Similar to unshielded thrusters, mode transitions and plasma oscillations affect magnetically shielded thruster performance and should be characterized during system development.

High-Efficiency Nested Hall Thrusters for Robotic Solar System Exploration

NASA Technical Reports Server (NTRS)

Hofer, Richard R.

2013-01-01

This work describes the scaling and design attributes of Nested Hall Thrusters (NHT) with extremely large operational envelopes, including a wide range of throttleability in power and specific impulse at high efficiency (>50%). NHTs have the potential to provide the game changing performance, powerprocessing capabilities, and cost effectiveness required to enable missions that cannot otherwise be accomplished. NHTs were first identified in the electric propulsion community as a path to 100- kW class thrusters for human missions. This study aimed to identify the performance capabilities NHTs can provide for NASA robotic and human missions, with an emphasis on 10-kW class thrusters well-suited for robotic exploration. A key outcome of this work has been the identification of NHTs as nearly constant-efficiency devices over large power throttling ratios, especially in direct-drive power systems. NHT systems sized for robotic solar system exploration are predicted to be capable of high-efficiency operation over nearly their entire power throttling range. A traditional Annular Hall Thruster (AHT) consists of a single annular discharge chamber where the propellant is ionized and accelerated. In an NHT, multiple annular channels are concentrically stacked. The channels can be operated in unison or individually depending on the available power or required performance. When throttling an AHT, performance must be sacrificed since a single channel cannot satisfy the diverse design attributes needed to maintain high thrust efficiency. NHTs can satisfy these requirements by varying which channels are operated and thereby offer significant benefits in terms of thruster performance, especially under deep power throttling conditions where the efficiency of an AHT suffers since a single channel can only operate efficiently (>50%) over a narrow power throttling ratio (3:1). Designs for 10-kW class NHTs were developed and compared with AHT systems. Power processing systems were

Life and Operating Range Extension of the BPT-4000 Qualification Model Hall Thruster

NASA Technical Reports Server (NTRS)

Welander, Ben; Carpenter, Christian; deGrys, Kristi; Hofer, Richard R.; Randolph, Thomas M.; Manzella, David H.

2006-01-01

Following completion of the 5,600 hr qualification life test of the BPT-4000 4.5 kW Hall Thruster Propulsion System, NASA and Aerojet have undertaken efforts to extend the qualified operating range and lifetime of the thruster to support a wider range of NASA missions. The system was originally designed for orbit raising and stationkeeping applications on military and commercial geostationary satellites. As such, it was designed to operate over a range of power levels from 3 to 4.5 kW. Studies of robotic exploration applications have shown that the cost savings provided by utilizing commercial technology that can operate over a wider range of power levels provides significant mission benefits. The testing reported on here shows that the 4.5 kW thruster as designed has the capability to operate efficiently down to power levels as low as 1 kW. At the time of writing, the BPT-4000 qualification thruster and cathode have accumulated over 400 hr of operation between 1 to 2 kW with an additional 600 hr currently planned. The thruster has demonstrated no issues with longer duration operation at low power.

Comparison of Hall Thruster Plume Expansion Model with Experimental Data (Preprint)

DTIC Science & Technology

2006-07-01

Cartesian mesh. AQUILA, the focus of this study, is a hybrid PIC model that tracks particles along an unstructured tetrahedral mesh. COLISEUM is capable...measurements of the ion current density profile, ion energy distributions, and ion species fraction distributions using a nude Faraday probe...Spacecraft and Rockets, Vol.37 No.1. 6 Oh, D. and Hastings, D., “Three Dimensional PIC -DSMC Simulations of Hall Thruster Plumes and Analysis for

On limitations of laser-induced fluorescence diagnostics for xenon ion velocity distribution function measurements in Hall thrusters

NASA Astrophysics Data System (ADS)

Romadanov, I.; Raitses, Y.; Diallo, A.; Hara, K.; Kaganovich, I. D.; Smolyakov, A.

2018-03-01

Hall thruster operation is characterized by strong breathing oscillations of the discharge current, the plasma density, the temperature, and the electric field. Probe- and laser-induced fluorescence (LIF) diagnostics were used to measure temporal variations of plasma parameters and the xenon ion velocity distribution function (IVDF) in the near-field plasma plume in regimes with moderate (<18%) external modulations of applied DC discharge voltage at the frequency of the breathing mode. It was shown that the LIF signal collapses while the ion density at the same location is finite. The proposed explanation for this surprising result is based on a strong dependence of the excitation cross-section of metastables on the electron temperature. For large amplitudes of oscillations, the electron temperature at the minimum enters the region of very low cross-section (for the excitation of the xenon ions); thus, significantly reducing the production of metastable ions. Because the residence time of ions in the channel is generally shorter than the time scale of breathing oscillations, the density of the excited ions outside the thruster is low and they cannot be detected. In the range of temperature of oscillations, the ionization cross-section of xenon atoms remains sufficiently large to sustain the discharge. This finding suggests that the commonly used LIF diagnostic of xenon IVDF can be subject to large uncertainties in the regimes with significant oscillations of the electron temperature, or other plasma parameters.

Low frequency azimuthal stability of the ionization region of the Hall thruster discharge. II. Global analysis

DOE Office of Scientific and Technical Information (OSTI.GOV)

Escobar, D.; Ahedo, E., E-mail: eduardo.ahedo@uc3m.es

2015-10-15

The linear stability of the Hall thruster discharge is analysed against axial-azimuthal perturbations in the low frequency range using a time-dependent 2D code of the discharge. This azimuthal stability analysis is spatially global, as opposed to the more common local stability analyses, already afforded previously (D. Escobar and E. Ahedo, Phys. Plasmas 21(4), 043505 (2014)). The study covers both axial and axial-azimuthal oscillations, known as breathing mode and spoke, respectively. The influence on the spoke instability of different operation parameters such as discharge voltage, mass flow, and thruster size is assessed by means of different parametric variations and compared againstmore » experimental results. Additionally, simplified models are used to unveil and characterize the mechanisms driving the spoke. The results indicate that the spoke is linked to azimuthal oscillations of the ionization process and to the Bohm condition in the transition to the anode sheath. Finally, results obtained from local and global stability analyses are compared in order to explain the discrepancies between both methods.« less

A review of studies on ion thruster beam and charge-exchange plasmas

NASA Technical Reports Server (NTRS)

Carruth, M. R., Jr.

1982-01-01

Various experimental and analytical studies of the primary beam and charge-exchange plasmas of ion thrusters are reviewed. The history of plasma beam research is recounted, emphasizing experiments on beam neutralization, expansion of the beam, and determination of beam parameters such as electron temperature, plasma density, and plasma potential. The development of modern electron bombardment ion thrusters is treated, detailing experimental results. Studies on charge-exchange plasma are discussed, showing results such as the relationship between neutralizer emission current and plasma beam potential, ion energies as a function of neutralizer bias, charge-exchange ion current collected by an axially moving Faraday cup-RPA for 8-cm and 30-cm ion thrusters, beam density and potential data from a 15-cm ion thruster, and charge-exchange ion flow around a 30-cm thruster. A 20-cm thruster electrical configuration is depicted and facility effects are discussed. Finally, plasma modeling is covered in detail for plasma beam and charge-exchange plasma.

Overview of NASA's Pulsed Plasma Thruster Development Program

NASA Technical Reports Server (NTRS)

Pencil, Eric J.; Kamhawi, Hani; Arrington, Lynn A.

2004-01-01

NASA's Pulsed Plasma Thruster Program consists of flight demonstration experiments, base research, and development efforts being conducted through a combination of in-house work, contracts, and collaborative programs. The program receives sponsorship from Energetics Project, the New Millennium Program, and the Small Business Innovative Research Program. The Energetics Project sponsors basic and fundamental research to increase thruster life, improve thruster performance, and reduce system mass. The New Millennium Program sponsors the in-orbit operation of the Pulsed Plasma Thruster experiment on the Earth Observing 1 spacecraft. The Small Business Innovative Research Program sponsors the development of innovative diamond-film capacitors, piezoelectric ignitors, and advanced fuels. Programmatic background, recent technical accomplishments, and future activities for each programmatic element are provided.

Development of a 13 kW Hall Thruster Propulsion System Performance Model for AEPS

NASA Technical Reports Server (NTRS)

Stanley, Steven; Allen, May; Goodfellow, Keith; Chew, Gilbert; Rapetti, Ryan; Tofil, Todd; Herman, Dan; Jackson, Jerry; Myers, Roger

2017-01-01

The Advanced Electric Propulsion System (AEPS) program will develop a flight 13kW Hall thruster propulsion system based on NASA's HERMeS thruster. The AEPS system includes the Hall Thruster, the Power Processing Unit (PPU) and the Xenon Flow Controller (XFC). These three primary components must operate together to ensure that the system generates the required combinations of thrust and specific impulse at the required system efficiencies for the desired system lifetime. At the highest level, the AEPS system will be integrated into the spacecraft and will receive power, propellant, and commands from the spacecraft. Power and propellant flow rates will be determined by the throttle set points commanded by the spacecraft. Within the system, the major control loop is between the mass flow rate and thruster current, with time-dependencies required to handle all expected transients, and additional, much slower interactions between the thruster and cathode temperatures, flow controller and PPU. The internal system interactions generally occur on shorter timescales than the spacecraft interactions, though certain failure modes may require rapid responses from the spacecraft. The AEPS system performance model is designed to account for all these interactions in a way that allows evaluation of the sensitivity of the system to expected changes over the planned mission as well as to assess the impacts of normal component and assembly variability during the production phase of the program. This effort describes the plan for the system performance model development, correlation to NASA test data, and how the model will be used to evaluate the critical internal and external interactions. The results will ensure the component requirements do not unnecessarily drive the system cost or overly constrain the development program. Finally, the model will be available to quickly troubleshoot any future unforeseen development challenges.

Theory for the anomalous electron transport in Hall-effect thrusters

NASA Astrophysics Data System (ADS)

Lafleur, Trevor; Baalrud, Scott; Chabert, Pascal

2016-09-01

Using insights from particle-in-cell (PIC) simulations, we develop a kinetic theory to explain the anomalous cross-field electron transport in Hall-effect thrusters (HETs). The large axial electric field in the acceleration region of HETs, together with the radially applied magnetic field, causes electrons to drift in the azimuthal direction with a very high velocity. This drives an electron cyclotron instability that produces large amplitude oscillations in the plasma density and azimuthal electric field, and which is convected downstream due to the large axial ion drift velocity. The frequency and wavelength of the instability are of the order of 5 MHz and 1 mm respectively, while the electric field amplitude can be of a similar magnitude to axial electric field itself. The instability leads to enhanced electron scattering many orders of magnitude higher than that from standard electron-neutral or electron-ion Coulomb collisions, and gives electron mobilities in good agreement with experiment. Since the instability is a strong function of almost all plasma properties, the mobility cannot in general be fitted with simple 1/B or 1/B2 scaling laws, and changes to the secondary electron emission coefficient of the HET channel walls are expected to play a role in the evolution of the instability. This work received financial support from a CNES postdoctoral research award.

Characterizing the Exhaust Plume of the Three-Electrode Micro Pulsed Plasma Thrusters

DTIC Science & Technology

2009-03-01

Plasma Thruster “, J Prop Power 1998;14:716-35 3 W. Andrew Hoskins, Christopher Rayburn, and Charles Sarmiento ” Pulsed Plasma Thruster...plasma thrusters are based on the previous PPT-4 and PPT-7 thruster designs. These thrusters used energy levels between 40 and 80 J generating several...PPT Programs 3 Program Year Energy Voltage Program Year Energy Voltage Zond-2 1964 50 J 1000 V TIP-III 1976 20 J 1630 V LES-6 1968 1.85 J 1360 V NOVA

Assessment of High-Voltage Photovoltaic Technologies for the Design of a Direct Drive Hall Effect Thruster Solar Array

NASA Technical Reports Server (NTRS)

Mikellides, I. G.; Jongeward, G. A.; Schneider, T.; Carruth, M. R.; Peterson, T.; Kerslake, T. W.; Snyder, D.; Ferguson, D.; Hoskins, A.

2004-01-01

A three-year program to develop a Direct Drive Hall-Effect Thruster system (D2HET) begun in 2001 as part of the NASA Advanced Cross-Enterprise Technology Development initiative. The system, which is expected to reduce significantly the power processing, complexity, weight, and cost over conventional low-voltage systems, will employ solar arrays that operate at voltages higher than (or equal to) 300 V. The lessons learned from the development of the technology also promise to become a stepping-stone for the production of the next generation of power systems employing high voltage solar arrays. This paper summarizes the results from experiments conducted mainly at the NASA Marshal Space Flight Center with two main solar array technologies. The experiments focused on electron collection and arcing studies, when the solar cells operated at high voltages. The tests utilized small coupons representative of each solar array technology. A hollow cathode was used to emulate parts of the induced environment on the solar arrays, mostly the low-energy charge-exchange plasma (1012-1013 m-3 and 0.5-1 eV). Results and conclusions from modeling of electron collection are also summarized. The observations from the total effort are used to propose a preliminary, new solar array design for 2 kW and 30-40 kW class, deep space missions that may employ a single or a cluster of Hall- Effect thrusters.

Experimental Investigation of a Direct-drive Hall Thruster and Solar Array System at Power Levels up to 10 kW

NASA Technical Reports Server (NTRS)

Snyder, John S.; Brophy, John R.; Hofer, Richard R.; Goebel, Dan M.; Katz, Ira

2012-01-01

As NASA considers future exploration missions, high-power solar-electric propulsion (SEP) plays a prominent role in achieving many mission goals. Studies of high-power SEP systems (i.e. tens to hundreds of kilowatts) suggest that significant mass savings may be realized by implementing a direct-drive power system, so NASA recently established the National Direct-Drive Testbed to examine technical issues identified by previous investigations. The testbed includes a 12-kW solar array and power control station designed to power single and multiple Hall thrusters over a wide range of voltages and currents. In this paper, single Hall thruster operation directly from solar array output at discharge voltages of 200 to 450 V and discharge powers of 1 to 10 kW is reported. Hall thruster control and operation is shown to be simple and no different than for operation on conventional power supplies. Thruster and power system electrical oscillations were investigated over a large range of operating conditions and with different filter capacitances. Thruster oscillations were the same as for conventional power supplies, did not adversely affect solar array operation, and were independent of filter capacitance from 8 to 80 ?F. Solar array current and voltage oscillations were very small compared to their mean values and showed a modest dependence on capacitor size. No instabilities or anomalous behavior were observed in the thruster or power system at any operating condition investigated, including near and at the array peak power point. Thruster startup using the anode propellant flow as the power 'switch' was shown to be simple and reliable with system transients mitigated by the proper selection of filter capacitance size. Shutdown via cutoff of propellant flow was also demonstrated. A simple electrical circuit model was developed and is shown to have good agreement with the experimental data.

Electromagnetic thrusters for spacecraft prime propulsion

NASA Technical Reports Server (NTRS)

Rudolph, L. K.; King, D. Q.

1984-01-01

The benefits of electromagnetic propulsion systems for the next generation of US spacecraft are discussed. Attention is given to magnetoplasmadynamic (MPD) and arc jet thrusters, which form a subset of a larger group of electromagnetic propulsion systems including pulsed plasma thrusters, Hall accelerators, and electromagnetic launchers. Mission/system study results acquired over the last twenty years suggest that for future prime propulsion applications high-power self-field MPD thrusters and low-power arc jets have the greatest potential of all electromagnetic thruster systems. Some of the benefits they are expected to provide include major reductions in required launch mass compared to chemical propulsion systems (particularly in geostationary orbit transfer) and lower life-cycle costs (almost 50 percent less). Detailed schematic drawings are provided which describe some possible configurations for the various systems.

Performance Characterization of the Air Force Transformational Satellite 12 kW Hall Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Haag, Thomas W.; Smith, Timothy; Herman, Daniel; Huang, Wensheng; Shastry, Rohit; Peterson, Peter; Mathers, Alex

2013-01-01

The STMD GCD ISP project is tasked with developing, maturing, and testing enabling human exploration propulsion requirements and potential designs for advanced high-energy, in-space propulsion systems to support deep-space human exploration and reduce travel time between Earth's orbit and future destinations for human activity. High-power Hall propulsion systems have been identified as enabling technologies and have been the focus of the activities at NASA Glenn-In-house effort to evaluate performance and interrogate operation of NASA designed and manufactured Hall thrusters. Evaluate existing high TRL EP devices that may be suitable for implementation in SEP TDM.

Design of a High-Energy, Two-Stage Pulsed Plasma Thruster

NASA Technical Reports Server (NTRS)

Markusic, T. E.; Thio, Y. C. F.; Cassibry, J. T.; Rodgers, Stephen L. (Technical Monitor)

2002-01-01

Design details of a proposed high-energy (approx. 50 kJ/pulse), two-stage pulsed plasma thruster are presented. The long-term goal of this project is to develop a high-power (approx. 500 kW), high specific impulse (approx. 7500 s), highly efficient (approx. 50%),and mechanically simple thruster for use as primary propulsion in a high-power nuclear electric propulsion system. The proposed thruster (PRC-PPT1) utilizes a valveless, liquid lithium-fed thermal plasma injector (first stage) followed by a high-energy pulsed electromagnetic accelerator (second stage). A numerical circuit model coupled with one-dimensional current sheet dynamics, as well as a numerical MHD simulation, are used to qualitatively predict the thermal plasma injection and current sheet dynamics, as well as to estimate the projected performance of the thruster. A set of further modelling efforts, and the experimental testing of a prototype thruster, is suggested to determine the feasibility of demonstrating a full scale high-power thruster.

Design and Preliminary Performance Testing of Electronegative Gas Plasma Thruster

NASA Technical Reports Server (NTRS)

Liu, Thomas M.; Schloeder, Natalie R.; Walker, Mitchell L. R.; Polzin, Kurt A.; Dankanich, John W.; Aanesland, Ane

2014-01-01

In classical gridded electrostatic ion thrusters, positively charged ions are generated from a plasma discharge of noble gas propellant and accelerated to provide thrust. To maintain overall charge balance on the propulsion system, a separate electron source is required to neutralize the ion beam as it exits the thruster. However, if high-electronegativity propellant gases (e.g., sulfur hexafluoride) are instead used, a plasma discharge can result consisting of both positively and negatively charged ions. Extracting such electronegative plasma species for thrust generation (e.g., with time-varying, bipolar ion optics) would eliminate the need for a separate neutralizer cathode subsystem. In addition for thrusters utilizing a RF plasma discharge, further simplification of the ion thruster power system may be possible by also using the RF power supply to bias the ion optics. Recently, the PEGASES (Plasma propulsion with Electronegative gases) thruster prototype successfully demonstrated proof-of-concept operations in alternatively accelerating positively and negatively charged ions from a RF discharge of a mixture of argon and sulfur hexafluoride.i In collaboration with NASA Marshall Space Flight Center (MSFC), the Georgia Institute of Technology High-Power Electric Propulsion Laboratory (HPEPL) is applying the lessons learned from PEGASES design and testing to develop a new thruster prototype. This prototype will incorporate design improvements and undergo gridless operational testing and diagnostics checkout at HPEPL in April 2014. Performance mapping with ion optics will be conducted at NASA MSFC starting in May 2014. The proposed paper discusses the design and preliminary performance testing of this electronegative gas plasma thruster prototype.

A unique control system simulator for the evaluation of pulsed plasma thrusters

NASA Technical Reports Server (NTRS)

Dahlgren, J. B.

1973-01-01

Because of the low thrust characteristics of solid-propellant pulsed plasma thrusters and their operational requirement to operate in a vacuum environment, unique and sensitive test techniques are required. A technique evolved for testing and evaluating pulsed plasma thrusters in an open- or closed-loop system mode employs a unique air bearing platform as a single-axis simulator on which the thruster is mounted. The simulator described was developed to evaluate pulsed plasma thrusters in the low micropound range; however, the simulator can be extended to cover the operational range of currently developed millipound thrusters.

Plasma Sheet Velocity Measurement Techniques for the Pulsed Plasma Thruster SIMP-LEX

NASA Technical Reports Server (NTRS)

Nawaz, Anuscheh; Lau, Matthew

2011-01-01

The velocity of the first plasma sheet was determined between the electrodes of a pulsed plasma thruster using three measurement techniques: time of flight probe, high speed camera and magnetic field probe. Further, for time of flight probe and magnetic field probe, it was possible to determine the velocity distribution along the electrodes, as the plasma sheet is accelerated. The results from all three techniques are shown, and are compared for one thruster geometry.

The effects of magnetic nozzle configurations on plasma thrusters

NASA Technical Reports Server (NTRS)

York, Thomas M.

1989-01-01

Plasma thrusters have been operated at power levels from 10kW to 0.1MW. When these devices have had magnetic fields applied to them which form a nozzle configuration for the expanding plasma, they have shown marked increases in exhaust velocity which is in direct proportion to the magnitude of the applied field. Further, recent results have shown that electrode erosion may be influenced by applied magnetic fields. This research is directed to the experimental and computational study of the effects of applied magnetic field nozzles in the acceleration of plasma flows. Plasma source devices which eliminate the plasma interaction in normal thrusters are studied as most basic. Normal thruster configurations will be studied without applied fields and with applied magnetic nozzle fields. Unique computational studies will utilize existing codes which accurately include transport processes. Unique diagnostic studies will support the experimental studies to generate new data. Both computation and diagnostics will be combined to indicate the physical mechanisms and transport properties that are operative in order to allow scaling and accurate prediction of thruster performance.

ECR plasma thruster research - Preliminary theory and experiments

NASA Technical Reports Server (NTRS)

Sercel, Joel C.; Fitzgerald, Dennis J.

1989-01-01

A preliminary theory of the operation of the electron-cyclotron-resonance (ECR) plasma thruster is described along with an outline of recent experiments. This work is presented to communicate the status of an ongoing research effort directed at developing a unified theory to quantitatively describe the operation of the ECR plasma thruster. The theory is presented as a set of nonlinear ordinary differential equations and boundary conditions which describe the plasma density, velocity, and electron temperature. Diagnostic tools developed to measure plasma conditions in the existing research device are described.

2-D Magnetohydrodynamic Modeling of A Pulsed Plasma Thruster

NASA Technical Reports Server (NTRS)

Thio, Y. C. Francis; Cassibry, J. T.; Wu, S. T.; Rodgers, Stephen L. (Technical Monitor)

2002-01-01

Experiments are being performed on the NASA Marshall Space Flight Center (MSFC) MK-1 pulsed plasma thruster. Data produced from the experiments provide an opportunity to further understand the plasma dynamics in these thrusters via detailed computational modeling. The detailed and accurate understanding of the plasma dynamics in these devices holds the key towards extending their capabilities in a number of applications, including their applications as high power (greater than 1 MW) thrusters, and their use for producing high-velocity, uniform plasma jets for experimental purposes. For this study, the 2-D MHD modeling code, MACH2, is used to provide detailed interpretation of the experimental data. At the same time, a 0-D physics model of the plasma initial phase is developed to guide our 2-D modeling studies.

Far-Field Plume Measurements of a Nested-Channel Hall-Effect Thruster (PREPRINT)

DTIC Science & Technology

2010-12-13

nude Faraday probe, retarding potential analyzer, and ExB probe. Data from these probes were used to calculate utilization efficiencies from existing...USA Far-field plume measurements were performed on the X2 nested-channel Hall-effect thruster using an ar- ray of diagnostics, including a nude Faraday...mode to nested-channel mode by utilizing a traditional array of far-field diagnostics, which include a nude Faraday probe, retarding potential analyzer

Effect of Segmented Electrode Length on the Performances of an Aton-Type Hall Thruster

NASA Astrophysics Data System (ADS)

Duan, Ping; Bian, Xingyu; Cao, Anning; Liu, Guangrui; Chen, Long; Yin, Yan

2016-05-01

The influences of the low-emissive graphite segmented electrode placed near the channel exit on the discharge characteristics of a Hall thruster are studied using the particle-in-cell method. A two-dimensional physical model is established according to the Hall thruster discharge channel configuration. The effects of electrode length on the potential, ion density, electron temperature, ionization rate and discharge current are investigated. It is found that, with the increasing of the segmented electrode length, the equipotential lines bend towards the channel exit, and approximately parallel to the wall at the channel surface, the radial velocity and radial flow of ions are increased, and the electron temperature is also enhanced. Due to the conductive characteristic of electrodes, the radial electric field and the axial electron conductivity near the wall are enhanced, and the probability of the electron-atom ionization is reduced, which leads to the degradation of the ionization rate in the discharge channel. However, the interaction between electrons and the wall enhances the near wall conductivity, therefore the discharge current grows along with the segmented electrode length, and the performance of the thruster is also affected. supported by National Natural Science Foundation of China (Nos. 11375039 and 11275034) and the Key Project of Science and Technology of Liaoning Province, China (No. 2011224007) and the Fundamental Research Funds for the Central Universities, China (No. 3132014328)

Electric thruster research

NASA Technical Reports Server (NTRS)

Kaufman, H. R.; Robinson, R. S.

1981-01-01

The multipole discharge chamber of an electrostatic ion thruster is discussed. No reductions in discharge losses were obtained, despite repeated demonstration of anode potentials more positive than the bulk of the discharge plasma. The penalty associated with biased anode operation was reduced as the magnetic integral above the biased anodes was increased. The hollow cathode is discussed. The experimental configuration of the Hall current thruster had a uniform field throughout the ion generation and acceleration regions. To obtain reliable ion generation, it was necessary to reduce the magnetic field strength, to the point where excessive electron backflow was required to establish ion acceleration. The theoretical study of ion acceleration with closed electron drift paths resulted in two classes of solutions. One class has the continuous potential variation in the acceleration region that is normally associated with a Hall current accelerator. The other class has an almost discontinuous potential step near the anode end of the acceleration region. This step includes a significant fraction of the total acceleration potential difference.

Experimental study of effect of magnetic field on anode temperature distribution in an ATON-type Hall thruster

NASA Astrophysics Data System (ADS)

Liu, Jinwen; Li, Hong; Mao, Wei; Ding, Yongjie; Wei, Liqiu; Li, Jianzhi; Yu, Daren; Wang, Xiaogang

2018-05-01

The energy deposition caused by the absorption of electrons by the anode is an important cause of power loss in a Hall thruster. The resulting anode heating is dangerous, as it can potentially reduce the thruster lifetime. In this study, by considering the ring shape of the anode of an ATON-type Hall thruster, the effects of the magnetic field strength and gradient on the anode ring temperature distribution are studied via experimental measurement. The results show that the temperature distribution is not affected by changes in the magnetic field strength and that the position of the peak temperature is essentially unchanged; however, the overall temperature does not change monotonically with the increase of the magnetic field strength and is positively correlated with the change in the discharge current. Moreover, as the magnetic field gradient increases, the position of the peak temperature gradually moves toward the channel exit and the temperature tends to decrease as a whole, regardless of the discharge current magnitude; in any case, the position of the peak temperature corresponds exactly to the intersection of the magnetic field cusp with the anode ring. Further theoretical analysis shows that the electrons, coming from the ionization region, travel along two characteristic paths to reach the anode under the guidance of the cusped magnetic field configuration. The change of the magnetic field strength or gradient changes the transfer of momentum and energy of the electrons in these two paths, which is the main reason for the changes in the temperature and distribution. This study is instructive for matching the design of the ring-shaped anode and the cusp magnetic field of an ATON-type Hall thruster.

Modeling of surface-dominated plasmas: from electric thruster to negative ion source.

PubMed

Taccogna, F; Schneider, R; Longo, S; Capitelli, M

2008-02-01

This contribution shows two important applications of the particle-in-cell/monte Carlo technique on ion sources: modeling of the Hall thruster SPT-100 for space propulsion and of the rf negative ion source for ITER neutral beam injection. In the first case translational degrees of freedom are involved, while in the second case inner degrees of freedom (vibrational levels) are excited. Computational results show how in both cases, plasma-wall and gas-wall interactions play a dominant role. These are secondary electron emission from the lateral ceramic wall of SPT-100 and electron capture from caesiated surfaces by positive ions and atoms in the rf negative ion source.

Elimination of Lifetime Limiting Mechanism of Hall Thrusters

NASA Technical Reports Server (NTRS)

Jacobson, David T. (Inventor); Manzella, David H. (Inventor)

2009-01-01

A Hall thruster includes inner and outer electromagnets, with the outer electromagnet circumferentially surrounding the inner electromagnet along a centerline axis and separated therefrom, inner and outer poles, in physical connection with their respective inner and outer electromagnets, with the inner pole having a mostly circular shape and the outer pole having a mostly annular shape, a discharge chamber separating the inner and outer poles, a combined anode electrode/gaseous propellant distributor, located at an upstream portion of the discharge chamber and supplying propellant gas and an actuator, in contact with a sleeve portion of the discharge chamber. The actuator is configured to extend the sleeve portion or portions of the discharge chamber along the centerline axis with respect to the inner and outer poles.

Transport properties of plasmas in microwave electrothermal thrusters. Master's thesis

DOE Office of Scientific and Technical Information (OSTI.GOV)

Haraburda, S.S.

1990-01-01

The microwave electrothermal thruster is a potential propulsion system for spacecraft applications such as platform station keeping. It is a thruster which allows no contact between the electrodes and the propellant. For this thruster, the electromagnetic energy is transferred to the electrons in the plasma region of the propellant using the TM011 and TM012 modes of a microwave cavity system. The collisional processes by the electrons with the propellant causes transfer of the energy. Work was done to study these processes using several diagnostic techniques - calorimetry, photography, and spectroscopy. Experimental results of these techniques for nitrogen and helium gasesmore » are included. These diagnostic techniques are important in understanding plasma phenomena and designing practical plasma rocket thrusters. In addition, a broad theoretical background is included to provide a fundamental description of the plasma phenomena.« less

Calibrating ion density profile measurements in ion thruster beam plasma

NASA Astrophysics Data System (ADS)

Zhang, Zun; Tang, Haibin; Ren, Junxue; Zhang, Zhe; Wang, Joseph

2016-11-01

The ion thruster beam plasma is characterized by high directed ion velocity (104 m/s) and low plasma density (1015 m-3). Interpretation of measurements of such a plasma based on classical Langmuir probe theory can yield a large experimental error. This paper presents an indirect method to calibrate ion density determination in an ion thruster beam plasma using a Faraday probe, a retarding potential analyzer, and a Langmuir probe. This new method is applied to determine the plasma emitted from a 20-cm-diameter Kaufman ion thruster. The results show that the ion density calibrated by the new method can be as much as 40% less than that without any ion current density and ion velocity calibration.

On matching the anode ring with the magnetic field in an ATON-type Hall effect thruster

NASA Astrophysics Data System (ADS)

Liu, Jinwen; Li, Hong; Zhang, Xu; Ding, Yongjie; Wei, Liqiu; Li, Jianzhi; Yu, Daren; Wang, Xiaogang

2018-06-01

In an ATON-type Hall effect thruster, a ring-shaped anode and a cusped magnetic field intersect the match between the anode shape and the field topology thus must be clarified to optimize the electron transport to the anode and consequently the design of a high-efficiency thruster. By changing the match pattern with both the change in the length of the anode ring and the axial displacement of the cusp magnetic field, this study experimentally investigated the influence of the match pattern on the discharge characteristics of an ATON-type thruster—P100—under the condition of a moderate discharge voltage. The experimental results show that there is a match pattern that always optimizes the performance of the P100 thruster. At the rated operation parameters (300 V of discharge voltage and 5 mg/s of propellant mass flow rate) and the rated magnetic field strength, the observed improvements on thrust (˜79 mN to ˜85 mN) and anode efficiency (˜46% to ˜55%) are significant. Through further theoretical analysis, this study revealed that the change in the characteristics of electron momentum and energy transfer in the near-anode region, induced by the change of the match pattern, is the basic reason. The findings of this work are instructive for both understanding the electron motion in a cusp magnetic field and guiding the design of the anode ring intersecting with a cusp magnetic field in an ATON-type Hall effect thruster.

Effect of matching between the magnetic field and channel length on the performance of low sputtering Hall thrusters

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Boyang, Jia; Sun, Hezhi; Wei, Liqiu; Peng, Wuji; Li, Peng; Yu, Daren

2018-02-01

Discharge characteristics of a non-wall-loss Hall thruster were studied under different channel lengths using a design based on pushing a magnetic field through a double permanent magnet ring. The effect of different magnetic field intensities and channel lengths on ionization, efficiency, and plume divergence angle were studied. The experimental results show that propellant utilization is improved for optimal matching between the magnetic field and channel length. While matching the magnetic field and channel length, the ionization position of the neutral gas changes. The ion flow is effectively controlled, allowing the thrust force, specific impulse, and efficiency to be improved. Our study shows that the channel length is an important design parameter to consider for improving the performance of non-wall-loss Hall thrusters.

Plasma propulsion for space applications

NASA Astrophysics Data System (ADS)

Fruchtman, Amnon

2000-04-01

The various mechanisms for plasma acceleration employed in electric propulsion of space vehicles will be described. Special attention will be given to the Hall thruster. Electric propulsion utilizes electric and magnetic fields to accelerate a propellant to a much higher velocity than chemical propulsion does, and, as a result, the required propellant mass is reduced. Because of limitations on electric power density, electric thrusters will be low thrust engines compared with chemical rockets. The large jet velocity and small thrust of electric thrusters make them most suitable for space applications such as station keeping of GEO communication satellites, low orbit drag compensation, orbit raising and interplanetary missions. The acceleration in the thruster is either thermal, electrostatic or electromagnetic. The arcjet is an electrothermal device in which the propellant is heated by an electric arc and accelerated while passing through a supersonic nozzle to a relatively low velocity. In the Pulsed Plasma Thruster a solid propellant is accelerated by a magnetic field pressure in a way that is similar in principle to pulsed acceleration of plasmas in other, very different devices, such as the railgun or the plasma opening switch. Magnetoplasmadynamic thrusters also employ magnetic field pressure for the acceleration but with a reasonable efficiency at high power only. In an ion thruster ions are extracted from a plasma through a double grid structure. Ion thrusters provide a high jet velocity but the thrust density is low due to space-charge limitations. The Hall thruster, which in recent years has enjoyed impressive progress, employs a quasi-neutral plasma, and therefore is not subject to a space-charge limit on the current. An applied radial magnetic field impedes the mobility of the electrons so that the applied potential drops across a large region inside the plasma. Methods for separately controlling the profiles of the electric and the magnetic fields will

Improvement of the low frequency oscillation model for Hall thrusters

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wang, Chunsheng, E-mail: wangcs@hit.edu.cn; Wang, Huashan

2016-08-15

The low frequency oscillation of the discharge current in Hall thrusters is a major aspect of these devices that requires further study. While the existing model captures the ionization mechanism of the low frequency oscillation, it unfortunately fails to express the dynamic characteristics of the ion acceleration. The analysis in this paper shows this is because of the simplification of the electron equation, which affects both the electric field distribution and the ion acceleration process. Additionally, the electron density equation is revised and a new model that is based on the physical properties of ion movement is proposed.

Effect of plasma distribution on propulsion performance in electrodeless plasma thrusters

NASA Astrophysics Data System (ADS)

Takao, Yoshinori; Takase, Kazuki; Takahashi, Kazunori

2016-09-01

A helicon plasma thruster consisting of a helicon plasma source and a magnetic nozzle is one of the candidates for long-lifetime thrusters because no electrodes are employed to generate or accelerate plasma. A recent experiment, however, detected the non-negligible axial momentum lost to the lateral wall boundary, which degrades thruster performance, when the source was operated with highly ionized gases. To investigate this mechanism, we have conducted two-dimensional axisymmetric particle-in-cell (PIC) simulations with the neutral distribution obtained by Direct Simulation Monte Carlo (DSMC) method. The numerical results have indicated that the axially asymmetric profiles of the plasma density and potential are obtained when the strong decay of neutrals occurs at the source downstream. This asymmetric potential profile leads to the accelerated ion towards the lateral wall, leading to the non-negligible net axial force in the opposite direction of the thrust. Hence, to reduce this asymmetric profile by increasing the neutral density at downstream and/or by confining plasma with external magnetic field would result in improvement of the propulsion performance. These effects are also analyzed by PIC/DSMC simulations.

Study of the catastrophic discharge phenomenon in a Hall thruster

NASA Astrophysics Data System (ADS)

Ding, Yongjie; Su, Hongbo; Li, Peng; Wei, Liqiu; Li, Hong; Peng, Wuji; Xu, Yu; Sun, Hezhi; Yu, Daren

2017-10-01

In a 1350-W Hall-effect thruster, in which a technique for pushing down the magnetic field is implemented, a catastrophic discharge phenomenon is identified by varying the magnetic field strength while keeping all other operating parameters constant. According to experiments, before and after the discharge catastrophe, the plume changes from focusing state to a divergent state, and discharge parameters such as discharge current and thrust exhibit noticeable changes. The divergence half-angle of the plume increases from 22° to 46°. The oscillation amplitude and mean values of the discharge current significantly increase from 0.8 A to 4 A and from 4.6 A to 6.3 A, respectively, while the thrust increases from 89.3 mN to 91 mN. Analysis of the experimental results shows that as the maximum magnetic field of the thruster we developed is in the plume region, the acceleration occurs in the plume region and a large number of Xe2+ ions appear in the plume area, the catastrophic discharge phenomenon observed.

Ion properties in a Hall current thruster operating at high voltage

DOE Office of Scientific and Technical Information (OSTI.GOV)

Garrigues, L., E-mail: laurent.garrigues@laplace.univ-tlse.fr

2016-04-28

Operation of a 5 kW-class Hall current Thruster for various voltages from 400 V to 800 V and a xenon mass flow rate of 6 mg s{sup −1} have been studied with a quasi-neutral hybrid model. In this model, anomalous electron transport is fitted from ion mean velocity measurements, and energy losses due to electron–wall interactions are used as a tuned parameter to match expected electron temperature strength for same class of thruster. Doubly charged ions production has been taken into account and detailed collisions between heavy species included. As the electron temperature increases, the main channel of Xe{sup 2+} ion production becomes stepwisemore » ionization of Xe{sup +} ions. For an applied voltage of 800 V, the mass utilization efficiency is in the range of 0.8–1.1, and the current fraction of doubly charged ions varies between 0.1 and 0.2. Results show that the region of ion production of each species is located at the same place inside the thruster channel. Because collision processes mean free path is larger than the acceleration region, each type of ions experiences same potential drop, and ion energy distributions of singly and doubly charged are very similar.« less

Stationary Plasma Thruster Plume Characteristics

NASA Technical Reports Server (NTRS)

Myers, Roger M.; Manzella, David H.

1994-01-01

Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions. This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements. Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties. For the design operating voltage of 300 V the centerline electron density was found to decrease from approximately 1.8 x 10 exp 17 cubic meters at a distance of 0.3 m to 1.8 X 10 exp 14 cubic meters at a distance of 4 m from the thruster. The electron temperature over the same region was between 1.7 and 3.5 eV. Ion current density measurements showed that the plume was sharply peaked, dropping by a factor of 2.6 within 22 degrees of centerline. The ion energy 4 m from the thruster and 15 degrees off-centerline was approximately 270 V. The thruster cathode flow rate and facility pressure were found to strongly affect the plume properties. In addition to the plume measurements, the data from the various probe types were used to assess the impact of probe design criteria

Magnetic Field Design for a Strongly Improved PHALL Thruster

NASA Astrophysics Data System (ADS)

Martins, Alexandre A.; Rodrigo, Miranda; Ferreira, José Leonardo

2017-10-01

In this article, we are going to go through some steps that we took in the refining of engineering work related to the development of a permanent magnet Hall thruster. The use of permanent magnets in these thrusters is mainly related to the decrease of used power for propulsion, especially important for low power thrusters as for micro-satellites. The advantage of our chosen configuration is that the magnetic field can be used either perpendicular or parallel to the thruster channel walls, whereas in the last case the generated erosion forces are strongly reduced by at least three orders of magnitude. We are going to show how each magnetic field configuration affects the generated plasma and consequently the generated propulsion force and efficiency.

Implementation and Initial Validation of a 100-Kilowatt Class Nested-Channel Hall Thruster

NASA Technical Reports Server (NTRS)

Hall, Scott J.; Florenz, Roland E.; Gallimore, Alec D.; Kamhawi, Hani; Brown, Daniel L.; Polk, James E.; Goebel, Dan; Hofer, Richard R.

2014-01-01

The X3 is a 100-kilowatt class nested-channel Hall thruster developed by the Plasmadynamics and Electric Propulsion Laboratory at the University of Michigan in collaboration with the Air Force Research Laboratory and NASA. The cathode, magnetic circuit, boron nitride channel rings, and anodes all required specific design considerations during thruster development, and thermal modeling was used to properly account for thermal growth in material selection and component design. A number of facility upgrades were required at the University of Michigan to facilitate operation of the X3. These upgrades included a re-worked propellant feed system, a completely redesigned power and telemetry break-out box, and numerous updates to thruster handling equipment. The X3 was tested on xenon propellant at two current densities, 37% and 73% of the nominal design value. It was operated to a maximum steady-state discharge power of 60.8 kilowatts. The tests presented here served as an initial validation of thruster operation. Thruster behavior was monitored with telemetry, photography and high-speed current probes. The photography showed a uniform plume throughout testing. At constant current density, reductions in mass flow rate of 18% and 26% were observed in the three-channel operating configuration as compared to the superposition of each channel running individually. The high-speed current probes showed that the thruster was stable at all operating points and that the channels influence each other when more than one is operating simultaneously. Additionally, the ratio of peak-to-peak AC-coupled discharge current oscillations to mean discharge current did not exceed 51% for any operating points reported here, and did not exceed 17% at the higher current density.

Multi-Axis Thrust Measurements of the EO-1 Pulsed Plasma Thruster

NASA Technical Reports Server (NTRS)

Arrington, Lynn A.; Haag, Thomas W.

1999-01-01

Pulsed plasma thrusters are low thrust propulsive devices which have a high specific impulse at low power. A pulsed plasma thruster is currently scheduled to fly as an experiment on NASA's Earth Observing-1 satellite mission. The pulsed plasma thruster will be used to replace one of the reaction wheels. As part of the qualification testing of the thruster it is necessary to determine the nominal thrust as a function of charge energy. These data will be used to determine control algorithms. Testing was first completed on a breadboard pulsed plasma thruster to determine nominal or primary axis thrust and associated propellant mass consumption as a function of energy and then later to determine if any significant off-axis thrust component existed. On conclusion that there was a significant off-axis thrust component with the bread-board in the direction of the anode electrode, the test matrix was expanded on the flight hardware to include thrust measurements along all three orthogonal axes. Similar off-axis components were found with the flight unit.

Performance Characterization of a Novel Plasma Thruster to Provide a Revolutionary Operationally Responsive Space Capability with Micro- and Nano-Satellites

DTIC Science & Technology

2011-03-24

and radiation resistance of rare earth permanent magnets for applications such as ion thrusters and high efficiency Stirling Radioisotope Generators...from Electron Transitioning Discharge Current Discharge Power Discharge Voltage Θ Divergence Angle Earths Gravity at Sea Level...Hall effect thruster HIVAC High Voltage Hall Accelerator LEO Low Earth Orbit LDS Laser Displacement System LVDT Linear variable differential

Plasma Measurements in an Integrated-System FARAD Thruster

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Rose, M. F.; Miller, R.; Best, S.

2007-01-01

Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a current sheet in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current and the induced magnetic field. The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster[1,2] is a type of pulsed inductive plasma accelerator in which the plasma is preionized by a mechanism separate from that used to form the current sheet and accelerate the gas. Employing a separate preionization mechanism allows for the formation of an inductive current sheet at much lower discharge energies and voltages than those used in previous pulsed inductive accelerators like the Pulsed Inductive Thruster (PIT). A benchtop FARAD thruster was designed following guidelines and similarity performance parameters presented in Refs. [3,4]. This design is described in detail in Ref. [5]. In this paper, we present the temporally and spatially resolved measurements of the preionized plasma and inductively-accelerated current sheet in the FARAD thruster operating with a Vector Inversion Generator (VIG) to preionize the gas and a Bernardes and Merryman circuit topology to provide inductive acceleration. The acceleration stage operates on the order of 100 J/pulse. Fast-framing photography will be used to produce a time-resolved, global view of the evolving current sheet. Local diagnostics used include a fast ionization gauge capable of mapping the gas distribution prior to plasma initiation; direct measurement of the induced magnetic field using B-dot probes, induced azimuthal current measurement using a mini-Rogowski coil, and direct probing of the number density and electron temperature using triple probes.

Prediction of plasma properties in mercury ion thrusters

NASA Technical Reports Server (NTRS)

Longhurst, G. R.

1978-01-01

A simplified theoretical model was developed which obtains to first order the plasma properties in the discharge chamber of a mercury ion thruster from basic thruster design and controllable operating parameters. The basic operation and design of ion thrusters is discussed, and the important processes which influence the plasma properties are described in terms of the design and control parameters. The conservation for mass, charge and energy were applied to the ion production region, which was defined as the region of the discharge chamber having as its outer boundary the surface of revolution of the innermost field line to intersect the anode. Mass conservation and the equations describing the various processes involved with mass addition and removal from the ion production region are satisfied by a Maxwellian electron density spatial distribution in that region.

Pulsed plasma thruster contamination studies

NASA Technical Reports Server (NTRS)

Rudolph, L. K.; Jones, R. M.

1979-01-01

The exhaust plume of the one millipound pulsed plasma thruster has a measurable backflow upstream of the nozzle exit plane which may deposit on and degrade the performance of exposed spacecraft surfaces. High speed photographs and Faraday cup measurements suggest that this backflow is predominantly an electrically neutral, relatively low energy vapor. Articulated collimator quartz crystal microbalance measurements of this backflow were made for a thruster with a radically modified nozzle and a flat plate backflow shield, to determine the backflow sensitivity to nozzle design changes. The results are compared with the original nozzle backflow and show a measurable reduction in the backflow directly upstream of the shield.

Interaction of a solar array with an ion thruster due to the charge-exchange plasma

NASA Technical Reports Server (NTRS)

Kaufman, H. R.

1976-01-01

The generation of a charge exchange plasma by a thruster, the transport of this plasma to the solar array, and the interaction of the solar array with the plasma after it arrives are all described. The generation of this plasma is described accurately from thruster geometry and operating conditions. The transport of the charge exchange plasma was studied experimentally with a 15 cm thruster. A model was developed for simple thruster array configurations. A variety of experiments were surveyed for the interaction of the plasma at the solar array.

Ion ejection from a permanent-magnet mini-helicon thruster

NASA Astrophysics Data System (ADS)

Chen, Francis F.

2014-09-01

A small helicon source, 5 cm in diameter and 5 cm long, using a permanent magnet (PM) to create the DC magnetic field B, is investigated for its possible use as an ion spacecraft thruster. Such ambipolar thrusters do not require a separate electron source for neutralization. The discharge is placed in the far-field of the annular PM, where B is fairly uniform. The plasma is ejected into a large chamber, where the ion energy distribution is measured with a retarding-field energy analyzer. The resulting specific impulse is lower than that of Hall thrusters but can easily be increased to relevant values by applying to the endplate of the discharge a small voltage relative to spacecraft ground.

Microwave processes in the SPD-ATON stationary plasma thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kirdyashev, K. P., E-mail: kpk@ms.ire.rssi.ru

2016-09-15

Results of experimental studies of microwave processes accompanying plasma acceleration in the SPD-ATON stationary plasma thruster are presented. Specific features of the generation of microwave oscillations in both the acceleration channel and the plasma flow outgoing from the thruster are analyzed on the basis of local measurements of the spectra of the plasma wave fields. Mechanisms for generation of microwave oscillations are considered with allowance for the inhomogeneity of the electron density and magnetic field behind the edge of the acceleration channel. The effect of microwave oscillations on the electron transport and the formation of the discharge current in themore » acceleration channel is discussed.« less

RHETT2/EPDM Hall Thruster Propulsion System Electromagnetic Compatibility Evaluation

NASA Technical Reports Server (NTRS)

Sarmiento, Charles J.; Sankovic, John M.; Freitas, Joseph; Lynn, Peter R.

1997-01-01

Electromagnetic compatibility measurements were obtained as part of the Electric Propulsion Demonstration Module (EPDM) flight qualification program. Tests were conducted on a Hall thruster system operating at a nominal 66O W discharge power. Measurements of conducted and radiated susceptibility and emissions were obtained and referenced to MEL-STD-461 C. The power processor showed some conducted susceptibility below 4 kHz for the magnet current and discharge voltage. Radiated susceptibility testing yielded a null result. Conducted emissions showed slight violations of the specified limit for MIL-461C CE03. Radiated emissions exceeded the RE02 standard at low frequencies, below 300 MHz, by up to 40 dB RV/m/MHz.

Development of High-Power Hall Thruster Power Processing Units at NASA GRC

NASA Technical Reports Server (NTRS)

Pinero, Luis R.; Bozak, Karin E.; Santiago, Walter; Scheidegger, Robert J.; Birchenough, Arthur G.

2015-01-01

NASA GRC successfully designed, built and tested four different power processor concepts for high power Hall thrusters. Each design satisfies unique goals including the evaluation of a novel silicon carbide semiconductor technology, validation of innovative circuits to overcome the problems with high input voltage converter design, development of a direct-drive unit to demonstrate potential benefits, or simply identification of lessonslearned from the development of a PPU using a conventional design approach. Any of these designs could be developed further to satisfy NASA's needs for high power electric propulsion in the near future.

Plasma property and performance prediction for mercury ion thrusters

NASA Technical Reports Server (NTRS)

Longhurst, G. R.; Wilbur, P. J.

1979-01-01

The discharge chambers of mercury ion thrusters are modelled so the principal effects and processes which govern discharge plasma properties and thruster performance are described. The conservation relations for mass, charge and energy when applied to the Maxwellian electron population in the ion production region yield equations which may be made one-dimensional by the proper choice of coordinates. Solutions to these equations with the appropriate boundary conditions give electron density and temperature profiles which agree reasonably well with measurements. It is then possible to estimate plasma properties from thruster design data and those operating parameters which are directly controllable. By varying the operating parameter inputs to the computer code written to solve these equations, perfromance curves are obtained which agree quite well with measurements.

Vacuum arc plasma thrusters with inductive energy storage driver

NASA Technical Reports Server (NTRS)

Schein, Jochen (Inventor); Gerhan, Andrew N. (Inventor); Woo, Robyn L. (Inventor); Au, Michael Y. (Inventor); Krishnan, Mahadevan (Inventor)

2004-01-01

An apparatus for producing a vacuum arc plasma source device using a low mass, compact inductive energy storage circuit powered by a low voltage DC supply acts as a vacuum arc plasma thruster. An inductor is charged through a switch, subsequently the switch is opened and a voltage spike of Ldi/dt is produced initiating plasma across a resistive path separating anode and cathode. The plasma is subsequently maintained by energy stored in the inductor. Plasma is produced from cathode material, which allows for any electrically conductive material to be used. A planar structure, a tubular structure, and a coaxial structure allow for consumption of cathode material feed and thereby long lifetime of the thruster for long durations of time.

Theory and Numerical Simulation of Plasma-wall Interactions in Electric Propulsion

NASA Astrophysics Data System (ADS)

Mikellides, Ioannis

2016-10-01

Electric propulsion (EP) can be an enabling technology for many science missions considered by NASA because it can produce high exhaust velocities, which allow for less propellant mass compared to typical chemical systems. Over the last decade two EP technologies have emerged as primary candidates for several proposed science missions, mainly due to their superior performance and proven record in space flight: the Ion and Hall thrusters. As NASA looks ahead to increasingly ambitious science goals, missions demand higher endurance from the propulsion system. So, by contrast to the early years of development of these thrusters, when the focus was on performance, considerable focus today is shifting towards extending their service life. Considering all potentially life-limiting mechanisms in Ion and Hall thrusters two are of primary concern: (a) the erosion of the acceleration channel in Hall thrusters and (b) the erosion of the hollow cathode. The plasma physics leading to material wear in these devices are uniquely challenging. For example, soon after the propellant is introduced into the hollow cathode it becomes partially ionized as it traverses a region of electron emission. Electron emission involves highly non-linear boundary conditions. Also, the sheath size is typically many times smaller than the characteristic physical scale of the device, yet energy gained by ions through the sheath must be accounted for in the erosion calculations. The plasma-material interactions in Hall thruster channels pose similar challenges that are further exacerbated by the presence of a strong applied magnetic field. In this presentation several complexities associated with plasma-wall interactions in EP will be discussed and numerical simulation results of key plasma properties in two examples, Hall thrusters and hollow cathodes, will be presented.

Evaluation of Low Power Hall Thruster Propulsion

NASA Technical Reports Server (NTRS)

Manzella, David; Oleson, Steve; Sankovic, John; Haag, Tom; Semenkin, Alexander; Kim, Vladimir

1996-01-01

Hall thruster systems based on the SPT-50 and the TAL D-38 were evaluated and mission studies were performed. The 0.3 kilowatt SPT-50 operated with a specific impulse of 1160 seconds and an efficiency of 0.32. The 0.8 kilowatt D-38 provided a specific impulse above 1700 seconds at an efficiency of 0.5. The D-38 system was shown to offer a 56 kilogram propulsion system mass savings over a 101 kilogram hydrazine monopropellant system designed to perform North-South station keeping maneuvers on board a 430 kilogram geostationary satellite. The SPIT-50 system offered a greater than 50% propulsion system mass reduction in comparison to the chemical system on board a 200 kilogram low Earth orbit spacecraft performing two orbit raises and drag makeup over two years. The performance characteristics of the SPF-50 were experimentally evaluated at a number of operating conditions. The ion current density distribution of this engine was measured. The performance and system mass benefits of advanced systems based on both engines were considered.

Experimental Investigation of the Near-Wall Region in the NASA HiVHAc EDU2 Hall Thruster

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Kamhawi, Hani; Huang, Wensheng; Haag, Thomas W.

2015-01-01

The HiVHAc propulsion system is currently being developed to support Discovery-class NASA science missions. Presently, the thruster meets the required operational lifetime by utilizing a novel discharge channel replacement mechanism. As a risk reduction activity, an alternative approach is being investigated that modifies the existing magnetic circuit to shift the ion acceleration zone further downstream such that the magnetic components are not exposed to direct ion impingement during the thruster's lifetime while maintaining adequate thruster performance and stability. To measure the change in plasma properties between the original magnetic circuit configuration and the modified, "advanced" configuration, six Langmuir probes were flush-mounted within each channel wall near the thruster exit plane. Plasma potential and electron temperature were measured for both configurations across a wide range of discharge voltages and powers. Measurements indicate that the upstream edge of the acceleration zone shifted downstream by as much as 0.104 channel lengths, depending on operating condition. The upstream edge of the acceleration zone also appears to be more insensitive to operating condition in the advanced configuration, remaining between 0.136 and 0.178 channel lengths upstream of the thruster exit plane. Facility effects studies performed on the original configuration indicate that the plasma and acceleration zone recede further upstream into the channel with increasing facility pressure. These results will be used to inform further modifications to the magnetic circuit that will provide maximum protection of the magnetic components without significant changes to thruster performance and stability.

Satellite propulsion spectral signature detection and analysis through Hall effect thruster plume and atmospheric modeling

NASA Astrophysics Data System (ADS)

Wheeler, Pamela; Cobb, Richard; Hartsfield, Carl; Prince, Benjamin

2016-09-01

Space Situational Awareness (SSA) is of utmost importance in today's congested and contested space environment. Satellites must perform orbital corrections for station keeping, devices like high efficiency electric propulsion systems such as a Hall effect thrusters (HETs) to accomplish this are on the rise. The health of this system is extremely important to ensure the satellite can maintain proper position and perform its intended mission. Electron temperature is a commonly used diagnostic to determine the efficiency of a hall thruster. Recent papers have coordinated near infrared (NIR) spectral measurements of emission lines in xenon and krypton to electron temperature measurements. Ground based observations of these spectral lines could allow the health of the thruster to be determined while the satellite is in operation. Another issue worth considering is the availability of SSA assets for ground-based observations. The current SSA architecture is limited and task saturated. If smaller telescopes, like those at universities, could successfully detect these signatures they could augment data collection for the SSA network. To facilitate this, precise atmospheric modeling must be used to pull out the signature. Within the atmosphere, the NIR has a higher transmission ratio and typical HET propellants are approximately 3x the intensity in the NIR versus the visible spectrum making it ideal for ground based observations. The proposed research will focus on developing a model to determine xenon and krypton signatures through the atmosphere and estimate the efficacy through ground-based observations. The model will take power modes, orbit geometries, and satellite altitudes into consideration and be correlated with lab and field observations.

Hall Propulsion Technology Development, NASA Glenn Research Center: 50 kW Thruster Technology EXPRESS Ground/Space Correlation

NASA Technical Reports Server (NTRS)

Jankovsky, Robert; Elliott, Fred

2000-01-01

It is the goal of this activity to develop 50 kW class Hall thruster technology in support of cost and time critical mission applications such as orbit insertion. NASA Marshall Space Flight Center is tasked to develop technologies that enable cost and travel time reduction of interorbital transportation. Therefore, a key challenge is development of moderate specific impulse (2000-3000 s), high thrust-to-power electric propulsion. NASA Glenn Research Center is responsible for development of a Hall propulsion system to meet these needs. First-phase, sub-scale Hall engine development completed. A 10 kW engine designed, fabricated, and tested. Performance demonstrated >2400 s, >500 mN thrust over 1000 hours of operation documented.

Low Cost Electric Propulsion Thruster for Deep Space Robotic Science Missions

NASA Technical Reports Server (NTRS)

Manzella, David

2008-01-01

Electric Propulsion (EP) has found widespread acceptance by commercial satellite providers for on-orbit station keeping due to the total life cycle cost advantages these systems offer. NASA has also sought to benefit from the use of EP for primary propulsion onboard the Deep Space-1 and DAWN spacecraft. These applications utilized EP systems based on gridded ion thrusters, which offer performance unequaled by other electric propulsion thrusters. Through the In-Space Propulsion Project, a lower cost thruster technology is currently under development designed to make electric propulsion intended for primary propulsion applications cost competitive with chemical propulsion systems. The basis for this new technology is a very reliable electric propulsion thruster called the Hall thruster. Hall thrusters, which have been flown by the Russians dating back to the 1970s, have been used by the Europeans on the SMART-1 lunar orbiter and currently employed by 15 other geostationary spacecraft. Since the inception of the Hall thruster, over 100 of these devices have been used with no known failures. This paper describes the latest accomplishments of a development task that seeks to improve Hall thruster technology by increasing its specific impulse, throttle-ability, and lifetime to make this type of electric propulsion thruster applicable to NASA deep space science missions. In addition to discussing recent progress on this task, this paper describes the performance and cost benefits projected to result from the use of advanced Hall thrusters for deep space science missions.

Performance, Facility Pressure Effects, and Stability Characterization Tests of NASA's Hall Effect Rocket with Magnetic Shielding Thruster

NASA Technical Reports Server (NTRS)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Herman, Daniel; Peterson, Peter Y.; Williams, George J.; Gilland, James; Hofer, Richard; Mikellides, Ioannis

2016-01-01

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. Part of the technology maturation effort included experimental evaluation of the TDU-1 thruster with conducting and dielectric front pole cover materials in two different electrical configurations. A graphite front magnetic pole cover thruster configuration with the thruster body electrically tied to cathode, and an alumina front pole cover thruster configuration with the thruster body floating were evaluated. Both configurations were also evaluated at different facility background pressure conditions to evaluate background pressure effects on thruster operation. Performance characterization tests found that higher thruster performance was attained with the graphite front pole cover configuration with the thruster electrically tied to cathode. A total thrust efficiency of 68% and a total specific impulse of 2,820 s was demonstrated at a discharge voltage of 600 V and a discharge power of 12.5 kW. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations and with maps of the discharge current-voltage-magnetic field (IVB). Analysis of TDU-1 discharge current waveforms found that lower normalized discharge current peak-to-peak and root mean square magnitudes were attained when the thruster was electrically floated with alumina front pole covers. Background pressure effects characterization tests indicated that the thruster performance and stability were mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 1×10-5 Torr-Xe (for thruster flow rates of 20.5 mg/s). Power spectral density analysis of the discharge current waveforms showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant breathing mode frequency. Finally, IVB

Ion ejection from a permanent-magnet mini-helicon thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Chen, Francis F.

2014-09-15

A small helicon source, 5 cm in diameter and 5 cm long, using a permanent magnet (PM) to create the DC magnetic field B, is investigated for its possible use as an ion spacecraft thruster. Such ambipolar thrusters do not require a separate electron source for neutralization. The discharge is placed in the far-field of the annular PM, where B is fairly uniform. The plasma is ejected into a large chamber, where the ion energy distribution is measured with a retarding-field energy analyzer. The resulting specific impulse is lower than that of Hall thrusters but can easily be increased to relevant valuesmore » by applying to the endplate of the discharge a small voltage relative to spacecraft ground.« less

Development of optical diagnostics for performance evaluation of arcjet thrusters

NASA Technical Reports Server (NTRS)

Cappelli, Mark A.

1995-01-01

Laser and optical emission-based measurements have been developed and implemented for use on low-power hydrogen arcjet thrusters and xenon-propelled electric thrusters. In the case of low power hydrogen arcjets, these laser induce fluorescence measurements constitute the first complete set of data that characterize the velocity and temperature field of such a device. The research performed under the auspices of this NASA grant includes laser-based measurements of atomic hydrogen velocity and translational temperature, ultraviolet absorption measurements of ground state atomic hydrogen, Raman scattering measurements of the electronic ground state of molecular hydrogen, and optical emission based measurements of electronically excited atomic hydrogen, electron number density, and electron temperature. In addition, we have developed a collisional-radiative model of atomic hydrogen for use in conjunction with magnetohydrodynamic models to predict the plasma radiative spectrum, and near-electrode plasma models to better understand current transfer from the electrodes to the plasma. In the final year of the grant, a new program aimed at developing diagnostics for xenon plasma thrusters was initiated, and results on the use of diode lasers for interrogating Hall accelerator plasmas has been presented at recent conferences.

Electric thruster research

NASA Technical Reports Server (NTRS)

Kaufman, H. R.; Robinson, R. S.

1982-01-01

It has been customary to assume that ions flow nearly equally in all directions from the ion production region within an electron-bombardment discharge chamber. In general, the electron current through a magnetic field can alter the electron density, and hence the ion density, in such a way that ions tend to be directed away from the region bounded by the magnetic field. When this mechanism is understood, it becomes evident that many past discharge chamber designs have operated with a preferentially directed flow of ions. Thermal losses were calculated for an oxide-free hollow cathode. At low electron emissions, the total of the radiation and conduction losses agreed with the total discharge power. At higher emissions, though, the plasma collisions external to the cathode constituted an increasingly greater fraction of the discharge power. Experimental performance of a Hall-current thruster was adversely affected by nonuniformities in the magnetic field, produced by the cathode heating current. The technology of closed-drift thrusters was reviewed. The experimental electron diffusion in the acceleration channel was found to be within about a factor of 3 of the Bohm value for the better thruster designs at most operating conditions. Thruster efficiencies of about 0.5 appear practical for the 1000 to 2000 s range of specific impulse. Lifetime information is limited, but values of several thousands of hours should be possible with anode layer thrusters operated or = to 2000 s.

Optimal Quasi-steady Plasma Thruster system characteristics.

NASA Technical Reports Server (NTRS)

Ludwig, D. E.; Kelly, A. J.

1972-01-01

The overall characteristics of a generalized Quasi-steady Plasma Thruster (QPT) system consisting of thruster head, power conditioning network, propellant supply subsystem are studied. Energy balance equations for the system are coupled with component mass relationships in order to determine overall system mass and performance. Power supply power levels varying from 100 to 10,000 watts with thruster power levels ranging from 300 kw to 30 Mw employing argon as the propellant are considered. The manner in which overall system mass, average thrust, and burn time vary as a function power supply power level, quasi-steady power level, and pulse time are studied. Results indicate the existence of optimum pulse times when system mass is employed as an optimization criterion.

Hybrid-PIC Simulation of Backsputtered Carbon Transport in the Near-Field Plume of a Hall Thruster

NASA Technical Reports Server (NTRS)

Choi, Maria; Yim, John T.; Williams, George J.; Herman, Daniel A.; Gilland, James H.

2017-01-01

Magnetic shielding has eliminated boron nitride erosion as the life limiting mechanism in a Hall thruster but has resulted in erosion of the front magnetic field pole pieces. Recent experiments show that the erosion of graphite pole covers, which are added to protect the magnetic field pole pieces, causes carbon to redeposit on other surfaces, such as boron nitride discharge channel and cathode keeper surfaces. As a part of the risk-reduction activities for AEPS thruster development, this study models transport of backsputtered carbon from the graphite front pole covers and vacuum facility walls. Fluxes, energy distributions, and redeposition rates of backsputtered carbon on the anode, discharge channel, and graphite cathode keeper surfaces are predicted.

Azimuthal Spoke Propagation in Hall Effect Thrusters

DTIC Science & Technology

2013-08-01

on mode transitions clearly shows that spoke behavior was dominant in so-called ”local oscillation mode” where the thruster exhibited lower mean...discharge current and discharge current oscillation amplitude. The H6 thrust-to-power are maximum when the thruster is operating in local mode with spokes...the H6 drives us to understand the fundamental mechanisms of spoke mechanics in order to improve thruster operation. II. Mode Transition Oscillations

Hardware in the Loop Testing of an Iodine-Fed Hall Thruster

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Peeples, Steven R.; Cecil, Jim; Lewis, Brandon L.; Molina Fraticelli, Jose C.; Clark, James P.

2015-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload,1 providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature ( less than100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum

Wear Testing of the HERMeS Thruster

NASA Technical Reports Server (NTRS)

Williams, George J.; Gilland, James H.; Peterson, Peter Y.; Kamhawi, Hani; Huang, Wensheng; Ahern, Drew W.; Yim, John; Herman, Daniel A.; Hofer, Richard R.; Sekerak, Michael

2016-01-01

The Hall-Effect Rocket with Magnetic Shielding (HERMeS) thruster is being developed and tested at NASA GRC and NASA JPL through support of the Space Technology Mission Directorate (STMD) as primary propulsion for the Asteroid Rendezvous and Redirect Mission (ARRM). This thruster is advancing the state of the art of hall-effect thrusters (HETs) and is intended to serve as a precursor to higher power systems for human interplanetary exploration. The HERMeS Thruster Demonstration Unit One (TDU-1) has entered a 2000-hour wear test campaign at NASA GRC and has completed the first three of four test segments totaling 728 hours of operation. This is the first test of a NASA-designed magnetically shielded thruster to extend beyond 300 hours of continuous operation.

Fluctuations, Electron Transport, and Flow Shear in 2D Axial, Azimuthal (z-θ) Hybrid Hall Thruster Simulations.

NASA Astrophysics Data System (ADS)

Fernandez, Eduardo; Gascon, Nicolas; Knoll, Aaron; Scharfe, Michelle; Cappelli, Mark

2007-11-01

Motivated by the inability of radial-axial (r-z) simulations to properly treat cross-field electron transport in Hall thrusters, a novel 2D z-θ model has been implemented. In common with many r-z descriptions, the simulation is hybrid in nature and assumes quasi-neutrality. Unlike r-z models, electron transport is not enhanced with an ad-hoc mobility coefficient; instead it is given by collisional or ``classical'' terms as well as ``anomalous'' contributions associated with azimuthal electric field fluctuations. Results indicate that anomalous transport dominates classical transport for most of the channel and near field, except in a strong electron flow shear region near the channel exit. The correlation between flow shear, fluctuation behavior, and electron transport will be examined, along with experimental data from the Stanford Hall Thruster. Our findings make a strong link to the turbulent transport suppression mechanism by flow shear seen in fusion devices. The scheme for combining the r-z and z-θ descriptions into an upcoming 3D hybrid model will be presented.

Diagnostics of Plasma Propulsion Devices

NASA Astrophysics Data System (ADS)

Cappelli, Mark A.

1998-11-01

Plasma rockets are rapidly emerging as critical technologies in future space flight. These devices take on various forms, ranging from electro-thermal to electromagnetic accelerators, generally categorized by the method in which electrical energy is converted to thrust. As is the case in many plasma devices, non-intrusive optical (emission, or laser-based) diagnostics is an essential element in the characterization of these plasma sources, as access to the discharges in these plasma engines is often limited. Furthermore, laser-based diagnostics offer additional benefits, including improved spatial resolution, and can provide state-specific measurements of species densities, velocities and energy distributions. In recent years, we have developed and applied a variety of emission and laser-based diagnostics strategies to the characterization of arcjet plasma and closed-drift xenon Hall plasma accelerators. Both of these types of plasma propulsion devices are of immediate interest to the space propulsion community, and are under varying stages of development. Arcjet thrusters have unique properties, with strong plasma density, temperature and velocity gradients, which enhance the coupling between the gasdynamic and plasma physics. Closed-drift Hall plasma thrusters are low density electrostatic devices that are inherently turbulent, and exhibit varying degrees of anomalous cross-field electron transport. Our most extensive, collective effort has been to apply laser-induced fluorescence, Doppler-free laser absorption, and Raman scattering to the characterization of hydrogen and helium arcjet flows. Detailed measurements of velocity, temperatures, and electron densities are compared to the results of magneto-hydrodynamic flowfield simulations. The results show that while the simulations capture many aspects of the flow, there are still some unresolved discrepancies. The database established for Hall thrusters is less extensive, as the laser absorption spectroscopy of

Plasma processes in inert gas thrusters

NASA Technical Reports Server (NTRS)

Kaufman, H. R.; Robinson, R. S.

1979-01-01

Inert gas thrusters, particularly with large diameters, have continued to be of interest for space propulsion applications. Two plasma processes are treated in this study: electron diffusion across magnetic fields and double ion production in inert-gas thrusters. A model is developed to describe electron diffusion across a magnetic field that is driven by both density and potential gradients, with Bohm diffusion used to predict the diffusion rate. This model has applications to conduction across magnetic fields inside a discharge chamber, as well as through a magnetic baffle region used to isolate a hollow cathode from the main chamber. A theory for double ion production is presented, which is not as complete as the electron diffusion theory described, but it should be a useful tool for predicting double ion sputter erosion. Correlations are developed that may be used, without experimental data, to predict double ion densities for the design of new and especially larger ion thrusters.

Mode transition induced by the magnetic field gradient in Hall thrusters

NASA Astrophysics Data System (ADS)

Han, Liang; Wei, Liqiu; Yu, Daren

2016-09-01

A mode transition phenomenon was found in Hall thrusters, which was induced by the increase of the magnetic field gradient. In the transition process, we observed experimentally that there have been obvious changes in the oscillation, the mean value of the discharge current, the thrust, the anode efficiency, and the plume pattern. The shifting and compression of the high magnetic field causes the electron density in the discharge channel to decrease and the ionization zone to move towards the exit plane. This also corresponds to a low atom density in the discharge channel, resulting in a loss of stability of the ionization at a high magnetic field gradient, which presents the transition of the discharge mode.

Three axis pulsed plasma thruster with angled cathode and anode strip lines

NASA Technical Reports Server (NTRS)

Cassady, R. Joseph (Inventor); Myers, Roger M. (Inventor); Osborne, Robert D. (Inventor)

2001-01-01

A spacecraft attitude and altitude control system utilizes sets of three pulsed plasma thrusters connected to a single controller. The single controller controls the operation of each thruster in the set. The control of a set of three thrusters in the set makes it possible to provide a component of thrust along any one of three desired axes. This configuration reduces the total weight of a spacecraft since only one controller and its associated electronics is required for each set of thrusters rather than a controller for each thruster. The thrusters are positioned about the spacecraft such that the effect of the thrusters is balanced.

Optimization of a coaxial electron cyclotron resonance plasma thruster with an analytical model

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cannat, F., E-mail: felix.cannat@onera.fr, E-mail: felix.cannat@gmail.com; Lafleur, T.; Laboratoire de Physique des Plasmas, CNRS, Sorbonne Universites, UPMC Univ Paris 06, Univ Paris-Sud, Ecole Polytechnique, 91128 Palaiseau

2015-05-15

A new cathodeless plasma thruster currently under development at Onera is presented and characterized experimentally and analytically. The coaxial thruster consists of a microwave antenna immersed in a magnetic field, which allows electron heating via cyclotron resonance. The magnetic field diverges at the thruster exit and forms a nozzle that accelerates the quasi-neutral plasma to generate a thrust. Different thruster configurations are tested, and in particular, the influence of the source diameter on the thruster performance is investigated. At microwave powers of about 30 W and a xenon flow rate of 0.1 mg/s (1 SCCM), a mass utilization of 60% and amore » thrust of 1 mN are estimated based on angular electrostatic probe measurements performed downstream of the thruster in the exhaust plume. Results are found to be in fair agreement with a recent analytical helicon thruster model that has been adapted for the coaxial geometry used here.« less

Engineering Considerations for the Self-Energizing Magnetoplasmadynamic (MPD)-Type Fusion Plasma Thruster

DTIC Science & Technology

1992-02-01

Feasibility studies Of dense plasma focus (DPF) device as a fusion propulsion thruster have been performed. Both conventional and spin-polarized D...uncertainties remain in the validity of scaling laws on capacitor mass at high current beyond 1 MA. Fusion Propulsion, Dense Plasma Focus , Magnetoplasmadynamic Thruster, Advanced Fuel, D-3He Fusion, Spin-Polarized Fusion.

Plume and Discharge Plasma Measurements of an NSTAR-type Ion Thruster

NASA Technical Reports Server (NTRS)

Foster, John E.; Soulas, George C.; Patterson, Michael J.

2000-01-01

The success of the NASA Deep Space 1 spacecraft has demonstrated that ion propulsion is a viable option for deep space science missions. More aggressive missions such as Comet Nuclear Sample Return and Europa lander will require higher power, higher propellant throughput and longer thruster lifetime than the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) engine. Presented here are thruster plume and discharge plasma measurements of an NSTAR-type thruster operated from 0.5 kW to 5 kW. From Faraday plume sweeps, beam divergence was determined. From Langmuir probe plume measurements on centerline, low energy ion production on axis due to charge-exchange and direct ionization was assessed. Additionally, plume plasma potential measurements made on axis were used to determine the upper energy limits at which ions created on centerline could be radially accelerated. Wall probes flush-mounted to the thruster discharge chamber anode were used to assess plasma conditions. Langmuir probe measurements at the wall indicated significant differences in the electron temperature in the cylindrical and conical sections of the discharge chamber.

Plume and Discharge Plasma Measurements of an NSTAR-type Ion Thruster

NASA Technical Reports Server (NTRS)

Foster, John E; Soulas, George C.; Patterson, Michael J.

2000-01-01

The success of the NASA Deep Space I spacecraft has demonstrated that ion propulsion is a viable option for deep space science missions. More aggressive missions such as Comet Nuclear Sample Return and Europa lander will require higher power, higher propellant throughput and longer thruster lifetime than the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) engine. Presented here are thruster plume and discharge plasma measurements of an NSTAR-type thruster operated from 0.5 kW to 5 kW. From Faraday plume sweeps, beam divergence was determined. From Langmuir probe plume measurements on centerline, low energy ion production on axis due to charge-exchange and direct ionization was assessed. Additionally, plume plasma potential measurements made on axis were used to determine the upper energy limits at which ions created on centerline could be radially accelerated. Wall probes flush-mounted to the thruster discharge chamber anode were used to assess plasma conditions. Langmuir probe measurements at the wall indicated significant differences in the electron temperature in the cylindrical and conical sections of the discharge chamber.

A HiPIMS plasma source with a magnetic nozzle that accelerates ions: application in a thruster

NASA Astrophysics Data System (ADS)

Bathgate, Stephen N.; Ganesan, Rajesh; Bilek, Marcela M. M.; McKenzie, David R.

2017-01-01

We demonstrate a solid fuel electrodeless ion thruster that uses a magnetic nozzle to collimate and accelerate copper ions produced by a high power impulse magnetron sputtering discharge (HiPIMS). The discharge is initiated using argon gas but in a practical device the consumption of argon could be minimised by exploiting the self-sputtering of copper. The ion fluence produced by the HiPIMS discharge was measured with a retarding field energy analyzer (RFEA) as a function of the magnetic field strength of the nozzle. The ion fraction of the copper was determined from the deposition rate of copper as a function of substrate bias and was found to exceed 87%. The ion fluence and ion energy increased in proportion with the magnetic field of the nozzle and the energy of the ions was found to follow a Maxwell-Boltzmann distribution with a directed velocity. The effectiveness of the magnetic nozzle in converting the randomized thermal motion of the ions into a jet was demonstrated from the energy distribution of the ions. The maximum ion exhaust velocity of at least 15.1 km/s, equivalent to a specific impulse of 1543 s was measured which is comparable to existing Hall thrusters and exceeds that of Teflon pulsed plasma thrusters.

Helicon thruster plasma modeling: Two-dimensional fluid-dynamics and propulsive performances

DOE Office of Scientific and Technical Information (OSTI.GOV)

Ahedo, Eduardo; Navarro-Cavalle, Jaume

2013-04-15

An axisymmetric macroscopic model of the magnetized plasma flow inside the helicon thruster chamber is derived, assuming that the power absorbed from the helicon antenna emission is known. Ionization, confinement, subsonic flows, and production efficiency are discussed in terms of design and operation parameters. Analytical solutions and simple scaling laws for ideal plasma conditions are obtained. The chamber model is then matched with a model of the external magnetic nozzle in order to characterize the whole plasma flow and assess thruster performances. Thermal, electric, and magnetic contributions to thrust are evaluated. The energy balance provides the power conversion between ionsmore » and electrons in chamber and nozzle, and the power distribution among beam power, ionization losses, and wall losses. Thruster efficiency is assessed, and the main causes of inefficiency are identified. The thermodynamic behavior of the collisionless electron population in the nozzle is acknowledged to be poorly known and crucial for a complete plasma expansion and good thrust efficiency.« less

Direct thrust measurements and modelling of a radio-frequency expanding plasma thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lafleur, T.; Charles, C.; Boswell, R. W.

2011-08-15

It is shown analytically that the thrust from a simple plasma thruster (in the absence of a magnetic field) is given by the maximum upstream electron pressure, even if the plasma diverges downstream. Direct thrust measurements of a thruster are then performed using a pendulum thrust balance and a laser displacement sensor. A maximum thrust of about 2 mN is obtained at 700 W for a thruster length of 17.5 cm and a flow rate of 0.9 mg s{sup -1}, while a larger thrust of 4 mN is obtained at a similar power for a length of 9.5 cm andmore » a flow rate of 1.65 mg s{sup -1}. The measured thrusts are in good agreement with the maximum upstream electron pressure found from measurements of the plasma parameters and in fair agreement with a simple global approach used to model the thruster.« less

Evaluation of High-Power Solar Electric Propulsion using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo Missions

NASA Technical Reports Server (NTRS)

Frisbee, Robert H.

2006-01-01

This paper presents the results of mission analyses that expose the advantages and disadvantages of high-power (MWe-class) Solar Electric Propulsion (SEP) for Lunar and Mars Cargo missions that would support human exploration of the Moon and Mars. In these analyses, we consider SEP systems using advanced Ion thrusters (the Xenon [Xe] propellant Herakles), Hall thrusters (the Bismuth [Bi] propellant Very High Isp Thruster with Anode Layer [VHITAL], magnetoplasmadynamic (MPD) thrusters (the Lithium [Li] propellant Advanced Lithium-Fed, Applied-field Lorentz Force Accelerator (ALFA2), and pulsed inductive thruster (PIT) (the Ammonia [NH3] propellant Nuclear-PIT [NuPIT]). The analyses include comparison of the advanced-technology propulsion systems (VHITAL, ALFA2, and NuPIT) relative to state-of-theart Ion (Herakles) propulsion systems and quantify the unique benefits of the various technology options such as high power-per-thruster (and/or high power-per-thruster packaging volume), high specific impulse (Isp), high-efficiency, and tankage mass (e.g., low tankage mass due to the high density of bismuth propellant). This work is based on similar analyses for Nuclear Electric Propulsion (NEP) systems.

Capillary Discharge Thruster Experiments and Modeling (Briefing Charts)

DTIC Science & Technology

2016-06-01

Martin1 ERC INC.1, IN-SPACE PROPULSION BRANCH, AIR FORCE RESEARCH LABORATORY EDWARDS AIR FORCE BASE, CA USA Electric propulsion systems June 2016... PROPULSION MODELS & EXPERIMENTS Spacecraft Propulsion Relevant Plasma: From hall thrusters to plumes and fluxes on components Complex reaction physics i.e... Propulsion Plumes FRC Chamber Environment R.S. MARTIN (ERC INC.) DISTRIBUTION A: APPROVED FOR PUBLIC RELEASE; DISTRIBUTION UNLIMITED; PA# 16279 3 / 30 ELECTRIC

A type of cylindrical Hall thruster with a magnetically insulated anode

NASA Astrophysics Data System (ADS)

Yongjie, Ding; Yu, Xu; Wuji, Peng; Liqiu, Wei; Hongbo, Su; Hezhi, Sun; Peng, Li; Hong, Li; Daren, Yu

2017-04-01

In this paper, a type of magnetically insulated anode structure is proposed for the design of a low-power cylindrical Hall thruster. The magnetic field distribution in the channel is guided by the magnetically insulated anode, altering the intersection status of the magnetic field line passing through the anode and wall. Experimental and simulation results show that a high potential is formed near the wall by the magnetically insulated anode. As the ionization moves towards the outlet, the energy and flux of the ions bombarding the channel wall can be reduced effectively. Due to the reduction in the bombardment of the wall from high-energy ions, the thrust and specific impulse greatly increase compared with those of the non-magnetically insulated anode. For anode mass flow rates of 0.3 and 0.35 mg s-1 and discharge voltages in the 100-200 V range, the thrust can be increased by more than 33% and the anode efficiency can be improved by more than 7%. Meanwhile, the length of the sputtering area is clearly reduced. The starting position of the sputtering area is in front of the magnetic pole, which can effectively prolong the service life of the thruster.

Hall Effect Thruster Interactions Data From the Russian Express-A2 and Express-A3 Satellites

NASA Technical Reports Server (NTRS)

Sitnikova, N.; Volkov, D.; Maximov, I.; Petrusevich, V.; Allen, D.

2003-01-01

This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80 E. and 11 W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3 99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.

Single electron dynamics in a Hall thruster electromagnetic field profile

NASA Astrophysics Data System (ADS)

Marini, Samuel; Pakter, Renato

2017-05-01

In this work, the single electron dynamics in a simplified three dimensional Hall thruster model is studied. Using Hamiltonian formalism and the concept of limiting curves, one is able to determine confinement conditions for the electron in the acceleration channel. It is shown that as a given parameter of the electromagnetic field is changed, the particle trajectory may transit from regular to chaotic without affecting the confinement, which allows one to make a detailed analysis of the role played by the chaos. The ionization volume is also computed, which measures the probability of an electron to ionize background gas atoms. It is found that there is a great correlation between chaos and increased effective ionization volume. This indicates that a complex dynamical behavior may improve the device efficiency by augmenting the ionization capability of each electron, requiring an overall lower electron current.

Effects of neutral distribution and external magnetic field on plasma momentum in electrodeless plasma thrusters

NASA Astrophysics Data System (ADS)

Takase, Kazuki; Takahashi, Kazunori; Takao, Yoshinori

2018-02-01

The effects of neutral distribution and an external magnetic field on plasma distribution and thruster performance are numerically investigated using a particle-in-cell simulation with Monte Carlo collisions (PIC-MCC) and the direct simulation Monte Carlo (DSMC) method. The modeled thruster consists of a quartz tube 1 cm in diameter and 3 cm in length, where a double-turn rf loop antenna is wound at the center of the tube and a solenoid is placed between the loop antenna and the downstream tube exit. A xenon propellant is introduced from both the upstream and downstream sides of the thruster, and the flow rates are varied while maintaining the total gas flow rate of 30 μg/s. The PIC-MCC calculations have been conducted using the neutral distribution obtained from the DSMC calculations, which were applied with different strengths of the magnetic field. The numerical results show that both the downstream gas injection and the external magnetic field with a maximum strength near the thruster exit lead to a shift of the plasma density peak from the upstream to the downstream side. Consequently, a larger total thrust is obtained when increasing the downstream gas injection and the magnetic field strength, which qualitatively agrees with a previous experiment using a helicon plasma source.

Hybrid-Particle-In-Cell Simulation of Backsputtered Carbon Transport in the Near-Field Plume of a Hall Thruster

NASA Technical Reports Server (NTRS)

Choi, Maria; Yim, John T.; Williams, George J.; Herman, Daniel A.; Gilland, James H.

2018-01-01

Magnetic shielding has eliminated boron nitride erosion as the life limiting mechanism in a Hall thruster but has resulted in erosion of the front magnetic field pole pieces. Recent experiments show that the erosion of graphite pole covers, which are added to protect the magnetic field pole pieces, causes carbon to redeposit on other surfaces, such as boron nitride discharge channel and cathode keeper surfaces. As a part of the risk-reduction activities for Advanced Electric Propulsion System thruster development, this study models transport of backsputtered carbon from the graphite front pole covers and vacuum facility walls. Fluxes, energy distributions, and redeposition rates of backsputtered carbon on the anode, discharge channel, and graphite cathode keeper surfaces are predicted.

The Iodine Satellite (iSAT) Hall Thruster Demonstration Mission Concept and Development

NASA Technical Reports Server (NTRS)

Dankanich, John W.; Polzin, Kurt A.; Calvert, Derek; Kamhawi, Hani

2014-01-01

The use of iodine propellant for Hall thrusters has been studied and proposed by multiple organizations due to the potential mission benefits over xenon. In 2013, NASA Marshall Space Flight Center competitively selected a project for the maturation of an iodine flight operational feed system through the Technology Investment Program. Multiple partnerships and collaborations have allowed the team to expand the scope to include additional mission concept development and risk reduction to support a flight system demonstration, the iodine Satellite (iSAT). The iSAT project was initiated and is progressing towards a technology demonstration mission preliminary design review. The current status of the mission concept development and risk reduction efforts in support of this project is presented.

Vacuum arc plasma thrusters with inductive energy storage driver

NASA Technical Reports Server (NTRS)

Krishnan, Mahadevan (Inventor)

2009-01-01

A plasma thruster with a cylindrical inner and cylindrical outer electrode generates plasma particles from the application of energy stored in an inductor to a surface suitable for the formation of a plasma and expansion of plasma particles. The plasma production results in the generation of charged particles suitable for generating a reaction force, and the charged particles are guided by a magnetic field produced by the same inductor used to store the energy used to form the plasma.

Performance of a quasi-steady, multi megawatt, coaxial plasma thruster

NASA Technical Reports Server (NTRS)

Scheuer, Jay T.; Schoenberg, Kurt F.; Henins, Ivars; Gerwin, Richard A.; Moses, Ronald W., Jr.; Garcia, Jose A.; Gribble, Robert F.; Hoyt, Robert P.; Black, Dorwin C.; Mayo, Robert M.

1994-01-01

The Los Alamos National Laboratory Coaxial Thruster Experiment (CTX) has been upgraded to enable the quasisteady operation of magnetoplasmadynamic (MPD) type thrusters at power levels from 2 to 40 MW for 10 ms. Diagnostics include an eight position, three axis magnetic field probe to measure magnetic field fluctuations during the pulse; a triple Langmuir probe to measure ion density, electron temperature, and plasma potential; and a time-of-flight neutral particle spectrometer to measure specific impulse. Here we report on the experimental observations and associated analysis and interpretation of long-pulse, quasisteady, coaxial thruster performance in the CTX device.

Internal Plasma Properties and Enhanced Performance of an 8 cm Ion Thruster Discharge

NASA Technical Reports Server (NTRS)

Foster, John E.; Patterson, Michael J.

1999-01-01

There is a need for a lightweight, low power ion thruster for space science missions. Such an ion thruster is under development at NASA Glenn Research Center. In an effort to better understand the discharge performance of this thruster. a version of this thruster with an anode containing electrically isolated electrodes at the cusps was fabricated and tested. Discharge characteristics of this ring cusp ion thruster were measured without ion beam extraction. Discharge current was measured at collection electrodes located at the cusps and at the anode body itself. Discharge performance and plasma properties were measured as a function of discharge power, which was varied between 20 and 50 W. It was found that ion production costs decreased by as much as 20 percent when the two most downstream cusp electrodes were allowed to float. Floating the electrodes did not give rise to a significant increase in discharge power even though the plasma density increased markedly. The improved performance is attributed to enhanced electron containment.

Influence of Triply-Charged Ions and Ionization Cross-Sections in a Hybrid-PIC Model of a Hall Thruster Discharge

NASA Technical Reports Server (NTRS)

Smith, Brandon D.; Boyd, Iain D.; Kamhawi, Hani

2014-01-01

The sensitivity of xenon ionization rates to collision cross-sections is studied within the framework of a hybrid-PIC model of a Hall thruster discharge. A revised curve fit based on the Drawin form is proposed and is shown to better reproduce the measured crosssections at high electron energies, with differences in the integrated rate coefficients being on the order of 10% for electron temperatures between 20 eV and 30 eV. The revised fit is implemented into HPHall and the updated model is used to simulate NASA's HiVHAc EDU2 Hall thruster at discharge voltages of 300, 400, and 500 V. For all three operating points, the revised cross-sections result in an increase in the predicted thrust and anode efficiency, reducing the error relative to experimental performance measurements. Electron temperature and ionization reaction rates are shown to follow the trends expected based on the integrated rate coefficients. The effects of triply-charged xenon are also assessed. The predicted thruster performance is found to have little or no dependence on the presence of triply-charged ions. The fraction of ion current carried by triply-charged ions is found to be on the order of 1% and increases slightly with increasing discharge voltage. The reaction rates for the 0?III, I?III, and II?III ionization reactions are found to be of similar order of magnitude and are about one order of magnitude smaller than the rate of 0?II ionization in the discharge channel.