Electrical Propulsion Group

Pasadena, CA, United States

Electrical Propulsion Group

Pasadena, CA, United States
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Hofer R.R.,Jet Propulsion Laboratory | Hofer R.R.,Electrical Propulsion Group | Randolph T.M.,Jet Propulsion Laboratory | Randolph T.M.,Mission Systems Concepts Group
47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011 | Year: 2011

A model of system mass and life cycle costs is used to determine the optimal number of thrusters for electric propulsion systems. The model is generalized for application with most electric propulsion systems and then applied to high-power Hall thruster systems in particular due to their relevance to human exploration of the solar system. Data from a number of flight and proposed systems were collected and used to construct mass and cost models for individual thruster strings using as inputs the number of active thrusters, the number of redundant thrusters, and the total system power to be divided between the active thruster strings. The string models incorporate all of the major components in the propulsion system: thruster, gimbals, power processing unit, cabling, xenon feed system, tanks, and propellant. Mass and cost are related through the launch cost of the propulsion system mass. This unifies the optimization to a single global parameter based on cost that provides guidance in selecting thruster power level. Fault-tolerance and string cost are driving factors determining the optimum thruster size for a given system power level. Results from the mass and cost models show that Hall thrusters capable of operation over 20-150 kW can support missions from 20 kW up to 1 MW. The flatness of the global cost function around the optimum provides an opportunity to select thruster power level based on other factors. After considering other factors such as fault-tolerance, cost uncertainty, complexity, ground test vacuum facility limitations, previously demonstrated power capabilities, and possible technology limitations, we conclude that thrusters operating less than 100 kW are strong candidates for supporting human exploration operating at several hundred kilowatts. The development of two thrusters to flight status is suggested. A low power model operating at 20-50 kW to support missions up to 500 kW and the development of a high power model operating at 50-100 kW to support missions up to 1 MW. The extensibility of Hall thruster technology to even higher power levels may also be realized with nested Hall thrusters operating at 200-500 kW per thruster that could potentially extend the applicable range of Hall thruster technology to 2-5 MW. It is expected that Hall thruster technology will meet the requirements of human exploration for the foreseeable future. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.


Mikellides I.G.,Jet Propulsion Laboratory | Mikellides I.G.,Electrical Propulsion Group | Katz I.,Jet Propulsion Laboratory | Katz I.,Electrical Propulsion Group | And 5 more authors.
46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit | Year: 2010

In a Qualification Life Test (QLT) of the BPT-4000 Hall thruster that recently accumulated >10,000 h it was found that the erosion of the acceleration channel practically stopped after ∼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." Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.


Sultan A.,Aerojet General Corporation | Koch B.,Aerojet General Corporation | Mathers A.,Aerojet General Corporation | Hofer R.R.,Jet Propulsion Laboratory | Hofer R.R.,Electrical Propulsion Group
47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011 | Year: 2011

Aerojet has completed the development and qualification of a 4.5 kW Hall thruster system to serve GEO satellite applications for station keeping and orbit raising as well as for use as primary propulsion on NASA missions. This 4.5 kW Hall thruster system is currently in use on the Air Force's first Advanced Extremely High Frequency (AEHF) satellite. The system includes the BPT-4000 Hall thruster, Power Processing unit (PPU), Xenon Flow Controller (XFC) and associated electrical harnessing. Hall thruster technology is applicable to a wide range of missions that will employ a variety of power systems and as such Aerojet and the Jet Propulsion Laboratory (JPL) have identified the need to develop a scalable power converter that can be tailored to a given application. The envisioned PPU architecture will be able to function over a wide input voltage range (~2:1), provide efficient throttling over a wide output range (100s of Watts - 10s of kilowatts), and employ a robust topology capable of meeting stringent radiation, derating, and component quality requirements. As an added benefit, a modular power converter will reduce the development time and cost for the next generation of PPUs that will be needed to serve applications requiring variable input voltages, higher output power, and a wider power and voltage throttling range. The modular power converters are capable of accepting unregulated 70-140 V input voltages with an output voltage range of 150-400 V when operated in parallel or 150-800 V when operated in series. A full bridge, phase-shifted, zero voltage switching topology was selected as the most suitable for satisfying the wide input and output range. Two converters were fabricated to demonstrate the individual capability of the module as well as the stacked capability of modules working in series or parallel configurations. Converter power efficiency was measured over the full input and output voltage range. Peak efficiencies meeting the target of 94% were measured. Testing under parallel operation of the converters demonstrated current sharing while series operation demonstrated the full 800 V output capability. Demonstration of these power modules retires the major risk associated with converting Aerojet's existing regulated PPU technology to an expanded architecture compatible with NASA missions and a variety of commercial platforms. © 2011 by Aerojet.


Jorns B.A.,Jet Propulsion Laboratory | Jorns B.A.,Electrical Propulsion Group | Hofer R.R.,Jet Propulsion Laboratory | Hofer R.R.,Electrical Propulsion Group
49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference | Year: 2013

The oscillations from 0-100 kHz in a 6-kW magnetically shielded thruster are experimentally 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 for the purpose of determining 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 7-12 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 centerline-mounted cathode 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.


Shastry R.,University of Michigan | Shastry R.,NASA | Gallimore A.D.,University of Michigan | Hofer R.R.,Jet Propulsion Laboratory | Hofer R.R.,Electrical Propulsion Group
47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011 | Year: 2011

In order to better understand interactions between the plasma and channel walls of a Hall thruster, the near-wall plasma was characterized within the H6 Hall thruster using five flush-mounted Langmuir probes. These probes were placed within the last 15% of the discharge channel and were used to measure plasma potential, electron temperature, and ion number density near the inner and outer channel walls. These data were then compared to prior internal measurements inside the channel using a High-speed Axial Reciprocating Probe stage. Comparison of these data has shown that, at the nominal operating condition of 300 V and 20 mg/s anode flow rate, the plasma near the wall begins to accelerate further upstream than plasma closer to centerline. This shift in acceleration zone creates large radial electric fields (~ 40-50 V/mm) that tend to defocus ions and drive them towards the walls. The shift is likely caused by large plasma density gradients between centerline and the channel walls, creating a significant deviation of equipotentials from magnetic field lines near the walls. Electron temperature axial profiles were found to be largely consistent across the channel, supporting the isothermal assumption along magnetic field lines. The experimental results were also compared to simulation results from the hybrid-PIC program HPHall-2. General agreement was found between simulation and experiment for axial profiles of plasma potential, electron temperature, and ion number density, with minor differences occurring in peak locations. Slight asymmetries in properties were found between the inner and outer channel walls despite the use of a symmetric magnetic field topology. This asymmetry was caused by a difference in the location of the maximum radial magnetic field, resulting in axial shifts of acceleration zone and peak electron temperature. This result is supported by asymmetric erosion profiles after 334 hours of operation, showing increased erosion along the outer wall where acceleration began further upstream. © 2011 by Rohit Shastry. Published by the American Institute of Aeronautics and Astronautics, Inc.


Goebel D.M.,Jet Propulsion Laboratory | Jameson K.K.,Jet Propulsion Laboratory | Jameson K.K.,California Polytechnic State University, San Luis Obispo | Hofer R.R.,Jet Propulsion Laboratory | Hofer R.R.,Electrical Propulsion Group
Journal of Propulsion and Power | Year: 2012

The cathode coupling voltage in Hall thrusters, which is the voltage difference between the cathode and the thruster beam plasma potential, is considered an indicator of the ease with which electrons flow from cathode to anode. Historically, the coupling voltage has been minimized by increasing the amount of propellant injected through the hollow cathode due to early observations that this maximizes the discharge (or anode) efficiency. However, recent experiments described here show that the total thruster efficiency is independent of the cathode flow over the range from 5 to 10% of the propellant injected into the thruster body through the anode. For this reason, cathode flow rates can be reduced closer to the classic plume mode limit characteristic of the hollow cathode design without impacting the total thruster efficiency. Such reductions in cathode flow rate can significantly extend the cathode life, especially for higher-power Hall thrusters with larger discharge currents, where the normal Hall thruster cathode flow split will significantly exceed the optimum level for cathode operation and life. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.


Goebel D.M.,Jet Propulsion Laboratory | Polk J.E.,Jet Propulsion Laboratory | Polk J.E.,Electrical Propulsion Group | Mikellides I.G.,Jet Propulsion Laboratory | Mikellides I.G.,Electrical Propulsion Group
Journal of Propulsion and Power | Year: 2011

It has been observed in ion thruster wear tests that the discharge loss increases with time and that the thruster performance correspondingly degrades. This behavioris usually attributed to an enlargement of the accelerator grid apertures during the tests due to ion erosion, which increases the grid transparency and thereby reduces the neutral gas pressure inside the thruster and decreases the ionization efficiency of the plasma generator. An analysis of thruster life test data using a discharge plasma model shows that this mechanism is in sufficient to explain the observed results. Tests at Jet Propulsion Laboratory in an ion thruster simulator used in a 16,000-h discharge cathode wear test show similar increases in discharge loss with time, in spite of the fact that there are no ion accelerator or grid apertures eroding and that the average pressure in the discharge chamber is essentially constant. Experiments show that increases in keeper orifice diameter cause increases in discharge loss, and this trend is reproduced by two-dimensional numerical simulations that show a reduced ion generation rate in the near-keeper plume regionas the electrode eroded. This cathode-electrode erosion mechanism islikely responsible for roughly half of the total discharge loss increases observed in ion thruster life tests. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc.


Ortega A.L.,Jet Propulsion Laboratory | Ortega A.L.,Electrical Propulsion Group | Mikellides I.G.,Jet Propulsion Laboratory | Mikellides I.G.,Electrical Propulsion Group | And 2 more authors.
52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 | Year: 2016

We present a model that quantifies the magnitude of the ion-acoustic turbulence (IAT) in the plume of hollow cathodes and its effect on the resistivity and ion heating. The model takes the form of a partial differential equation (PDE) that can be solved concurrently with the equations of motion for a partially ionized plasma already included in our numerical code for the simulation of the plasma discharge in hollow cathodes, OrCa2D. We also determine that self-induced magnetic fields are not negligible in hollow cathodes operating at large discharge currents and implement in our code Ampere’s law and modifications to Ohm’s law that account for this effect. Numerical simulations that employed these models show large improvements in our agreement with experimental measurements with respect to a previous model, which assumed complete saturation of the IAT and did not account for the growth stage of the waves. In particular, the model is able to accurately predict the location and magnitude of the maximum resitivity to the electron current along the cathode centerline. © 2016 California Institute of Technology.


Sekerak M.J.,University of Michigan | Sekerak M.J.,Plasmadynamics and Electrical Propulsion Laboratory | Gallimore A.D.,University of Michigan | Gallimore A.D.,Plasmadynamics and Electrical Propulsion Laboratory | And 4 more authors.
Journal of Propulsion and Power | Year: 2016

Mode transitions in a 6 kW laboratory Hall-effect thruster were induced by varying the magnetic field intensity while holding all other operating parameters constant. Ultrafast imaging, discharge current, and thrust measurements were used to characterize the change in discharge channel current density and thruster performance through mode transitions. The modes are described here as global oscillation mode and local oscillation mode. In global mode, the entire discharge channel is oscillating in unison and spokes are either absent or negligible with discharge current oscillation amplitude (root mean square) greater than 10% of the mean value and can even be as high as 100%. In local oscillation mode, perturbations in the discharge current density are seen to propagate in the E × B direction. Spokes are localized oscillations that are typically 10-20% of the mean discharge current density value. The discharge current oscillation amplitude and mean values are significantly lower than global mode. The mode transitions changed with operating conditions, where the transition between global mode and local mode occurred at higher relative magnetic field strengths for higher mass flow rate or higher discharge voltage. The thrust was approximately constant through the mode transition, but the thrust-to-power ratio and anode efficiency decreased significantly in global mode. The peaks in thrust to power and anode efficiency typically occur near the transition point. Thruster performance maps should include variation in discharge current, discharge voltage, and magnetic field, known as ID - VD - Bmaps, at different flow rates to identify transition regions throughout the life of a thruster. These results are used to calculate a transition surface for use by operators to keep the thruster operating in an optimal mode. © Copyright 2015 by University of Michigan.


Snyder J.S.,Jet Propulsion Laboratory | Snyder J.S.,Electrical Propulsion Group | Goebel D.M.,Jet Propulsion Laboratory | Hofer R.R.,Jet Propulsion Laboratory | And 6 more authors.
Journal of Propulsion and Power | Year: 2012

The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type thruster that is baselined for the BepiColombo mission to Mercury and is being qualified for application on high-power communications satellite platforms. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5-4.5 kW. In addition to the typical electrical and flow measurements, an E × B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes, were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. At full power, the T6 produced 143 mN of thrust at a specific impulse of 4120 s and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 s and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) thruster. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium. © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

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