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NASA NTRS1992

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

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.

Gallimore, Alec D.↗

NASA NTRS2015

High Input Voltage, Silicon Carbide Power Processing Unit Performance Demonstration

A silicon carbide brassboard power processing unit has been developed by the NASA Glenn Research Center in Cleveland, Ohio. The power processing unit operates from two sources - a nominal 300-Volt high voltage input bus and a nominal 28-Volt low voltage input bus. The design of the power processing unit includes four low voltage, low power supplies that provide power to the thruster auxiliary supplies, and two parallel 7.5 kilowatt power supplies that are capable of providing up to 15 kilowatts of total power at 300-Volts to 500-Volts to the thruster discharge supply. Additionally, the unit contains a housekeeping supply, high voltage input filter, low voltage input filter, and master control board, such that the complete brassboard unit is capable of operating a 12.5 kilowatt Hall Effect Thruster. The performance of unit was characterized under both ambient and thermal vacuum test conditions, and the results demonstrate the exceptional performance with full power efficiencies exceeding 97. With a space-qualified silicon carbide or similar high voltage, high efficiency power device, this design could evolve into a flight design for future missions that require high power electric propulsion systems.

Technology↗

NASA NTRS1981

Electric thruster research

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.

Kaufman, H. R.↗

NASA NTRS2007

Sensitivity Testing of the NSTAR Ion Thruster

During the Extended Life Test of the DS1 flight spare ion thruster, the engine was subjected to sensitvity testing in order to characterize the macroscopic dependence of discharge chamber sensitivity to a +\-3% vatiation in main flow, cathode flow and beam current, and to +\5% variation in beam and accelerator voltage, was determined for the minimum- (THO), half- (TH8) and full power (TH15) throttle levels. For each power level investigared, 16 high/low operating conditions were chosen to vary the flows, beam current, and grid voltages in in a matrix that mapped out the entire parameter space. The matrix of data generated was used to determine the partial derivative or senitivity of the dependent parameters--discharge voltage, discharge current, discharge loss, double-to-single-ion current ratio, and neutralizer-keeper voltage--to the variation in the independent parameters--main flow, cathode flow, beam current, and beam voltage. The sensititivities of each dependent parameter with respect to each independent parameter were determined using a least-square fit routine. Variation in these sensitivities with thruster runtime was recorded over the duration of the ELT, to detemine if discharge performance changed with thruster wear. Several key findings have been ascertained from the sensitivity testing. Discharge operation is most sensitve to changes in cathode flow and to a lesser degree main flow. The data also confirms that for the NSTAR configuration plasma production is limited by primary electron input due to the fixed neutral population. Key sensitivities along with their change with thruster wear (operating time) will be presented. In addition double ion content measurements with an ExB probe will also be presented to illustrate beam ion production and content sensitivity to the discharge chamber operating parameteres.

electric propulsion↗

NASA NTRS2015

Testing of an Arcjet Thruster with Capability of Direct-Drive Operation

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft provided power to the voltage-current that a thruster needs for operation. Testing has been initiated to study whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The work aims to identify and address technical questions related to power conditioning and noise suppression in the system and heating of the thruster in long-duration operation. The apparatus under investigation has a target power level from 400-1,000 W. However, the proposed direct-drive arcjet is potentially a highly scalable concept, applicable to solar-electric spacecraft with up to 100's of kW and beyond. A direct-drive electric propulsion system would be comprised of a thruster that operates with the power supplied directly from the power source (typically solar arrays) with no further power conditioning needed between those two components. Arcjet thrusters are electric propulsion devices, with the power supplied as a high current at low voltage; of all the different types of electric thruster, they are best suited for direct drive from solar arrays. One advantage of an arcjet over Hall or gridded ion thrusters is that for comparable power the arcjet is a much smaller device and can provide more thrust and orders of magnitude higher thrust density (approximately 1-10 N/sq m), albeit at lower I(sub sp) (approximately 800-1000 s). In addition, arcjets are capable of operating on a wide range of propellant options, having been demonstrated on H2, ammonia, N2, Ar, Kr, Xe, while present SOA Hall and ion thrusters are primarily limited to Xe propellant. Direct-drive is often discussed in terms of Hall thrusters, but they require 250-300 V for operation, which is difficult even with high-voltage solar arrays. The arcjet requires under 100 V, which is more in-line with what is easily possible with a solar array. Direct-drive of an electric propulsion system confers the advantage of reducing or eliminating the power processing unit (PPU) that is typically needed to convert the spacecraft-provided power to the voltage and current needed for thruster operation. Since the PPU is typically the most expensive piece of an electric thruster system, from both a fabrication and qualification standpoint, its elimination offers the potential for major reductions in system cost and risk. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets. It has a precision machined 99.95% pure tungsten anode which also serves as the nozzle. The anode constrictor region is 1 mm (0.040-in) diameter and 1 mm (0.040-in) long. The cathode is a tungsten welding electrode doped with LaO2; its tip was precision ground to a 30◦ angle ending in a blunt end. The two electrodes are separated by a boron-nitride insulator which also serves as the propellant injection manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer-body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1A shows the assembled thruster in the vacuum chamber; figure 1B shows the thruster in operation on argon at a flow rate of 676 sccm (20 mg/s). Initial testing was conducted in a 3.5-ft diameter vacuum chamber; the ultimate pressure reached during quasi-steady operation of the thruster was about 330 millitorr. The thruster was powered with a high-current, 0-100A, 15 kW power supply. The discharge was initiated with a high-voltage (approximately 10 kV) spark initiator that was isolated from the supply by a stack of diodes. The testing indicated that an operating point exists within the I-V characteristics that is compatible with direct-drive solar-electric operation; for a flow rate of 20 mg/s (argon) the arc could be sustained at a voltage of about 20 V and a current of 25 A (500W).

Martin, Adam K.↗

NASA NTRS2016

The Ion Propulsion System for the Asteroid Redirect Robotic Mission

The Asteroid Redirect Robotic Mission is a Solar Electric Propulsion Technology Demonstration Mission (ARRM) whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of NASA's future beyond-low-Earth-orbit, human-crewed exploration plans. This presentation presents the conceptual design of the ARRM ion propulsion system, the status of the NASA in-house thruster and power processing development activities, the status of the planned technology maturation for the mission through flight hardware delivery, and the status of the mission formulation and spacecraft acquisition.

Flight Systems↗

NASA NTRS2016

The Ion Propulsion System for the Asteroid Redirect Robotic Mission

The Asteroid Redirect Robotic Mission is a Solar Electric Propulsion Technology Demonstration Mission (ARRM) whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of NASA'a future beyond-low-Earth-orbit, human-crewed exploration plans. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. This paper presents the conceptual design of the ARRM ion propulsion system, the status of the NASA in-house thruster and power processing development activities, the status of the planned technology maturation for the mission through flight hardware delivery, and the status of the mission formulation and spacecraft acquisition.

Electric Propulsion↗

NASA NTRSDate not supplied

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

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

DSMC↗

NASA NTRS1982

Electric thruster research

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.

Kaufman, H. R.↗

NASA NTRSDate not supplied

Evaluating the Performance of Hybrid Vehicles between LEO, Cislunar Space, and Mars

Nuclear power and electric propulsion technologies can enable a broad range of existing and future space mission concepts. This paper evaluates how these two technology classes, when combined into systems with various performance levels (and paired with a chemical propulsion system) affect vehicle sizing and mission feasibility for round-trip, opposition-class, crewed missions to Mars. The technology trade-space evaluated in this paper encompasses: various nuclear power system mass-efficiency levels (or power-specific mass), a range of output power levels available to the electric propulsion system, two electric propulsion thruster technologies (Hall and Magnetoplasmadynamic) at various characteristic specific impulse levels, and a range of propellant tank mass sizing coefficient. Each combination of the technologies is evaluated for the 2039 and the 2042 Mars mission opportunities for a range of imposed total mission duration constraints. The results show that synergies between technology performance levels exist and that increasing the performance of some parameters might have undesired consequences for the trajectory and vehicle. It was found that increasing the mass-efficiency of the nuclear power system is broadly beneficial, resulting in lower vehicle masses and enabling shorter mission durations. However, increasing the specific impulse of a thruster or increasing the available power level do not always yield a net benefit to the vehicle or mission; since increasing the specific impulse or power level (holding power system mass-efficiency constant) tends to decrease the vehicle thrust-to-mass ratio. Furthermore, the results show that mission feasibility is highly dependent on the mission duration; the required technology performance level to enable a mission is relaxed as the mission duration increases. Given the optimally sized vehicles resulting from the trade-space for a crewed mission to Mars, this paper also evaluates the impact of aggregating the Mars mission vehicle in Low-Earth Orbit and transferring it to the Mars mission departure node (using the on-board power and propulsion systems). The trade-space for this portion of the analysis includes the option of in-space vehicle refueling at the Mars mission departure node. The results show that on-orbit refueling can relax the required technology performance levels to achieve lower launch masses and orbit-raising durations at the cost of increased architecture complexity and risk.

Nuclear Electric Propulsion↗

NASA NTRS2013

Guide to Flow Measurement for Electric Propulsion Systems

In electric propulsion (EP) systems, accurate measurement of the propellant mass flow rate of gas or liquid to the thruster and external cathode is a key input in the calculation of thruster efficiency and specific impulse. Although such measurements are often achieved with commercial mass flow controllers and meters integrated into propellant feed systems, the variability in potential propellant options and flow requirements amongst the spectrum of EP power regimes and devices complicates meter selection, integration, and operation. At the direction of the Committee on Standards for Electric Propulsion Testing, a guide was jointly developed by members of the electric propulsion community to establish a unified document that contains the working principles, methods of implementation and analysis, and calibration techniques and recommendations on the use of mass flow meters in laboratory and spacecraft electric propulsion systems. The guide is applicable to EP devices of all types and power levels ranging from microthrusters to high-power ion engines and Hall effect thrusters. The establishment of a community standard on mass flow metering will help ensure the selection of the proper meter for each application. It will also improve the quality of system performance estimates by providing comprehensive information on the physical phenomena and systematic errors that must be accounted for during the analysis of flow measurement data. This paper will outline the standard methods and recommended practices described in the guide titled "Flow Measurement for Electric Propulsion Systems."

spacecraft electric propulsion (EP) systems↗

NASA NTRSDate not supplied

Qualification Progress and Mission-Enabling Capabilities of the 12-kW Hall-Effect, Advanced Electric Propulsion System (AEPS) Thruster

An update of the 12 kW Advanced Electric Propulsion System (AEPS) qualification and flight thruster status is provided. Three flight thrusters completed acceptance testing in 2025 and have been delivered to NASA. Environmental qualification is complete and preparations are under way for long-duration life-demonstration testing. Programmatic lessons learned associated with risk management, contract oversight, and requirement definition during the fabrication and assembly are discussed. In addition, the extensibility of the AEPS thruster to enable a variety of NASA missions is presented.

Electric Propulsion↗

NASA NTRSDate not supplied

Qualification Progress and Mission-Enabling Capabilities of the 12-kW Hall-Effect, Advanced Electric Propulsion System (AEPS) Thruster

An update of the 12 kW Advanced Electric Propulsion System (AEPS) qualification and flight thruster status is provided. Three flight thrusters completed acceptance testing in 2025 and have been delivered to NASA. Environmental qualification is complete and preparations are under way for long-duration life-demonstration testing. A brief overview of three of SEP’s programmatic lessons learned are discussed: risk management, contract oversight, and requirement definition. In addition, the extensibility of the AEPS thruster to enable a variety of NASA missions is presented.

Mars Exploration↗

NASA NTRSDate not supplied

Development and Qualification Status of the Electric Propulsion Systems for the NASA PPE Mission and Gateway Program

NASA is charged with landing the first American woman and next American man on the South Pole of the Moon and establishing sustainable lunar exploration by the end of the decade. To meet this challenge, NASA’s Gateway will develop and deploy critical infrastructure required for operations on the lunar surface and that enables a sustained presence on and around the moon. NASA’s Power and Propulsion Element (PPE), the first planned element of NASA’s cis-lunar Gateway, leverages prior and ongoing NASA and U.S. industry investments in high-power, long-life solar electric propulsion technology investments. NASA awarded a PPE contract to Maxar Technologies to provide a 50 kW-class SEP spacecraft that meets Gateway’s needs, aligns with industry’s heritage spacecraft buses, and allows extensibility for NASA’s Mars exploration goals. Maxar’s PPE concept design, is based on their high heritage, modular, and highly reliable 1300-series bus architecture. The electric propulsion system features three 12 kW Advanced Electric Propulsion (AEPS) thrusters from Aerojet Rocketdyne and four BHT-6000 thrusters from Busek. Maxar-provided power electronics and xenon flow controllers from Moog are utilized in both the 12kW and 6kW electric propulsion strings on the spacecraft. The paper will present overviews of NASA’s Gateway and the PPE Project, status of the development and qualification activities for the two electric propulsion system, and the planned implementation of PPE electric propulsion system as keystone of NASA’s Gateway. The PPE spacecraft is currently heading into the Critical Design Review, with the qualification and flight electric propulsion hardware fabrication already initiated and significant progress being made toward planned qualifications in support of the planned PPE spacecraft co-manifest launch in 2024.

Electric Propulsion↗

NASA NTRSDate not supplied

Development and Qualification Status of the Electric Propulsion Systems for the NASA PPE Mission and Gateway Program

Electric Propulsion↗

NASA NTRSDate not supplied

Edge Wear of the Advanced Electric Propulsion System Pole Covers

This work summarizes a wear test focused on characterizing the erosion of the edges and side walls of the 12-kW Advanced Electric Propulsion System (AEPS) pole covers as well as the impact of magnetic field strength on component erosion for operation at a discharge voltage of 600 V. Testing was performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster and accumulated approximately 325 hours of operation at the 600 V/12 kW condition with the magnetic field strength set to 75% of its nominal value. Thruster performance was invariant throughout the wear test for all thruster throttle conditions and matched previous measurements at the nominal magnetic field strength. Contrary to past results, inner front pole cover erosion rates were invariant to magnetic field strength at the 600 V/12 kW condition and no erosion was measured on the outer front pole cover. The erosion profile of the downstream inner front pole cover faces was continuous across the cover, indicating no significant change in erosion processes at the pole cover edges. Side wall erosion was only detected on the cathode-facing surfaces of the inner front pole cover and was equal in magnitude to the rates measured on the adjacent downstream edges. Taken together, these results suggest that erosion of the pole cover edges and sidewalls is driven by cathode-borne ions and will ultimately not impact AEPS life estimates.

Hall thrusters↗

NASA NTRSDate not supplied

Edge Wear of the Advanced Electric Propulsion System Pole Covers

This work presents a summary of the first wear test of the 12 kW Advanced Electric Propulsion System (AEPS) focused on characterizing the erosion of the edges and side walls of the thruster pole covers. Testing was performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster and accumulated approximately 325 hours of operation at the 600 V/12 kW condition with the magnetic field strength set to 75% of its nominal value. Thruster performance was invariant throughout the wear test for all thruster throttle conditions and was found to be equal to previous measurements. Contrary to past results, inner front pole cover erosion rates were invariant to magnetic field strength at the 600 V/12 kW condition and no erosion was measured on the outer front pole cover. The erosion profile of the downstream inner front pole cover faces was continuous across the cover indicating no significant change in erosion process at the pole cover edges. Side wall erosion was only detected on the cathode-facing surfaces of the inner front pole cover and was equal in magnitude to the rates measured on the adjacent downstream edges. Taken together, these results suggest that erosion of the pole cover edges and sidewalls is driven by cathode-borne ions and will ultimately not impact AEPS life estimates.

Hall thrusters↗

NASA NTRS2019

Characterization of a Fixed-Volume Release System for Initiating an Arc Discharge in a Heaterless Hollow Cathode

Heaterless hollow cathodes may provide improved reliability, simplicity, and portability when compared with traditional heater-equipped hollow cathodes, traits which are well suited for low-power Hall-effect thrusters that are currently being developed for small satellite propulsion. Despite the advantages, there are concerns that the ignition process in heaterless hollow cathodes may impose an excessive burden on the propellant feed and/or electrical systems of a small satellite. To address this concern, a fixed-volume release flow protocol, which can be used to temporarily increase the propellant mass flow rate, was developed, modeled, and experimentally evaluated. The new protocol allowed for a heaterless hollow cathode to be ignited reliably with a moderate bias voltage and a minimal electrical power requirement. Specifically, a xenon fed heaterless hollow cathode was ignited with a 375 V bias using 17.3 mg of propellant. Repeating the tests with krypton showed that ignition could be achieved in the same heaterless hollow cathode assembly with a 300 V bias using 13.1 mg of propellant. We judge that a fixed-volume release system could be implemented in a satellite feed system while introducing minimal additional complexity.

Heaterless↗