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NASA NTRS2019

Ion Thruster Produced Roll Torque

Disturbances to the ion engine's thrust vector will cause a spacecraft to spin about its axis if left unmanaged. Spin about the yaw and pitch axis can be easily handled by a gimbal with enough authority. Spin about the roll axis however must be handled by additional thrusters or reaction wheels. In order to capitalize on the high efficiency of their thrusters, missions utilizing electric propulsion as primary propulsion generally include long periods of thrusting (several years). It is necessary to quantify and understand the ion thruster produced roll torque because it will define the amount of chemical propellant that must be carried or the lifetime and quantity of momentum wheels required for the mission. The roll torque produced by the NEXT ion thruster is analyzed through a combination of theoretical calculations and magnetic field simulations. Experimental techniques for measuring roll torque and past flight data are also discussed.

Arthur, Neil A.↗

NASA NTRS2006

Effects of an internally-mounted cathode on hall thruster plume properties

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.

Fitzgerald, Dennis J.↗

NASA NTRS1993

Power Processing Units for High Powered Solar Electric Propulsion Using MPD Thrusters

The power of various multimegawatt (MMW) vehicles required for deep space missions such asthe Mars Missions in the stipulated time frame of years 2000-2015 ranges from 5 to 100 MWe. Thispaper summarizes an evaluation of high-power processing units (PPUs) for MMW solar electricpropulsion (SEP) vehicles using advanced magnetoplasma- dynamic (MPD) thrusters. Each PPUprovides DC-DC conversion and voltage matching between the phovoltaic power system and the MPDthrusters. The power system consists of DC-DC converter using MCTs. The PPUs are thencombined with contractors to provide PPU input and output isolation, and to allow switching betweenoperating and spare (redundant) PPUs and thrusters as needed. Based on analyses, it is found thatsignificant economies of scale are possible for PPUs that supply power to MPD thrusters operating at0.1 to 5 MWe per thruster.

Frisbee, R. H.↗

NASA NTRS2018

The Effects of Background Pressure on SPT-140 Hall Thruster Performance

NASA’s planned Psyche mission is scheduled to launch in 2022 and begin a 3.5-year cruise to the metallic asteroid Psyche, where it would examine this unique body. The baseline spacecraft design is a hybrid of JPL’s deep-space heritage subsystems with commercial partner SSL’s power, structure, and SPT-140 electric propulsion subsystems. Since the deepspace implementation of the SPT-140 differs from the commercial implementation, primarily in the need for deep power throttling, characterization of the system at lower powers is necessary. One specific area of interest is the sensitivity of thruster performance to background pressure in ground-based test facilities, which can have an impact on the prediction of in-space performance. Measurements of this pressure dependence were performed on a qualification-model SPT-140 thruster over the 0.9-4.5 kW range of interest for the Psyche mission. Thrust sensitivity to pressure, in an absolute sense, was largest at 4.5 kW and decreased with power until there was little-to-no measurable effect at 0.9 kW. In a relative sense, thrust sensitivity was similar at all powers above 0.9 kW with about 2-4% higher thrust measured at 10 µ µTorr than at the lowest operating pressure. Thruster magnetic field sensitivity, examined as a function of magnet current, did not have a strong dependence on facility pressure. Finally, an investigation of low-power operation at the lowest facility pressure showed that a combination of added cathode keeper current and additional cathode propellant flow significantly mitigated the larger negative cathode-to-ground voltages that were observed. These test results, combined with thruster life test results, inform the selection of proper low-power operating conditions for Psyche.

Mackey, Jonathan A.↗

NASA NTRSDate not supplied

Extended Wear Testing of the 12-kW Advanced Electric Propulsion System Engineering Test Unit Hall Thruster

This work presents a summary of the first wear test of the 12 kW Advanced Electric Propulsion System (AEPS) intended to isolate the impact of discharge current on component wear rates. Testing was led by Aerojet Rocketdyne and performed with the AEPS Engineering Test Unit 2 (ETU-2) thruster. In total, approximately 860 hours of operation were accumulated split between operating conditions of 600 V/9 kW, 600 V/11 kW, and 600 V/12 kW. Thruster performance and stability were invariant throughout the wear test for all thruster throttle conditions and shown to be equal to the values previously measured with ETU-2. Inner front pole cover erosion rates were found to be invariant to discharge current as the measurements at all three operating conditions were equal to within the empirical uncertainty. Outer front pole cover erosion rates were found to be equal for the 600 V/9 kW and 600 V/12 kW conditions with operation at 600 V/11 kW yielding elevated erosion rates. Overall, the data shows that the AEPS thruster design has a high probability of meeting lifetime and performance requirements as the design proceeds to flight production and qualification.

Jason D Frieman↗

NASA NTRS1996

3-D Simulations of NSTAR Ion Thruster Plasma Interactions

Described is a Particle-in-Cell with Monte Carlo Collision code developed to perform detailed three-dimensional ion thruster simulations. To capture the full kinetic behavior of ion thruster plumes, both the electrons and ions are treated as test particles. Simulation results are given of the NSTAR ion thruster under ground test and in space conditions. Numerical results are compared.

NSTAR Ion Thruster Plasma Ionization Particles↗

NASA NTRS1995

(abstract) ARGOS: a System to Monitor Ulysses Nutation and Thruster Firings from Variations of the Spacecraft Radio Signal

Ulysses is a spin-stabilized spacecraft that experienced significant nutation after its launch in October 1990. This was due to the Sun-spacecraft-Earth geometry, and a study of the phenomenon predicted that the nutation would again be a problem during 1994-95. The difficulty of obtaining nutation estimates in real time from the spacecraft telemetry forced the ESA/NASA Ulysses Team to explore alternative information sources. The work performed by the ESA Operations Team provided a model for a system that uses the radio signal strength measurements to monitor the spacecraft dynamics. These measurements (referred to as AGC) are provided once per second by the tracking stations of the DSN. The system was named ARGOS (Attitude Reckoning from Ground Observable Signals) after the ever-vigilant, hundred-eyed giant of Greek Mythology. The ARGOS design also included Doppler processing, because Doppler shifts indicate thruster firings commanded by the active nutation control carried out onboard the spacecraft. While there is some visibility into thruster activity from telemetry, careful processing of the high-sample-rate Doppler data provides an accurate means of detecting the presence and time of thruster firings. DSN Doppler measurements are available at a ten-per-second rate in the same tracking data block as the AGC data.

Ulysses nutation Doppler radio signal strength mea↗

NASA NTRS2005

Hollow cathode and thruster discharge chamber plasma measurements

Due to the successful performance of the NSTAR ion thruster in Deep Space 1 mission, coupled with the recently completed 30,352 hour extended life test (ELT) of the NSTAR flight spare thruster, ion thrusters have become a viable option for future NASA missions. In this paper, detailed measurements of the plasma parameters internal and external to the cathode will presented for the NSTAR cathode up to 13.1A of discharge current and for the NEXIS cathode up to 30A of discharge current.

hollow cathode keeper erosion↗

NASA NTRS2006

VHITAL-160 Thruster Development Status

A general overview on the status of the Very High Isp Thruster with Anode Layer (VHITAL)-160 program is presented. The topics include: 1) Bi TAL Overview; 2) VHITAL Program Overview; 3) Thruster Fabrication; and 4) Thruster Testing.

hall thruster↗

NASA NTRS2006

Ion Thruster Discharge Performance Per Magnetic Field Topography

DC-ION is a detailed computational model for predicting the plasma characteristics of rain-cusp ion thrusters. The advanced magnetic field meshing algorithm used by DC-ION allows precise treatment of the secondary electron flow. This capability allows self-consistent estimates of plasma potential that improves the overall consistency of the results of the discharge model described in Reference [refJPC05mod1]. Plasma potential estimates allow the model to predict the onset of plasma instabilities, and important shortcoming of the previous model for optimizing the design of discharge chambers. A magnetic field mesh simplifies the plasma flow calculations, for both the ions and the secondary electrons, and significantly reduces numerical diffusion that can occur with meshes not aligned with the magnetic field. Comparing the results of this model to experimental data shows that the behavior of the primary electrons, and the precise manner of their confinement, dictates the fundamental efficiency of ring-cusp. This correlation is evident in simulations of the conventionally sized NSTAR thruster (30 cm diameter) and the miniature MiXI thruster (3 cm diameter).

plasma processes↗

NASA NTRS2013

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

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.

Electric Propulsion↗

NASA NTRS2013

Real-Tme Boron Nitride Erosion Measurements of the HiVHAc Thruster via Cavity Ring-Down Spectroscopy

Cavity ring-down spectroscopy was used to make real-time erosion measurements from the NASA High Voltage Hall Accelerator thruster. The optical sensor uses 250 nm light to measure absorption of atomic boron in the plume of an operating Hall thruster. Theerosion rate of the High Voltage Hall Accelerator thruster was measured for discharge voltages ranging from 330 to 600 V and discharge powers ranging from 1 to 3 kW. Boron densities as high as 6.5 x 10(exp 15) per cubic meter were found within the channel. Using a very simple boronvelocity model, approximate volumetric erosion rates between 5.0 x 10(exp -12) and 8.2 x 10(exp -12) cubic meter per second were found.

Hall thruster↗

NASA NTRS2013

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

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.

Electric propulsion↗

NASA NTRS2014

Influence of Triply-Charged Ions and Ionization Cross-Sections in a Hybrid-PIC Model of a Hall Thruster Discharge

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.

Hall Thrusters↗

NASA NTRS2013

Azimuthal Spoke Propagation in Hall Effect Thrusters

Spokes are azimuthally propagating perturbations in the plasma discharge of Hall Effect Thrusters (HETs) that travel in the E x B direction and have been observed in many different systems. The propagation of azimuthal spokes are investigated in a 6 kW HET known as the H6 using ultra-fast imaging and azimuthally spaced probes. A spoke surface is a 2-D plot of azimuthal light intensity evolution over time calculated from 87,500 frames/s videos. The spoke velocity has been determined using three methods with similar results: manual fitting of diagonal lines on the spoke surface, linear cross-correlation between azimuthal locations and an approximated dispersion relation. The spoke velocity for three discharge voltages (300, 400 and 450 V) and three anode mass flow rates (14.7, 19.5 and 25.2 mg/s) yielded spoke velocities between 1500 and 2200 m/s across a range of normalized magnetic field settings. The spoke velocity was inversely dependent on magnetic field strength for low B-field settings and asymptoted at B-field higher values. The velocities and frequencies are compared to standard drifts and plasma waves such as E x B drift, electrostatic ion cyclotron, magnetosonic and various drift waves. The empirically approximated dispersion relation yielded a characteristic velocity that matched the ion acoustic speed for 5 eV electrons that exist in the near-anode and near-field plume regions of the discharge channel based on internal measurements. Thruster performance has been linked to operating mode where thrust-to-power is maximized when azimuthal spokes are present so investigating the underlying mechanism of spokes will benefit thruster operation.

Cylindrical Hall Thruster (CHT)↗

NASA NTRS2015

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

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.

Hall Thrusters↗

NASA NTRS2016

Status of the Development of Flight Power Processing Units for the NASAs Evolutionary Xenon Thruster - Commercial (NEXT-C) Project

A pathfinder prototype unit and two flight power processing units (PPUs) are being developed by the Aerojet Rocketdyne Corporation in Redmond, Washington and ZIN Technologies in Cleveland, Ohio, in support of the NEXT-C Project. This project is being led by the NASA Glenn Research Center in Cleveland, Ohio, and will also yield two flight thrusters. This hardware is being considered to be provided as Government Furnished Equipment for the New Frontiers Program, and is applicable to a variety of planetary science missions and astrophysics science missions. The design of the NEXT-C PPU evolves from the hardware fabricated under the NEXT technology development project. The power processing unit operates from two sources: a wide input 80 to 160 V high-power bus and a nominal 28 V low-power bus. The unit includes six power supplies. Four power supplies (beam, accelerator, discharge, and neutralizer keeper) are needed for steady state operation, while two cathode heater power supplies (neutralizer and discharge) are utilized during thruster startup. The unit in total delivers up to 7 kW of regulated power to a single gridded-ion thruster. Significant modifications to the initial design include: high-power adaptive-delay control, upgrade of design to EEE-INST-002 compliance, telemetry accuracy improvements, incorporation of telemetry to detect plume-mode operation, and simplification of the design in select areas to improve manufacturability and commercialization potential. The project is presently in the prototype phase and preparing for qualification level environmental testing.

NASAaEuro(TM)s Evolutionary Xenon Thruster↗

NASA NTRS2016

Post-Test Inspection of NASA's Evolutionary Xenon Thruster Long-Duration Test Hardware: Discharge and Neutralizer Cathodes

The NEXT Long-Duration Test is part of a comprehensive thruster service life assessment intended to demonstrate overall throughput capability, validate service life models, quantify wear rates as a function of time and operating condition, and identify any unknown life-limiting mechanisms. The test was voluntarily terminated in February 2014 after demonstrating 51,184 hours of high-voltage operation, 918 kg of propellant throughput, and 35.5 MN-s of total impulse. The post-test inspection of the thruster hardware began shortly afterwards with a combination of non-destructive and destructive analysis techniques, and is presently nearing completion. This paper presents relevant results of the post-test inspection for both discharge and neutralizer cathodes. Discharge keeper erosion was found to be significantly reduced from what was observed in the NEXT 2 kh wear test and NSTAR Extended Life Test, providing adequate protection of vital cathode components throughout the test with ample lifetime remaining. The area of the discharge cathode orifice plate that was exposed by the keeper orifice exhibited net erosion, leading to cathode plate material building up in the cathode-keeper gap and causing a thermally-induced electrical short observed during the test. Significant erosion of the neutralizer cathode orifice was also found and is believed to be the root cause of an observed loss in flow margin. Deposition within the neutralizer keeper orifice as well as on the downstream surface was thicker than expected, potentially resulting in a facility-induced impact on the measured flow margin from plume mode. Neutralizer keeper wall erosion on the beam side was found to be significantly lower compared to the NEXT 2 kh wear test, likely due to the reduction in beam extraction diameter of the ion optics that resulted in decreased ion impingement. Results from the post-test inspection have led to some minor thruster design improvements.

Electric Propulsion↗