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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Performance Evaluation of an Expanded Range XIPS Ion Thruster System for NASA Science Missions

This paper examines the benefit that a solar electric propulsion (SEP) system based on the 5 kW Xenon Ion Propulsion System (XIPS) could have for NASA's Discovery class deep space missions. The relative cost and performance of the commercial heritage XIPS system is compared to NSTAR ion thruster based systems on three Discovery class reference missions: 1) a Near Earth Asteroid Sample Return, 2) a Comet Rendezvous and 3) a Main Belt Asteroid Rendezvous. It is found that systems utilizing a single operating XIPS thruster provides significant performance advantages over a single operating NSTAR thruster. In fact, XIPS performs as well as systems utilizing two operating NSTAR thrusters, and still costs less than the NSTAR system with a single operating thruster. This makes XIPS based SEP a competitive and attractive candidate for Discovery class science missions.

ion thrusters↗
NASA NTRS2005

The Minimum Impulse Thruster

The Minimum Impulse Thruster (MIT) was developed to improve the state-of-the-art minimum impulse capability of hydrazine monopropellant thrusters. Specifically, a new fast response solenoid valve was developed, capable of responding to a much shorter electrical pulse width, thereby reducing the propellant flow time and the minimum impulse bit. The new valve was combined with the Aerojet MR-103, 0.2 lbf (0.9 N) thruster and put through an extensive Delta-qualification test program, resulting in a factor of 5 reduction in the minimum impulse bit, from roughly 1.1 milli-lbf-seconds (5 milliNewton seconds) to - 0.22 milli-lbf-seconds (1 mN-s). To maintain it's extensive heritage, the thruster itself was left unchanged. The Minimum Impulse Thruster provides mission and spacecraft designers new design options for precision pointing and precision translation of spacecraft.

thruster↗
NASA NTRS2001

Propulsion Instruments for Small Hall Thruster Integration

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.

ion probes↗
NASA NTRS2014

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

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.

Hall Thrusters↗
NASA NTRS2011

Numerical Simulations of a 20-kW Class Hall Thruster Using the Magnetic-Field-Aligned-Mesh Code Hall2De

This paper reports on numerical simulations of the NASA-300M, a 20-kW class Hall thruster developed at the NASA Glenn Research Center (GRC). The numerical simulations have been performed with a 2-D axisymmetric, magnetic field-aligned-mesh (MFAM) plasma solver developed at the Jet Propulsion Laboratory (JPL). The main objective of the collaborative effort is to combine physics-based simulation, plasma diagnostics and recent findings on erosion physics to design and demonstrate a high-power, high-performance Hall thruster that exceeds the life of state-of-the-art Hall thrusters by more than one order of magnitude. The thruster simulations have been carried out at a discharge voltage of 500 V and discharge current of 40 A. The results indicate that although the impact energy of ions may attain values that are comparable to the discharge voltage along the downstream portions of the channel, a withdrawn ionization region and significant ion focusing combine to sustain erosion rates below 1 mm/kh. A more extensive evaluation of the baseline NASA-300M configuration and re-design of this thruster with magnetically shielded walls constitute the main focus of our work in the coming months.

plasma simulations↗
NASA NTRS2015

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

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.

Electric Propulsion↗
NASA NTRS2016

Wear Testing of the HERMeS Thruster

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.

Thruster↗
NASA NTRSDate not supplied

Effects of Background Pressure and Electrical Configuration on the Velocity Field of the HERMeS Hall Thruster

A Laser-Induced Fluorescence test was performed during risk reduction testing on the HERMeS Hall thruster. This article focuses on the portion of the test to study trends in the ion acceleration characteristics with varying anode mass flow rate, background pressure, and electrical configuration. The acceleration zone of the HERMeS Hall thruster was observed to move upstream as the anode mass flow rate decreased and upstream as the background pressure increased. Examination of the characteristics of the ions near the pole covers suggested the possibility that thruster was no longer magnetically shielded at the extreme conditions of very low anode mass flow rate and high background pressure. The acceleration zone was also observed to move slightly upstream when the thruster body was floated as opposed to cathode-tied and did not move appreciably when the thruster body was grounded to the vacuum facility. Characteristics of the ions bombarding the pole covers did not vary across the tested electrical configurations.

Electric Propulsion↗
NASA NTRS2023

Facility Effects on the Ion Characteristics of A 12.5-Kw Hall Thruster

During a laser-induced fluorescence test of a 12.5-kW magnetically-shielded Hall thruster, ion characteristics in the discharge channel and near the poles were measured as the background pressure and electrical configuration were varied. The acceleration zone of the thruster moved upstream by 2% and 10% of the channel length when the background pressure was raised to 1.8 times and 7 times the lowest achievable pressure, respectively. Examination of the characteristics of the ions near the pole covers suggested that as the background pressure decreased the pole covers may be experiencing more erosion. When operating at a discharge voltage of 300 V, the acceleration zone was observed to be at the same location for all electrical configurations. When operating at a discharge voltage of 600V, the acceleration zone was observed to move 3% of the channel length upstream when the thruster body was floated instead of tied to the cathode or grounded to the facility. Characteristics of the ions bombarding the pole covers did not vary across the tested electrical configurations. This observation combined with thruster body voltage measurements suggested that varying the electrical configuration only affected the thruster body sheath voltage and did not affect the plasma potential beyond the sheath.

Electric propulsion↗
NASA NTRSDate not supplied

Testing the Effectiveness of Integrated Plume Shield for a Hall Thruster

Hall thrusters have been shown to exhibit side plumes with varying degree of erosive power. This article presents the concept of the integrated plume shield for the purpose of mitigating the erosive power of said side plume. A test was conducted on a low-power Hall thruster with and without integrated plume shield and the effect on thruster performance and plume characteristics were measured. Collected data shows that the thruster performed identically to within uncertainty in terms of thrust, specific impulse, and discharge oscillation characteristics. Plume measurements showed up to a 48% reduction in ion current density when averaging over 80° to 105° from the firing axis. A brief analytical study was performed to determine the advantages and disadvantages of using integrated plume shield as opposed to using spacecraft surface covering and placing thrusters on booms. The integrated plume shield was found to provide varying level of advantages in mass, design restrictions, and complexity, which are likely to lead to major cost savings in the long term.

Electric Propulsion↗
NASA NTRS2024

The Effect of Impingement Angle on Bipropellant Thruster Plume Degradation of Spacecraft Materials

Unburned and partially combusted liquid propellant present in bipropellant thruster plumes can contaminate and damage external spacecraft materials. Microscopic impact features due to high-speed droplet impacts have been observed on space shuttle flight experiments and International Space Station (ISS) returned flight hardware. Analytical results have shown that particle impingement angle greatly affects surface damage, with normal impacts being the most severe and highly oblique impacts being more benign. The effect of thruster plume impingement angle on material degradation has now been further studied following a bipropellant thruster plume test performed at the high-vacuum plume test facility for chemical thrusters at DLR Göttingen, STG-CT. Several spacecraft material samples were exposed to the freely expanding pulsed thruster plume at a range of impingement angles (from 0 to 75 deg). Results of post-test evaluations and potential spaceflight applications are summarized.

Gateway↗
NASA NTRSDate not supplied

Sub-kW Class Hall-Effect Thruster Power Processing Unit for Wide Output Range Applications

The National Aeronautics and Space Administration (NASA) Small Spacecraft Electric Propulsion (SSEP) project is maturing high-propellant throughput sub-kilowatt Hall-effect thruster technologies to enable small spacecraft deep space science and exploration missions with high delta-v requirements. In support of this effort, development of a power processing unit (PPU) capable of providing discharge power of up to 1 kW continues to be pursued at the NASA Glenn Research Center (GRC). Previous reported work included a successful integrated test of a scalable, modular breadboard discharge power supply with the NASA-H64M laboratory model Hall-effect thruster and presentation of notional designs for the various auxiliary power supplies needed for thruster operation. Since that time, auxiliary power supply designs have been completed and fabricated, with the cathode heater and keeper power supplies being successfully tested with a hollow cathode assembly (HCA) in the NASA GRC Vacuum Facility 56 (VF-56). The desire for a lower mass, higher efficiency, and more versatile PPU to maximize performance of power and mass-limited small spacecraft has led to the exploration of a discharge power supply based on a series-parallel (LCC) resonant topology. This topology has enabled the discharge power supply to operate over a wider output range at switching frequencies 4-5 times higher than previous design iterations. Simulation models of the topology have been developed and a breadboard of the topology has been fabricated and evaluated on both resistive loads and an integrated Hall thruster test. This paper will present collected performance and integrated test data from both the fabricated auxiliary and resonant discharge power supplies. Advantages of the resonant converter architecture over more traditional pulse-width modulated (PWM) techniques in Hall-effect thruster discharge power supply applications will also be described.

electric propulsion↗
NASA NTRSDate not supplied

Sub-kW Class Hall-Effect Thruster Power Processing Unit for Wide Output Range Applications

The National Aeronautics and Space Administration (NASA) Small Spacecraft Electric Propulsion (SSEP) project is maturing high-propellant throughput sub-kilowatt Hall-effect thruster technologies to enable small spacecraft deep space science and exploration missions with high delta-v requirements. In support of this effort, development of a power processing unit (PPU) capable of providing discharge power of up to 1 kW continues to be pursued at the NASA Glenn Research Center (GRC). Previous reported work included a successful integrated test of a scalable, modular breadboard discharge power supply with the NASA-H64M laboratory model Hall-effect thruster and presentation of notional designs for the various auxiliary power supplies needed for thruster operation. Since that time, auxiliary power supply designs have been completed and fabricated, with the cathode heater and keeper power supplies being successfully tested with a hollow cathode assembly (HCA) in the NASA GRC Vacuum Facility 56 (VF-56). The desire for a lower mass, higher efficiency, and more versatile PPU to maximize performance of power and mass-limited small spacecraft has led to the exploration of a discharge power supply based on a series-parallel (LCC) resonant topology. This topology has enabled the discharge power supply to operate over a wider output range at switching frequencies 4-5 times higher than previous design iterations. Simulation models of the topology have been developed and a breadboard of the topology has been fabricated and evaluated on both resistive loads and an integrated Hall thruster test. This paper will present collected performance and integrated test data from both the fabricated auxiliary and resonant discharge power supplies. Advantages of the resonant converter architecture over more traditional pulse-width modulated (PWM) techniques in Hall-effect thruster discharge power supply applications will also be described.

electric propulsion↗
NASA NTRSDate not supplied

The Effect of Impingement Angle on Bipropellant Thruster Plume Degradation of Spacecraft Materials

Unburned and partially combusted liquid propellant present in bipropellant thruster plumes can contaminate and damage external spacecraft materials. Microscopic impact features due to high-speed droplet impacts have been observed on space shuttle flight experiments and International Space Station (ISS) returned flight hardware. Analytical results have shown that particle impingement angle greatly affects surface damage, with normal impacts being the most severe and highly oblique impacts being more benign. The effect of thruster plume impingement angle on material degradation has now been further studied following a bipropellant thruster plume test performed at the high-vacuum plume test facility for chemical thrusters at DLR Göttingen, STG-CT. Several spacecraft material samples were exposed to the freely expanding pulsed thruster plume at a range of impingement angles (from 0 to 75 deg). Results of post-test evaluations and potential spaceflight applications are summarized.

Gateway↗
DOE OSTI2021

Quasi-steady testing approach for high-power Hall thrusters

Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have been identified as enabling technologies for applications such as lunar tugs, large satellite orbital transfer vehicles, and solar system exploration. These large thrusters introduce significant testing challenges due to the propellant flow rate exceeding the pumping speed available in most laboratories. Even with proposed upgrades in mind, the likelihood that multiple vacuum facilities will exist in the near future to allow long duration testing of high-power Hall thrusters operating at power levels in excess of 100 kW remains extremely low. In this article, we numerically explore the feasibility of testing Hall thrusters in a quasi-steady mode defined by pulsing the mass flow rate between a nominal and a low value. Here, our simulations indicate that sub-second durations available before the chamber reaches critical pressure are sufficiently long to achieve the steady-state current and flow field distributions, allowing us to characterize thruster performance and the near plume region.

42 ENGINEERING↗
DOE OSTI2021

Mitigation of breathing oscillations and focusing of the plume in a segmented electrode wall-less Hall thruster

We report in the absence of the channel walls bounding the plasma, a wall-less Hall thruster is a promising configuration with a potentially longer lifetime and easier scalability than conventional Hall thrusters. Because the ion acceleration takes place in the fringing magnetic field with a strong axial component, the operation of a typical wall-less thruster is characterized by a large beam divergence of the plasma flow, which reduces the thrust. In this work, the addition of a biased segmented electrode to the wall-less thruster is shown to significantly narrow the plasma plume and suppress large amplitude breathing oscillations of the discharge current commonly associated with ionization instability. Both effects result in improvements to the thruster performance. Physical mechanisms responsible for these effects are unclear, but they are apparently associated with the reduction of the electron cross field transport to the anode and a transition in the breathing mode frequency.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
NASA NTRS1971

Performance of a neutralizer for electron bombardment thruster

Results of the SERT II flight indicate that the hollow cathode neutralizer not only represents a power and propellant weight penalty but can be a contributing cause to accelerator grid erosion. Tests with a 30-cm diameter thruster show that a neutralizer position of approximately 9 cm axially downstream of the accelerator grid and approximately 9 cm radially away from the outer edge of the accelerator grid and pointing parallel to the thruster axis provides the best overall performance. The estimated grid wear rate was less than 0.08 mm in 10,000 hours. The coupling voltage was approximately 17 volts at a neutralizer flow rate of 22 equivalent milliamperes of mercury and a beam current of 1.5 amperes. Neutralizer power was 31 watts and the effect of neutralizer flow on overall propellant utilization efficiency is a 1.2 percentage point reduction at a thruster utilization efficiency of 90 percent. The neutralizer position defined in tests with a 30 cm thruster was tested with a 15 cm SERT II thruster. When the neutralizer was relocated further downstream with this different orientation, accelerator impingement current due to neutralizer operation was reduced by approximately a factor of seven and was nearly independent of neutralizer operation.

Bechtel, R. T.↗
NASA NTRS1973

Review of Kaufman thruster development at the Lewis Research Center, 1973

Two thruster sizes are studied. One, a small 5-cm or 8-cm size is for spacecraft station keeping. The other, 30-cm (130 mN thrust), is for a thruster array to do primary solar electric propulsion. A 5-cm thruster (1.8 mN) has recently completed 9715 hr of life testing. Use of dished grids in the 30-cm thruster has increased beam current from 2 to 5 A. The total thrust system mass is compared for present small thrusters at different operating conditions for station keeping of synchronous satellites.

Kerslake, W. R.↗