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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
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↗
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.↗
NASA NTRS1973

Performance characterization tests of three 0.44-N (0.1 lbf) hydrazine catalytic thrusters

The 0.44-N (0.1-lbf) class of hydrazine catalytic thruster has been evaluated to assess its capability for spacecraft limit-cycle attitude control with thruster pulse durations on the order of 10 milliseconds. Dynamic-environment and limit-cycle simulation tests were performed on three commercially available thruster/valve assemblies, purchased from three different manufacturers. The results indicate that this class of thruster can sustain a launch environment and, when properly temperature-conditioned, can perform limit-cycle operations over the anticipated life span of a multi-year mission. The minimum operating temperature for very short pulse durations was determined for each thruster. Pulsing life tests were then conducted on each thruster under a thermally controlled condition which maintained the catalyst bed at both a nominal 93 C (200 F) and 205 C (400 F). These were the temperatures believed to be slightly below and very near the minimum recommended operating temperature, respectively. The ensuing life tests ranged from 100,000 to 250,000 pulses at these temperatures, as would be required for spacecraft limit-cycle attitude control applications.

Moynihan, P. I.↗
NASA NTRS1972

Electrostatic thrusters.

The current status of research and development programs on electrostatic thrusters is reviewed. Current programs that utilize mercury electron-bombardment thrusters range from 5- to 30-cm in diameter. Recent progress on the 5-cm thruster has emphasized durability, with accelerator time exceeding 6300 hours and total time on the rest of the thruster exceeding 8300 hours. Recent progress on the 30-cm thruster has been outstanding in dished-grid accelerator systems. Ion beams up to 5 amperes have been obtained for short periods with 1000 volts net accelerating potential difference. The cesium electron-bombardment and cesium contact programs are also described.

Kaufman, H. R.↗
NASA NTRS1973

Test facility for 6000 hour life test of 30 cm mercury ion thruster

The environmental and instrumentation requirements for long term testing of electrical propulsion thrusters which impose severe and unusual requirements upon the simulation facility were studied. High speed ions ejected from a mercury thruster erode material from collecting surfaces, which is then scattered and redeposited upon other surfaces, with resultant damage to the chamber and test article. By collecting the thruster plume on a frozen mercury surface damage to the thruster and chamber by back-scattered erosion products was minimized. Provisions for unattended operation, remote data acquisition, personnel safety, and instrumentation for assessing thruster performance are also discussed.

Caldwell, J. J.↗
NASA NTRS1974

Study of monopropellants for electrothermal thrusters: Design and fabrication task summary report

The feasibility of operating small thrust level electrothermal thrusters with monopropellants other than MIL-grade hydrazine was studied. Analytical study, design, and fabrication of demonstration thrusters was performed, and an evaluation test program was initiated to evaluate monopropellants with freezing points lower than MIL-grade hydrazine, and to determine their applicability to electrothermal thrusters for spacecraft attitude control. Five demonstration thrusters were fabricated to determine the feasibility of operation with monomethylhydrazine, Aerozine-50, 77 percent hydrazine-23 percent hydrazine azide, and a mixture of hydrazine monopropellants consisting of 35 percent hydrazine-50 percent monomethylhydrazine-15 percent ammonia. The present thruster is designed to produce a steady-state thrust level of 0.344 N at 1.724 x 1 million N/M sq feed pressure. Vacuum specific impulse goals were set at 1961 N-s/kg steady-state and 1716 N-s/kg pulsed-mode.

Kuenzly, J. D.↗
NASA NTRS1974

Development of a short pulsed solid propellant plasma thruster

The experimental results are summarized that were obtained in the development of a Teflon solid propellant pulsed plasma thruster. The feasibility was established of storing and feeding solid propellant in the form of an open circular loop into an operational thruster. This technique was verified to be practical by feeding over 20 inches of Teflon into a micro-thruster over an accumulated life test of 1858 hours. High energy density capacitors were evaluated under vacuum conditions when the capacitor is coupled directly to a plasma thruster. Numerous early capacitor failures were encountered. It was concluded that essentially all of the failures encountered in a vacuum environment are due to an internal electrical breakdown that will occur inside a capacitor that is not truly hermetically sealed. A steady input power significantly in excess of 130 watts can safely be tolerated if heat conduction can be provided to a sink whose temperature is about 16 C. A vacuum life test of the capacitor bank was carried out while discharging into a milli-lb. (milli-Newton) type pulsed plasma thruster. More than 1500 hours of vacuum testing of this milli-Newton type system has been accumulated without any capacitor problems. Recommendations are made for future capacitor designs.

Guman, W. J.↗
NASA NTRS1974

Status of 30 cm mercury ion thruster development

Two engineering model 30-cm ion thrusters were assembled, calibrated, and qualification tested. This paper discusses the thruster design, performance, and power system. Test results include documentation of thrust losses due to doubly charged mercury ions and beam divergence by both direct thrust measurements and beam probes. Diagnostic vibration tests have led to improved designs of the thruster backplate structure, feed system, and harness. Thruster durability is being demonstrated over a thrust range of 97 to 113 mN at a specific impulse of about 2900 seconds. As of August 15, 1974, the thruster has successfully operated for over 4000 hours.

Sovey, J. S.↗
NASA NTRS1973

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

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.

Dahlgren, J. B.↗
NASA NTRS1975

SERT 2 thruster space restart, 1974

The results of testing the flight thrusters on the SERT spacecraft during the 1974 test period are presented. The most notable result was the clearing of the high voltage short from thruster 2 and the successful stable operation of its ion beam. Test periods were limited to 70 minutes or less by earth eclipse of the spacecraft solar array and by ground station coverage limitations. Thruster 2 was restarted 26 times with an ion beam produced 21 times. The high voltage short remains in thruster 1, but the cathodes were restarted 12 times to demonstrate continued restart capability. The propellant feed systems, power processors, and spacecraft ancillary equipment were demonstrated to be functional after 4 1/2 years in space. In addition to the thruster tests, a neutralizer cathode was operated separately to demonstrate that the potential level of a spacecraft could be controlled by the neutralizer alone.

Kerslake, W. R.↗
NASA NTRS1975

Auxiliary propulsion thruster performance with ion machined accelerator grids

A substantial improvement in the performance of an 8-centimeter-diameter auxiliary propulsion thruster was achieved by reducing the diameter of the accelerator grid apertures. The accelerator grid hole geometry was defined by ion machining accelerator grids on an 8-centimeter thruster at thrust levels of 2.2, 4.4, and 6.7 millinewtons was (mN). A thruster with an ion machined accelerator grid was operated at a thrust of 4.4 mN for 1000 hours. The discharge propellant utilization was 92% at an eV/ion of 338. Thruster performance and accelerator grid hole geometry was documented as a function of thrust level. It was also determined that the small hole accelerator grid has a very low backstreaming voltage limit. In fact the thruster can be operated with the acclerator grid held at neutralizer tip potential.

Hudson, W. R.↗
NASA NTRS1975

Measurement of sputtered efflux from 5-, 8-, and 30-cm diameter mercury ion thrusters

A study was undertaken to investigate the sputtered efflux from 5-, 8-, and 30-cm diameter mercury ion thrusters. Quartz crystal microbalances and fused silica samples were used to analyze the sputtered flux. Spectral transmittance measurements and spectrographic analysis of the samples were made after they were exposed to different thruster effluence by operating the thrusters at various conditions and durations of time. These measurements were used to locate the source of the efflux and determine its accumulated effect at various locations near the thruster. Comparisons of in situ and ex situ transmittance measurements of samples exposed to thruster efflux are also presented.

Weigand, A. J.↗
NASA NTRS1974

Status of 30 cm mercury ion thruster development

Two engineering model 30-cm ion thrusters have been assembled, calibrated, and qualification tested. This paper discusses the thruster design, performance, and power system. Test results include documentation of thrust losses due to doubly charged mercury ions and beam divergence by both direct thrust measurements and beam probes. Diagnostic vibration tests have led to improved designs of the thruster backplate structure, feed system, and harness. Thruster durability is being demonstrated over a thrust range of 97 to 113 mN at a specific impulse of about 2900 seconds. As of August 15, 1974, the thruster has successfully operated for over 4000 hours.

Sovey, J. S.↗