Hani Kamhawi

NEXT Ion Engine 2000 Hour Wear Test Results

The results of the NEXT 2000 h wear test are presented. This test was conducted with a 40 cm engineering model ion engine, designated EM1, at a 3.52 A beam current and 1800 V beam power supply voltage. Performance tests, which were conducted over a throttling range of 1.1 to 6.9 kW throughout the wear test, demonstrated that EM1 satisfied all thruster performance requirements. The ion engine accumulated 2038 h of operation at a thruster input power of 6.9 kW, processing 43 kg of xenon. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The ion engine was also inspected following the test. This paper presents these findings.

The High Power Electric Propulsion (HiPEP) Ion Thruster

John E. Foster, Tom Haag, and Michael PattersonGlenn Research Center, Cleveland, OhioGeorge J. Williams, Jr.Ohio Aerospace Institute, Brook Park, OhioJames S. SoveyAlpha-Port, Inc., Cleveland, OhioChristian CarpenterQSS Group, Inc., Cleveland, OhioHani Kamhawi, Shane Malone, and Fred ElliotGlenn Research Center, Cleveland, Ohio

Performance Test Results of the NASA-457M v2 Hall Thruster

Abstract Performance testing of a second generation, 50 kW-class Hall thruster labeled NASA-457M v2 was conducted at the NASA Glenn Research Center. This NASA-designed thruster is an excellent candidate for a solar electric propulsion system that supports human exploration missions. Thruster discharge power was varied from 5 to 50 kW over discharge voltage and current ranges of 200 to 500 V and 15 to 100 A, respectively. Anode efficiencies varied from 0.56 to 0.71. The peak efficiency was similar to that of other state-of-the-art high power Hall thrusters, but outperformed these thrusters at lower discharge voltages. The 0.05 to 0.18 higher anode efficiencies of this thruster compared to its predecessor were primarily due to which of two stable discharge modes the thruster was operated. One stable mode was at low magnetic field strengths, which produced high anode efficiencies, and the other at high magnetic fields where its predecessor was operated. Cathode keeper voltages were always within 2.1 to 6.2 V and cathode voltages were within 13 V of tank ground during high anode efficiency operation. However, during operation at high magnetic fields, cathode-to-ground voltage magnitudes increased dramatically, exceeding 30 V, due to the high axial magnetic field strengths in the immediate vicinity of the centrally-mounted cathode. The peak thrust was 2.3 N and this occurred at a total thruster input power of 50.0 kW at a 500 V discharge voltage. The thruster demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 46.1 mN/kW at full power, and a specific impulse range of 1420 to 2740 s. For a discharge voltage of 300 V, where specific impulses would be about 2000 s, thrust efficiencies varied from 0.57 to 0.63.

Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

NASA is developing mission concepts for a solar electric propulsion technology demonstration mission. A number of mission concepts are being evaluated including ambitious missions to near Earth objects. The demonstration of a high-power solar electric propulsion capability is one of the objectives of the candidate missions under consideration. In support of NASAs exploration goals, a number of projects are developing extensible technologies to support NASAs near and long term mission needs. Specifically, the Space Technology Mission Directorate Solar Electric Propulsion Technology Demonstration Mission project is funding the development of a 12.5-kW magnetically shielded Hall thruster system to support future NASA missions. This paper presents the design attributes of the thruster that was collaboratively developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory. The paper provides an overview of the magnetic, plasma, thermal, and structural modeling activities that were carried out in support of the thruster design. The paper also summarizes the results of the functional tests that have been carried out to date. The planned thruster performance, plasma diagnostics (internal and in the plume), thermal, wear, and mechanical tests are outlined.

Current-mode triple and quadruple Langmuir probe methods with applications to flowing pulsed plasmas

A current-mode method for triple and quadruple Langmuir probes was developed and implemented in flowing, pulsed, collisionless plasmas. The current-mode method involves biasing all probe electrodes, and requires the measurement of probe currents providing the electron temperature, the electron density, and the ratio of ion speed to most probable thermal speed. The current-mode theory is developed for a single species, two-temperature, collisionless plasma. The current collection model for a probe aligned with the flow and radius to Debye length ratios of 5/spl les/r/sub p///spl lambda//sub D//spl les/100 accounts for finite-sheath effects while for r/sub p///spl lambda//sub D/>100, current collection is based on the thin-sheath assumption. The ion current to the perpendicular probe assumes a thin-sheath and is given as a function of the ion speed ratio. The numerical procedure for the solution of the nonlinear current-mode equations, as well sensitivity and uncertainty analysis are presented. The plasma source used in the experiments is a laboratory Teflon pulsed plasma thruster, operating at discharge energies of 5, 20, and 40 J, with a pulse duration of 10-15 /spl mu/s, ablating 20-50 /spl mu/g/pulse. Current-mode triple and quadruple probe measurements obtained at various locations in the plume of the plasma source are presented. Extensive comparisons between double probe and current-mode probe measurements validate the new method.

In-Space Propulsion High Voltage Hall Accelerator Development Project Overview

NASA’s Science Mission Directorate In-Space Propulsion Technology Project is funding the development of a high specific impulse long life Hall thruster. The goal of the high voltage Hall accelerator (HiVHAc) project is to develop a flight-like, engineering model (EM) Hall thruster that can meet future NASA science mission requirements. These requirements are met by a thruster that operates over an input power range from 0.3 to 3.5 kW, attains specific impulses from 1,000 to 2,700 seconds, and processes at least 300 kg of xenon propellant at full power. To demonstrate the HiVHAc project goal, two laboratory thrusters have been built and tested. The latest laboratory thruster, the NASA-103M.XL, incorporated a life-extending discharge channel replacement innovation and has been operated for approximately 5,000 hours at a discharge voltage of 700 volts. In 2007, NASA Glenn Research Center teamed with Aerojet to design and manufacture a flight-like HiVHAc EM thruster which incorporated this life-extending channel replacement innovation. The EM thruster was designed to withstand the structural and thermal loads encountered during NASA science missions and to attain performance and lifetime levels consistent with NASA missions. Aerojet and NASA Glenn Research Center have completed the EM thruster design, structural and thermal analysis, fabrication of thruster components, and have assembled and extensively tested one EMl thruster. Performance and thermal characterization of the engineering model thruster has been performed for discharge power levels up to 3.5 kW. The results indicate discharge efficiencies up to of 63% and discharge specific impulse up to 2,930 seconds. In addition to the thruster development, the HiVHAc project is leveraging power processing unit and xenon flow system developments sponsored by other projects but that can apply directly to a HiVHAc system. The goal is to advance the technology readiness level of a HiVHAc propulsion system to 6.

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

Abstract : During a 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 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 will focus 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.

NASA HERMeS Hall Thruster Electrical Configuration Characterization

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in-space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This paper describes the electrical configuration testing of the HERMeS TDU-1 Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined were 1) thruster body tied to facility ground, 2) thruster floating, and 3) thruster body electrically tied to cathode common. The HERMeS TDU-1 Hall thruster was also configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Advanced Xenon Feed System (AXFS) Development and Hot-fire Testing

NASA’s In-Space Propulsion Technology project has been investing in advanced xenon feed system technologies to significantly reduce the cost, mass, and volume while increasing the reliability over the state-of-the-art alternatives. VACCO industries was competitively selected to develop an AXFS under a NASA Research Announcement solicitation and completed their effort with the delivery of an AXFS and hot-fire demonstration. The AXFS development produced two flow control modules, one pressure control module, the AXFS controller and LabVIEW software. The baseline AXFS design is more than a 90 percent reduction in mass, cost, and volume over the Dawn flight system and over 50 percent reduction over comparable TRL 6 flow control systems. The component modules completed environmental testing and integrated hot-fire testing in March of 2009. An overview of the development effort and results of testing are presented.

Performance, Facility Pressure Effects, and Stability Characterization Tests of NASA's Hall Effect Rocket with Magnetic Shielding Thruster

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. Part of the technology maturation effort included experimental evaluation of the TDU-1 thruster with conducting and dielectric front pole cover materials in two different electrical configurations. A graphite front magnetic pole cover thruster configuration with the thruster body electrically tied to cathode, and an alumina front pole cover thruster configuration with the thruster body floating were evaluated. Both configurations were also evaluated at different facility background pressure conditions to evaluate background pressure effects on thruster operation. Performance characterization tests found that higher thruster performance was attained with the graphite front pole cover configuration with the thruster electrically tied to cathode. A total thrust efficiency of 68% and a total specific impulse of 2,820 s was demonstrated at a discharge voltage of 600 V and a discharge power of 12.5 kW. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations and with maps of the discharge current-voltage-magnetic field (IVB). Analysis of TDU-1 discharge current waveforms found that lower normalized discharge current peak-to-peak and root mean square magnitudes were attained when the thruster was electrically floated with alumina front pole covers. Background pressure effects characterization tests indicated that the thruster performance and stability were mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 110-5 Torr-Xe (for thruster flow rates of 20.5 mg/s). Power spectral density analysis of the discharge current waveforms showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant breathing mode frequency. Finally, IVB maps of the TDU-1 thruster indicated that the discharge current became more oscillatory with higher discharge current peak-to-peak and RMS values with increased facility background pressure at lower thruster mass flow rates; thruster operation at higher flow rates resulted in less change to the thrusters IVB characteristics with elevated background pressure.