Thomas W. Haag

NEXT: NASA's Evolutionary Xenon Thruster

NASA’s Glenn Research Center has been selected to lead development of NASA’s Evolutionary Xenon Thruster (NEXT) system. The central feature of the NEXT system is an electric propulsion thruster (EPT) that inherits the knowledge gained through the NSTAR thruster that successfully propelled the Deep Space 1 to asteroid Braille and comet Borrelly, while significantly increasing the thruster power level and making improvements in performance parameters associated with NSTAR. The EPT concept under development has a 40 cm beam diameter, twice the effective area of the Deep-Space 1 thruster, while maintaining a relatively-small volume. It incorporates mechanical features and operating conditions to maximize the design heritage established by the flight NSTAR 30 cm engine, while incorporating new technology where warranted to extend the power and throughput capability.

Development of an Ion Thruster and Power Processor for New Millennium's Deep Space 1 Mission

James S. Sovey, John A. Hamley, Thomas W. Haag, Michael J. Patterson, Eric J. Pencil,Todd T. Peterson, Luis R. Pinero, John L. Power, Vincent K. Rawlin, and Charles J. SarmientoNASA Lewis Research Center, Cleveland, OhioJohn R. Anderson, Raymond A. Becker, John R. Brophy, and James E. PolkJet Propulsion Laboratory, Pasadena, CaliforniaGerald Benson, Thomas A. Bond, G. I. Cardwell, Jon A. Christensen, Kenneth J. Freick,David J. Hamel, Stephen L. Hart, John McDowell, Kirk A. Norenberg, T. Keith Phelps,Ezequiel Solis, and Harold YostHughes Electron Dynamics Division, Torrance, CaliforniaMichael MatrangaSpectrum Astro Incorporated, Gilbert, ArizonaPrepared for the33rd Joint Propulsion Conference and Exhibitcosponsored by AIAA, ASME, SAE, and ASEESeattle, Washington, July 6-9, 1997National Aeronautics andSpace AdministrationLewis Research Center

Next: NASA's Evolutionary Xenon Thruster development status

NASA’s Glenn Research Center (GRC) is leading the development of NASA’s Evolutionary Xenon Thruster (NEXT) system. The central feature of the NEXT system is an electric propulsion thruster that inherits the knowledge gained through the NSTAR thruster that successfully propelled the Deep Space 1 to asteroid Braille and comet Borrelly, while significantly increasing the thruster power level and making improvements in performance parameters associated with NSTAR. The thruster under development has a 40 cm beam diameter, twice the effective area of the Deep-Space 1 thruster. It incorporates mechanical features and operating conditions to maximize the design heritage established by the flight NSTAR 30 cm thruster, while incorporating new technology where warranted to extend the power and throughput capability. To date three engineering-model thrusters have been manufactured at NASA GRC, and performance, wear, vibration, and integration testing of these thrusters is underway.

NASA 30 Cm Ion Thruster Development Status

A 30 cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for missions of national interest and it is an element of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) program established to validate ion propulsion for space flight applications. The thruster has been developed to an engineering model level and it incorporates innovations in design, materials, and fabrication techniques compared to those employed to conventional ion thrusters. The performance of both functional and engineering model thrusters has been assessed including thrust stand measurements, over an input power range of 0.5-2.3 kW. Attributes of the engineering model thruster include an overall mass of 6.4 kg, and an efficiency of 65 percent and thrust of 93 mN at 2.3 kW input power. This paper discusses the design, performance, and lifetime expectations of the functional and engineering model thrusters under development at NASA.

Operating characteristics of the Russian D-55 thruster with anode layer

Performance measurements of a Russian engineering-model Thruster with Anode Layer (TAL) were obtained as part of a program to evaluate the operating characteristics of Russian Hall-thruster technology. The TAL model D-55 was designed to operate in the 1-2 kW power range on xenon. When received, the thruster had undergone only a few hours of acceptance testing by the manufacturer. Direct thrust measurements were obtained at a background pressure of 0.0003 Pa (2 x 10(exp -6) torr) at power levels ranging from 0.3 kW to 2.1 kW. At the nominal power level of 1.3 kW, a specific impulse level of 1600 s with a corresponding efficiency of 0.48 was attained. At all flow rates tested, the efficiency increased linearly with specific impulse until a maximum was reached, and then the efficiency leveled off. Increasing the anode flow rate shifted the efficiency upward, reaching 0.50 at 1850 s specific impulse. The thruster was equipped with inner and outer electromagnets which were isolated from the discharge and from each other. Variation of the magnetic field, obtained by changing the currents through the magnets, had little effect on performance, except at current levels below 70 percent of nominal. For a given operating condition, the performance was slightly affected by facility pressure. As the pressure was increased by a factor of thirty to 0.008 Pa (6 x 10(exp -5) torr), the current steadily increased by 4 percent, and the thrust increased by 2 percent. Performance comparisons were made with the Stationary Plasma Thruster, and the efficiency and specific impulse values were similar at power levels ranging from 0.9 kW to 1.5 kW. Endurance testing was not performed, and comparisons of lifetime were not made.

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.

Ion propulsion development activities at the NASA Glenn Research Center

The NASA Glenn Research Center (GRC) ion propulsion program addresses the need for high specific impulse ion propulsion systems and technology across a broad range of mission applications and power levels. Development areas include high-throughput NSTAR derivative engine and power processing technology, lightweight high-efficiency sub-kilowatt ion propulsion, micro-ion propulsion concepts, engine and component technologies for highpower (30 kW class) ion engines, and fundamentals. NASA GRC is also involved in two highly focussed activities: development of 5/10-kW class next-generation ion propulsion system technology, and development of high-specific impulse (> 10,000 seconds) ion propulsion technology applicable to deep-space and interstellar-precursor missions.

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.

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.

Results of the 2000 hr Wear Test of the HiPEP Ion Thruster with Pyrolytic Graphite Ion Optics

A 25 kW, long-life ion thruster was developed and wear tested at the NASA Glenn Research Center in support of Project Prometheus. The 2000 hr wear test was undertaken to quantify known erosion phenomena such as ion optics erosion due to charge exchange ion impingement and discharge cathode keeper erosion and to identify unknown wear mechanisms associated with such high-specific impulse, high-power thrusters. The discussion provides a comparison between predicted wear and deposition rates and an analysis of the impact of the various phenomena observed. Trends in observed erosion of the ion optics were consistent with expectations and the negligible wear of the discharge keeper and neutralizer keeper was less than expected. The HiPEP thruster was designed and developed at the NASA Glenn Research Center (GRC) during the ongoing development of the NASA Evolutionary Xenon Thruster (NEXT) 2 and following the successful demonstration of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) ion thruster on the Deep Space 1 spacecraft. 3 Both of these programs incorporated relatively short duration wear tests (~2000 hr) at GRC early in their development efforts. 4,5 These tests were conducted to validate design approaches, identify unknown wear mechanisms, and quantify wear rates before the thrusters were developed to higher fidelity. It is in this context that the HiPEP thruster was incorporated into a wear test very early in its development. While this test was ongoing, a higher fidelity development model thruster (DMT) was being developed in collaboration between GRC and the Aerojet Corporation. An ion optics assembly electrostatically identical to the one being wear tested was successfully vibration tested. 6 A thermo -mechanical model of the DMT was also developed. The results of the 2000 hr wear test would then support and augment a detailed design process and significantly accelerate the delivery of high -fidelity hardware. However, before the assembly of the HiPEP DMT, NASAs Project Prometheus redirected the HiPEP design work to support the design of the Herakles ion thruster, which was in part based on advances made under the HiPEP program. The HiPEP thruster was designed to accommodate a large range of operational requirements and to facilitate the future development of higher -power ion thrusters. To this end, it has rectangular discharge chamber and incorporates pyrolytic graphite (PG) ion optics. HiPEP versions have been successfully operated with both dc (i.e., hollow cathode-based) and microwave discharges. 7,8 Operation at 25 kW over a specific impulse range of 6000 to 9000 s using a dc discharge was demonstrated. 8 Following the performance demonstration of the HiPEP thruster with PG ion optics and a dc discharge, 9 the thruster entered a 2000 hr wear test. The objectives of the test were to