Stationary Plasma Thruster Plume Emissions

The emission spectrum from a xenon plasma produced by a Stationary Plasma Thruster provided by the Ballistic Missile Defense Organization (BMDO) was measured. Approximately 270 individual Xe I, Xe II, and XE III transitions were identified. A total of 250 mW of radiated optical emission was estimated from measurements taken at the thruster exit plane. There was no evidence of erosion products in the emission signature. Ingestion and ionization of background gas at elevated background pressure was detected. The distribution of excited states could be described by temperatures ranging from fractions of 1 eV to 4 eV with a high degree of uncertainty due to the nonequilibrium nature of this plasma. The plasma was over 95 percent ionized at the thruster exit plane. Between 10 and 20 percent of the ions were doubly charged. Two modes of operation were identified. The intensity of plasma emission increased by a factor of two during operation in an oscillatory mode. The transfer between the two modes of operation was likely related to unidentified phenomena occurring on a time scale of minutes

Simplified Numerical Description of SPT Operations

A simplified numerical model of the plasma discharge within the SPT-100 stationary plasma thruster was developed to aid in understanding thruster operation. A one dimensional description was used. Non-axial velocities were neglected except for the azimuthal electron velocity. A nominal operating condition of 4.5 mg/s of xenon anode flow was considered with 4.5 Amperes of discharge current, and a peak radial magnetic field strength of 130 Gauss. For these conditions, the calculated results indicated ionization fractions of 0.99 near the thruster exit with a potential drop across the discharge of approximately 250 Volts. Peak calculated electron temperatures were found to be sensitive to the choice of total ionization cross section for ionization of atomic xenon by electron bombardment and ranged from 51 eV to 60 eV. The calculated ionization fraction, potential drop, and electron number density agree favorably with previous experiments. Calculated electron temperatures are higher than previously measured

Stationary Plasma Thruster Ion Velocity Distribution

A nonintrusive velocity diagnostic based on laser induced fluorescence of the 5d4F(5/2)-6p4D(5/2) singly ionized xenon transition was used to interrogate the exhaust of a 1.5 kW Stationary Plasma Thruster (SPT). A detailed map of plume velocity vectors was obtained using a simplified, cost-effective, nonintrusive, semiconductor laser based scheme. Circumferential velocities on the order of 250 m/s were measured which implied induced momentum torques of approximately 5 x 10(exp -2) N-cm. Axial and radial velocities were evaluated one mm downstream of the cathode at several locations across the width of the annular acceleration channel. Radial velocities varied linearly with radial distance. A maximum radial velocity of 7500 m/s was measured 8 mm from the center of the channel. Axial velocities as large as 16,500 m/s were measured

Performance characterization of a segmented anode arcjet thruster

A modular, 1 to 2 kW class arcjet thruster incorporating a segmented anode/nozzle was operated on a thrust stand to obtain performance characteristics of the device and further study its operating characteristics under a number of experimental conditions. The nozzle was composed of five axial conducting segments isolated from one another by boron nitride spacers. The electrical configuration allowed the current delivered to the arcjet to be collected at any combination of segments. Both the current collected by each segment, and the potential difference between the cathode and each segment were monitored throughout the test period. As in previous tests a similar device, current appeared to attach diffusely in the anode when all of the segments were allowed to conduct. Improvements to the device allowed long term (4 to 8 hour) operation at steady-state and operating characteristics were repeatable over extended periods. Performance characteristics indicated that the segmented anode reasonably simulates the behavior of solid anodes of similar geometry. Current distribution depended on flow rate as the arc attachment moved downstream in the nozzle with increases in the mass flow rate. The current level had little effect on current distribution on the anode segments. Thrust measurements indicated that the current distribution in the nozzle did not significantly affect performance of the device

Stationary Plasma Thruster Plume Characteristics

Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions. This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements. Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties. For the design operating voltage of 300 V the centerline electron density was found to decrease from approximately 1.8 x 10 exp 17 cubic meters at a distance of 0.3 m to 1.8 X 10 exp 14 cubic meters at a distance of 4 m from the thruster. The electron temperature over the same region was between 1.7 and 3.5 eV. Ion current density measurements showed that the plume was sharply peaked, dropping by a factor of 2.6 within 22 degrees of centerline. The ion energy 4 m from the thruster and 15 degrees off-centerline was approximately 270 V. The thruster cathode flow rate and facility pressure were found to strongly affect the plume properties. In addition to the plume measurements, the data from the various probe types were used to assess the impact of probe design criteri

Hall thruster ion beam characterization

The distribution of ion current density as a function of angle from the thruster axis was measured at nominal operating conditions of 300 Volts discharge voltage and 4.5 Amperes discharge current for two SPT-100 stationary plasma thrusters from Fakel Enterprises, a D-55 anode layer thruster from the Central Scientific Research Institute for Machine Building (TsNIIMASH), and a first, second, and third generation T-100 stationary plasma thruster from the Scientific Research Institute of Thermal Process (NIITP). The data showed that the current density distributions of these thrusters were similar. Some differences in peak ion current density were observed. Multiply charged ions were found to be a small fraction of the plasma plume for all of the thrusters. The effect of facility pressure on ion current density distribution was found to be nonnegligible at pressures above 2 x 10(exp -6) Torr. The ion current density distributions of a new SPT-100 and a 6000 hour wear tested SPT-100 exhibited no discernible difference. Ion current density measurements were also taken at off-nominal thruster operating conditions

Magnetic circuit for hall effect plasma accelerator

A Hall effect plasma accelerator includes inner and outer electromagnets, circumferentially surrounding the inner electromagnet along a thruster centerline axis and separated therefrom, inner and outer magnetic conductors, in physical connection with their respective inner and outer electromagnets, with the inner magnetic conductor having a mostly circular shape and the outer magnetic conductor having a mostly annular shape, a discharge chamber, located between the inner and outer magnetic conductors, a magnetically conducting back plate, in magnetic contact with the inner and outer magnetic conductors, and a combined anode electrode/gaseous propellant distributor, located at a bottom portion of the discharge chamber. The inner and outer electromagnets, the inner and outer magnetic conductors and the magnetically conducting back plate form a magnetic circuit that produces a magnetic field that is largely axial and radially symmetric with respect to the thruster centerline

High-Power Krypton Hall Thruster Technology Being Developed for Nuclear-Powered Applications

The NASA Glenn Research Center has been performing research and development of moderate specific impulse, xenon-fueled, high-power Hall thrusters for potential solar electric propulsion applications. These applications include Mars missions, reusable tugs for low-Earth-orbit to geosynchronous-Earth-orbit transportation, and missions that require transportation to libration points. This research and development effort resulted in the design and fabrication of the NASA-457M Hall thruster that has been tested at input powers up to 95 kW. During project year 2003, NASA established Project Prometheus to develop technology in the areas of nuclear power and propulsion, which are enabling for deep-space science missions. One of the Project-Prometheus-sponsored Nuclear Propulsion Research tasks is to investigate alternate propellants for high-power Hall thruster electric propulsion. The motivation for alternate propellants includes the disadvantageous cost and availability of xenon propellant for extremely large scale, xenon-fueled propulsion systems and the potential system performance benefits of using alternate propellants. The alternate propellant krypton was investigated because of its low cost relative to xenon. Krypton propellant also has potential performance benefits for deep-space missions because the theoretical specific impulse for a given voltage is 20 percent higher than for xenon because of krypton's lower molecular weight. During project year 2003, the performance of the high-power NASA-457M Hall thruster was measured using krypton as the propellant at power levels ranging from 6.4 to 72.5 kW. The thrust produced ranged from 0.3 to 2.5 N at a discharge specific impulse up to 4500 sec

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

High-Power Hall Propulsion Development at NASA Glenn Research Center

The NASA Office of the Chief Technologist Game Changing Division is sponsoring the development and testing of enabling technologies to achieve efficient and reliable human space exploration. High-power solar electric propulsion has been proposed by NASA's Human Exploration Framework Team as an option to achieve these ambitious missions to near Earth objects. NASA Glenn Research Center is leading the development of mission concepts for a solar electric propulsion Technical Demonstration Mission. The mission concepts are highlighted in this paper but are detailed in a companion paper. There are also multiple projects that are developing technologies to support a demonstration mission and are also extensible to NASA's goals of human space exploration. Specifically, the In-Space Propulsion technology development project at the NASA Glenn has a number of tasks related to high-power Hall thrusters including performance evaluation of existing Hall thrusters; performing detailed internal discharge chamber, near-field, and far-field plasma measurements; performing detailed physics-based modeling with the NASA Jet Propulsion Laboratory's Hall2De code; performing thermal and structural modeling; and developing high-power efficient discharge modules for power processing. This paper summarizes the various technology development tasks and progress made to date