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At least 757 records · Page 42

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NASA NTRS1992

Anode power deposition in a MPD thruster with a magnetically annulled Hall parameter anode

Results from previous studies indicate that the anode fall increases monotonically with the electron Hall parameter. In an attempt to reduce the anode fall by decreasing the local electron Hall parameter, a proof-of-concept test was performed in which an array of 36 permanent magnets were imbedded within the anode of a high power quasi-steady MPD thruster to decrease the local azimuthal component of the induced magnetic field. The modified thruster was operated at power levels between 150 kW and 4 MW with Ar and He propellants. Terminal voltage, triple probe, floating probe, and magnetic probe measurements were made to characterize the performance of the thruster with new anode. Incorporation of the modified anode resulted in a reduction of the anode fall by up to 15 V with Ar and 20 V with He, which corresponded to decreased anode power fractions of 40 and 45 percent with Ar and He, respectively.

Gallimore, Alec D.↗

NASA NTRS1992

Scaling of 100 kW class applied-field MPD thrusters

Three cylindrical applied-field magnetoplasmadynamic thrusters were tested with argon propellant over a broad range of operating conditions to establish empirical scaling laws for thruster performance. Argon flow rates, discharge currents, and applied-field strengths were varied between 0.025 and 0.14 g/s, 750 to 2000 A, and 0.034 to 0.20 T, respectively. The results showed that the thrust reached over five times the self-field value, and that thrust increased linearly with the product of discharge current and applied-field strength, and quadratically with the anode radius. While increasing the propellant flow rate increased the thrust, it did not affect the rate of thrust increase with applied-field strength, and, at low propellant flow rates, the self-field thrust approached 30 percent of the measured thrust. The voltage increased linearly with applied-field strength but was insensitive to the discharge current. The rate of voltage increase with applied-field strength was strongly dependent on anode radius. Thruster efficiency increased monotonically with applied-field strength and propellant flow rate.

Myers, Roger M.↗

NASA NTRS1992

Characterization of ion accelerating systems on NASA LeRC's ion thrusters

An investigation is conducted regarding ion-accelerating systems for two NASA thrusters to study the limits of ion-extraction capability or perveance. A total of nine two-grid ion-accelerating systems are tested with the 30- and 50-cm-diam ring-cusp inert-gas ion thrusters emphasizing the extension of ion-extraction. The vacuum-tank testing is described using xenon, krypton, and argon propellants, and thruster performance is computed with attention given to theoretical design considerations. Reductions in perveance are noted with decreasing accelerator-hole-to-screen-hole diameter ratios. Perveance values vary indirectly with the ratio of discharge voltage to total accelerating voltage, and screen/accelerator electrode hole-pair alignment is also found to contribute to perveance values.

Rawlin, Vincent K.↗

NASA NTRS1992

Ion Thruster Development at NASA Lewis Research Center

Recent ion propulsion technology efforts at NASA's Lewis Research Center including development of kW-class xenon ion thrusters, high power xenon and krypton ion thrusters, and power processors are reviewed. Thruster physical characteristics, performance data, life projections, and power processor component technology are summarized. The ion propulsion technology program is structured to address a broad set of mission applications from satellite stationkeeping and repositioning to primary propulsion using solar or nuclear power systems.

Sovey, James S.↗

NASA NTRS1992

Second Magnetoplasmadynamic Thruster Workshop

The meeting focused on progress made in establishing performance and lifetime expectations of magnetoplasmadynamic (MPD) thrusters as functions of power, propellant, and design; models for the plasma flow and electrode components; viability and transportability of quasi-steady thruster testing; engineering requirements for high power, long life thrusters; and facilities and their requirements for performance and life testing.

Source record↗

NASA NTRS1992

OL-AC Phillips Laboratory MPD thruster research program

The topics are presented in viewgraph form and include the following: facility construction; quadruple langmuir probe measurements; hollow/porous anode magnetoplasmadynamic (MPD) thruster; the measurement of the ionization fraction inside of the MPD thruster; and the experimental investigation of the effects of microturbulence on MPD thruster performance.

Tilley, Dennis L.↗

NASA NTRS1991

Thruster configurations for maneuvering heavy payloads

The Cargo Transfer Vehicle (CTV) will be required to perform six degree of freedom (6 DOF) maneuvers while carrying a wide range of payloads varying from 100,000 lbm to no payload. The current baseline design configuration for the CTV uses a forward propulsion module (FPM) mounted in front of the payload with the CTV behind the payload so that the center of gravity (CG) of the combined stack is centered between the thruster sets. This allows for efficient rotations and translations of heavy payloads in all directions; however, the FPM is a costly item, so it is desirable to find design solutions that do not require the FPM. This presentation provides an overview of the analysis of the FPM requirements for the CTV. In this study, only the reaction control system (RCS) thruster configurations are considered for 6 DOF maneuvers of various CTV cargo configurations. An important output of this study are the viable alternative thruster configurations that eliminate the need for the FPM. Initial results were derived using analytical techniques and simulation analysis tools. Results from the preliminary analysis were validated using our 6 DOF simulation.

Tsugawa, Roy K.↗

NASA NTRS1993

100-kW class applied-field MPD thruster component wear

Component erosion and material deposition sites were identified and analyzed during tests of various configurations of 100 kW class, applied-field, water-cooled magnetoplasmadynamic (MPD) thrusters. Severe erosion of the cathode and the boron nitride insulator was observed for the first series of tests, which was significantly decreased by reducing the levels of propellant contamination. Severe erosion of the copper anode resulting from sputtering by the propellant was also observed. This is the first observation of this phenomenon in MPD thrusters. The anode erosion indicates that development of long life MPD thrusters requires the use of light gas propellants such as hydrogen, deuterium, or lithium.

Mantenieks, Maris A.↗

NASA NTRS1993

Preliminary investigation of power flow and electrode phenomena in a multi-megawatt coaxial plasma thruster

This paper summarizes preliminary experimental and theoretical research that was directed towards the study of quasisteady-state power flow in a large, un-optimized, multi-megawatt coaxial plasma thruster. The report addresses large coaxial thruster operation and includes evaluation and interpretation of the experimental results with a view to the development of efficient, steady-state megawatt-class magnetoplasmadynamic (MPD) thrusters.

Schoenberg, Kurt↗

NASA NTRS1992

The effects of 1 kW class arcjet thruster plumes on spacecraft charging and spacecraft thermal control materials

Arcjet thrusters are soon to be used for north/south stationkeeping on commercial communications satellites. A series of tests was performed to evaluate the possible effects of these thrusters on spacecraft charging and the degradation of thermal control material. During the tests the interaction between arcjet plumes and both charged and uncharged surfaces did not cause any significant material degradation. In addition, firing an arcjet thruster benignly reduced the potential of charged surfaces to near zero.

Bogorad, A.↗

NASA NTRS1993

Performance of a low-power subsonic-arc-attachment arcjet thruster

A subsonic-arc-attachment thruster design was scaled from a 30 kW 1960's vintage thruster to operate at nominally 3 kW. Performance measurements were obtained over a 1-4 kW power range using hydrogen as the propellant. Several modes of operation were identified and were characterized by varying degrees of voltage instability. A stability map was developed showing that the voltage oscillations were brought upon by elevated current or propellant levels. At a given specific energy level the specific impulse increased asymptotically with increased flow rates. Comparisons of performance were made between radial and tangential propellant injection. When the vortex flow was eliminated using radial injection, the operating voltages were lower at a given current, and the specific impulse and efficiency decreased. Tests were also conducted to determine the effects of background pressure on operation. Performance data were obtained at pressures of 0.047 Pa and 18 Pa. For a given specific energy level, the performance increased with a decrease in facility background pressure. Lowering the background pressure also caused a dramatic change in the voltage-current characteristic and the voltage stability, a phenomenon not previously reported with conventional supersonic-arc-attachment thrusters.

Sankovic, John M.↗

NASA NTRS1993

Recent testing of 30 KW hydrogen arcjet thrusters

NASA is conducting efforts to evaluate high-power hydrogen arcjets for orbit transfer propulsion applications. As part of this program, an attempt was made to reexamine both radiatively- and regeneratively-cooled, 30 kW thrusters first demonstrated by the Giannini Scientific Corp. in 1963. The arcjets were configured to force arc attachment upstream of the throat in a subsonic chamber region. While thruster currents were steady, the voltage traces exhibited sawtooth waveforms at frequencies on the order of 20 kHz. Voltage variations per cycle were typically between 100 and 310 volts, indicating major changes in the position of the arc attachment with time. When operated at their respective design points, the performance of both thrusters fell below the values listed in the 1960's development reports. The reason for the discrepancies is not currently understood and further investigations are in progress. However, the recently measured efficiencies were high compared to those obtained with constricted-arc designs at similar conditions, and further arcjet performance optimizations may be possible.

Haag, Thomas W.↗

NASA NTRS1993

Experimental studies of an ECR plasma thruster

The Electron Cyclotron Resonance (ECR) thruster is a proposed electrodeless space electric propulsion device with interesting and little understood physics. A laboratory ECR thruster was run in a vacuum tank at pressures in the 10 exp -5 torr range using 2.12 GHz microwave beam and Ar gas propellant. Movable diagnostic probes (a Faraday cup and a gridded energy analyzer) measured plasma characteristics as propellant gas flow rate and input microwave power level were varied. Ion energy and flux data were used to calculate I(sp), propulsive efficiency, and thrust. The ion flux profiles show an unexpected depression on the thruster axis for low tank pressures that disappears as the tank pressure increases. Ion energies decrease as the flow rate and pressure increase, but the microwave power level affects the energy only negligibly. The calculated propulsion parameters demonstrate that the efficiency of the laboratory device is low, and that tank pressure greatly changes the performance.

Kaufman, D. A.↗

NASA NTRS1993

Arcjet space thrusters

Electrically powered arc jets which produce thrust at high specific impulse could provide a substantial cost reduction for orbital transfer and station keeping missions. There is currently a limited understanding of the complex, nonlinear interactions in the plasma propellant which has hindered the development of high efficiency arc jet thrusters by making it difficult to predict the effect of design changes and to interpret experimental results. A computational model developed at the University of Tennessee Space Institute (UTSI) to study laser powered thrusters and radio frequency gas heaters has been adapted to provide a tool to help understand the physical processes in arc jet thrusters. The approach is to include in the model those physical and chemical processes which appear to be important, and then to evaluate our judgement by the comparison of numerical simulations with experimental data. The results of this study have been presented at four technical conferences. The details of the work accomplished in this project are covered in the individual papers included in the appendix of this report. We present a brief description of the model covering its most important features followed by a summary of the effort.

Keefer, Dennis↗

NASA NTRS1993

Recent testing of 30 kW hydrogen arcjet thrusters

Haag, Thomas W.↗

NASA NTRS1994

Preliminary far-field plume sputtering characterization of the Stationary Plasma Thruster (SPT-100)

For electric propulsion devices to be considered for use on communications satellites, integration impacts must be examined in detail. Two phenomena of concern associated with highly energetic plumes are contamination via sputtered material from the thruster and sputter erosion of downstream surfaces. In order to characterize the net effect of both phenomena, an array of witness plates were mounted in several types of holders and were exposed to the SPT-100 thruster plume for 50 hours. Surface analysis of the witness plates revealed that in the most energetic regions of the plume, there was a net removal of material from the samples facing the thruster. In the peripheral regions, net deposits were observed and characterized by the changes in optical properties of these samples. Changes in surface properties of samples located in collimators were within experimental uncertainty.

Pencil, Eric J.↗

NASA NTRS1994

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.

Manzella, David H.↗

NASA NTRS1994

Performance of a quasi-steady, multi megawatt, coaxial plasma thruster

The Los Alamos National Laboratory Coaxial Thruster Experiment (CTX) has been upgraded to enable the quasisteady operation of magnetoplasmadynamic (MPD) type thrusters at power levels from 2 to 40 MW for 10 ms. Diagnostics include an eight position, three axis magnetic field probe to measure magnetic field fluctuations during the pulse; a triple Langmuir probe to measure ion density, electron temperature, and plasma potential; and a time-of-flight neutral particle spectrometer to measure specific impulse. Here we report on the experimental observations and associated analysis and interpretation of long-pulse, quasisteady, coaxial thruster performance in the CTX device.

Scheuer, Jay T.↗