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At least 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Test facility and preliminary performance of a 100 kW class MPD thruster

A 260 kW magnetoplasmadynamic (MPD) thruster test facility was assembled and used to characterize thrusters at power levels up to 130 kW using argon and helium propellants. Sensitivities of discharge characteristics to arc current, mass flow rate, and applied magnetic field were investigated. A thermal efficiency correlation developed by others for low power MPD thrusters defined parametric guidelines to minimize electrode losses in MPD thrusters. Argon and helium results suggest that a parameter defined as the product of arc voltage and the square root of the mass flow rate must exceed .7 V-kg(1/2)-s(-1/2) in order to obtain thermal efficiencies in excess of 60 percent.

Sovey, James S.↗
NASA NTRS1989

Performance of large area xenon ion thrusters for orbit transfer missions

Studies have indicated that xenon ion propulsion systems can enable the use of smaller Earth-launch vehicles for satellite placement which results in significant cost savings. These analyses have assumed the availability of advanced, high power ion thrusters operating at about 10 kW or higher. A program was initiated to explore the viability of operating 50 cm diameter ion thrusters at this power level. Operation with several discharge chamber and ion extraction grid set combinations has been demonstrated and data were obtained at power levels to 16 kW. Fifty cm diameter thrusters using state of the art 30 cm diameter grids or advanced technology 50 cm diameter grids allow discharge power and beam current densities commensurate with long life at power levels up to 10 kW. In addition, 50 cm diameter thrusters are shown to have the potential for growth in thrust and power levels beyond 10 KW.

Rawlin, Vincent K.↗
NASA NTRS1988

Experiments and analysis of a compact electrothermal thruster

The description and experimental performance of a compact microwave electrothermal thruster (MET) are presented. This thruster uses a coaxial applicator to couple microwave power into a high pressure discharge. Unlike earlier experiments, it uses no fused quartz in the discharge chamber or the nozzle. This allows high temperatures in the discharge chamber without quartz erosion and melting, thereby improving thruster performance and lifetime. The thruster design is compact, enhancing its potential as a space engine. Experimental tests using nitrogen and helium propellants with input powers levels of 200 W to 1.5 kW are presented. Experimental results, which produce energy efficiencies of 20 to 60 percent and specific impulse of 250 to 450 sec, compare favorably to previous experimental MET performance.

Asmussen, Jes↗
NASA NTRS1989

Deterministic errors in the Magellan orbit due to attitude control thruster activity

The pitch and yaw attitude control thrusters of the Magellan spacecraft generate uncoupled moments about the spacecraft center of mass, perturbing the spacecraft's orbit. A strategy to model these perturbing forces in the spacecraft equations of motion is presented. This strategy can model the thruster forces occurring in the nominal and contingency modes of attitude control operation. In the nominal mode, the thrusters fire only to balance the daily unloading of the momentum wheels. The contingency case will occur if a pitch or yaw axis momentum wheel fails, and the thrusters could fire up to 1200 times around the orbit to replace the failed wheel's momentum contribution. Deterministic errors are computed and shown to be under specified error requirements.

Engelhardt, Douglas B.↗
NASA NTRS1989

Performance of Large Area Xenon Ion Thrusters for Orbit Transfer Missions

Studies have indicated that xenon ion propulsion systems can enable the use of smaller earth-launch vehicles for satellite placement which results in significant cost savings. These analyses have assumed the availability of advanced, high power ion thrusters-operating at about 10 kW or higher. A program was initiated to explore the viability of operating 50 cm diameter ion thrusters at this power level. Operation with several discharge chamber and ion extraction grid set combinations was demonstrated and data were obtained at power levels to 16 kW. Fifty cm diameter thrusters using state of the art 30 cm diameter grids or advanced technology 50 cm diameter grids allow discharge power and beam current densities commensurate with long life at power levels up to 10 kW. In addition, 50 cm diameter thrusters are shown to have potential for growth in thrust and power levels beyond 10 kW.

Rawlin, Vincent K.↗
NASA NTRS1989

Test Facility and Preliminary Performance of a 100 kW Class MPD Thruster

A 260 kW magnetoplasmadynamic (MPD) thruster test facility was assembled and used to characterize thrusters at power levels up to 130 kW using argon and helium propellants. Sensitivities of discharge characteristics to arc current, mass flow rate, and applied magnetic field were investigated. A thermal efficiency correlation developed by others for low power MPD thrusters defined parametric guidelines to minimize electrode losses in MPD thrusters. Argon and helium results suggest that a parameter defined as the product of arc voltage and the square root of the mass flow rate must exceed 0.7 V/kg(sup 1/2)/sec(sup 1/2) in order to obtain thermal efficiencies in excess of 60 percent.

Sovey, J. S.↗
NASA NTRS1990

The effects of magnetic nozzle configurations on plasma thrusters

Plasma thrusters have been operated at power levels from 10 kw to 0.1 MW. When these devices have had magnetic fields applied to them which form a nozzle configuration for the expanding plasma, they have shown marked increases in exhaust velocity which is in direct proportion to the magnitude of the applied field. Further, recent results have shown that electrode erosion may be influenced by applied magnetic fields. This research effort is directed to the experimental and computational study of the effects of applied magnetic field nozzles in the acceleration of plasma flows. Plasma source devices which eliminate the plasma interaction in normal thrusters are studied as most basic. Normal thruster configurations were studied without applied fields and with applied magnetic nozzle fields. Unique computational studies utilize existing codes which accurately include transport processes. Unique diagnostic studies supported the experimental studies to generate new data. Both computation and diagnostics were combined to indicate the physical mechanisms and transport properties that are operative in order to allow scaling and accurate prediction of thruster performance.

York, Thomas M.↗
NASA NTRS1990

Anode power deposition in quasi-steady MPD thrusters

Spatially resolved anode heat flux measurements of a pulsed quasi-steady MPD thruster have been made by embedding thermocouples to the inner surface of a hollowed anode. Results obtained using argon propellant at mass flow rates of 4 and 16 g/s with the thruster operating at currents between 8 and 24 kA are presented. The results show that the anode fall voltage increases linearly with thruster current and does not depend on propellant mass flow rate. The fraction of thruster power deposited into the anode is 42 percent at 1 MW and falls to less than 20 percent at 6 MW. At any given operating condition, there is an inverse relationship between the anode fall and anode current density.

Kelly, Arnold J.↗
NASA NTRS1990

U.S. Space Station Freedom waste fluid disposal system with consideration of hydrazine waste gas injection thrusters

The results are reported of a study of various methods for propulsively disposing of waste gases. The options considered include hydrazine waste gas injection, resistojets, and eutectic salt phase change heat beds. An overview is given of the waste gas disposal system and how hydrozine waste gas injector thruster is implemented within it. Thruster performance for various gases are given and comparisons with currently available thruster models are made. The impact of disposal on station propellant requirements and electrical power usage are addressed. Contamination effects, reliability and maintainability assessments, safety issues, and operational scenarios of the waste gas thruster and disposal system are considered.

Winters, Brian A.↗
NASA NTRS1990

Calculations of gaseous H2/O2 thruster

Calculations were made for the gaseous H2/O2 thruster for the Space Station by using the RPLUS code which employs an implicit finite volume, LUSSOR scheme to solve the Navier-Stokes equations and the species equations. The combustion processes of hydrogen and oxygen are modeled by a 9 species and 18 step reaction mechanism and the turbulence is simulated by the Baldwin-Lomax turbulence model for the thruster wall boundary layer and the modified Prandtl's mixing length model for the reacting shear layer. Results are presented for different mixture ratios and fuel film cooling percents and compared with the experimental data. The calculated performance predictions for the thruster agree well with the experimental data and the results demonstrate that the RPLUS code can be used for design and analysis of thrusters and rockets.

Kim, S. C.↗
NASA NTRS1990

A 5-kW xenon ion thruster lifetest

The results of the first life test of a high power ring-cusp ion thruster are presented. A 30-cm laboratory model thruster was operated steady-state at a nominal beam power of 5 kW on xenon propellant for approximately 900 hours. This test was conducted to identify life-timing erosion modifications, and to demonstrate operation using simplified power processing. The results from this test are described including the conclusions derived from extensive post-test analyses of the thruster. Modifications to the thruster and ground support equipment, which were incorporated to solve problems identified by the lifetest, are also described.

Patterson, Michael J.↗
NASA NTRS1991

MPD thruster technology

MPD (MagnetoPlasmaDynamic) thrusters demonstrated between 2000 and 7000 seconds specific impulse at efficiencies approaching 40 percent, and were operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. To date, however, only a limited number of thruster configurations, propellants, and operating conditions were studied. The present status of MPD research is reviewed, including developments in the measured performance levels and electrode erosion rates. Theoretical studies of the thruster dynamics are also described. Significant progress was made in establishing empirical scaling laws, performance and lifetime limitations and in the development of numerical codes to simulate the flow field and electrode processes.

Myers, Roger M.↗
NASA NTRS1991

Anode power deposition in an applied-field segmented anode MPD thruster

Anode heat flux measurements of a water cooled segmented anode applied-field MPD thruster were made to investigate anode heat transfer phenomena. Pure argon and argon-hydrogen mixtures were used as propellants for a variety of thruster currents, propellant mass flow rates, and axial applied magnetic field strengths. The thruster was operated in two modes; with all four segments active, and with two of the segments floating. In addition, thrust and specific impulse were determined for each operating condition. The results show that the heat flux to the anode increases monotonically with axial magnetic field strength and thruster current. Between 50 and 75 percent of the anode heat flux is transported by the current carrying electrons. Convective and radiative heat transfer account for the remaining portion of the power deposited in the anode. The addition of hydrogen to the argon propellant results in the reduction of the fraction of anode power deposited by the anode fall to a level equivalent to that deposited by convection and radiation.

Myers, R. M.↗
NASA NTRS1991

MPD thruster technology

MPD thrusters have demonstrated between 2000 and 7000 sec specific impulse at efficiencies approaching 40 percent, and have been operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. This work reviews the present status of MPD thruster research, including developments in the measured performance levels and electrode erosion rates, and theoretical studies of the thruster dynamics. Significant progress has been made in establishing empirical scaling laws, performance and lifetime limitations, and in the development of numerical codes to simulate the flowfield and the electrode processes.

Myers, Roger M.↗
NASA NTRS1991

Coaxial plasma thrusters for high specific impulse propulsion

A fundamental basis for coaxial plasma thruster performance is presented and the steady-state, ideal MHD properties of a coaxial thruster using an annular magnetic nozzle are discussed. Formulas for power usage, thrust, mass flow rate, and specific impulse are acquired and employed to assess thruster performance. The performance estimates are compared with the observed properties of an unoptimized coaxial plasma gun. These comparisons support the hypothesis that ideal MHD has an important role in coaxial plasma thruster dynamics.

Schoenberg, Kurt F.↗
NASA NTRS1992

Multimegawatt MPD thruster design considerations

Performance and lifetime requirements for multimegawatt magnetoplasmadynamic (MPD) thrusters were used to establish a baseline 2.5 MW thruster design. The chamber surface power deposition resulting from current conduction, plasma and surface radiation, and conduction from the hot plasma was then evaluated to establish the feasibility of thruster operation. It was determined that state of the art lithium heat pipes were adequate to cool the anode electrode, and that the liquid hydrogen propellant could be used to cool the applied field magnet, cathode, and backplate. Unresolved issues having an impact of thruster design are discussed to help focus future research.

Myers, Roger M.↗
NASA NTRS1991

Derated ion thruster design issues

Preliminary activities to develop and refine a lightweight 30 cm engineering model ion thruster are discussed. The approach is to develop a 'derated' ion thruster capable of performing both auxiliary and primary propulsion roles over an input power range of at least 0.5 to 5.0 kilo-W. Design modifications to a baseline thruster to reduce mass and volume are discussed. Performance data over an order of magnitude input power range are presented, with emphasis on the performance impact of engine throttling. Thruster design modifications to optimize performance over specific power envelopes are discussed. Additionally, lifetime estimates based on wear test measurements are made for the operation envelope of the engine.

Patterson, Michael J.↗
NASA NTRS1992

Numerical simulation of geometric scale effects in cylindrical self-field MPD thrusters

A 2-D, two-temperature, single fluid magnetohydrodynamic code which incorporates classical plasma transport coefficients and Hall effects was developed to predict steady-state, self-field magnetoplasmadynamic (MPD) thruster performance. The governing equations and numerical methods of solution are outlined and discussed. Experimental comparisons are used to validate model predictions. The model accurately predicts thrust and reproduces trends in the discharge voltage for discharge currents below experimentally measured onset values. However, because the model does not include electrode effects, the calculated voltage drops are significantly lower than experimentally measured values. Predictions of thrust and flow efficiency are made for a matrix of fifteen cylindrical thruster geometries, assuming a fully ionized argon propellant. A maximum predicted specific impulse of 1680 s is obtained for a thruster with an anode radius of 2.5 cm, a cathode radius of 0.5 cm, and equal electrode lengths of 2.5 cm. A scaling relation is developed to predict, within limits, the onset of cylindrical, self-field thruster instability as a function of geometry and operating condition.

Lapointe, M.↗