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At least 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

Internal erosion rates of a 10-kW xenon ion thruster

A 30 cm diameter divergent magnetic field ion thruster, developed for mercury operation at 2.7 kW, was modified and operated with xenon propellant at a power level of 10 kW for 567 h to evaluate thruster performance and lifetime. The major differences between this thruster and its baseline configuration were elimination of the three mercury vaporizers, use of a main discharge cathode with a larger orifice, reduction in discharge baffle diameter, and use of an ion accelerating system with larger acceleration grid holes. Grid thickness measurement uncertainties, combined with estimates of the effects of reactive residual facility background gases gave a minimum screen grid lifetime of 7000 h. Discharge cathode orifice erosion rates were measured with three different cathodes with different initial orifice diameters. Three potential problems were identified during the wear test: the upstream side of the discharge baffle eroded at an unacceptable rate; two of the main cathode tubes experienced oxidation, deformation, and failure; and the accelerator grid impingement current was more than an order of magnitude higher than that of the baseline mercury thruster. The charge exchange ion erosion was not quantified in this test. There were no measurable changes in the accelerator grid thickness or the accelerator grid hole diameters.

Rawlin, Vincent K.↗
NASA NTRS1988

Internal erosion rates of a 10-kW xenon ion thruster

A 30 cm diameter divergent magnetic field ion thruster, developed for mercury operation at 2.7 kW, was modified and operated with xenon propellant at a power level of 10 kW for 567 h to evaluate thruster performance and lifetime. The major differences between this thruster and its baseline configuration were elimination of the three mercury vaporizers, use of a main discharge cathode with a larger orifice, reduction in discharge baffle diameter, and use of an ion accelerating system with larger acceleration grid holes. Grid thickness measurement uncertainties, combined with estimates of the effects of reactive residual facility background gases gave a minimum screen grid lifetime of 7000 h. Discharge cathode orifice erosion rates were measured with three different cathodes with different initial orifice diameters. Three potential problems were identified during the wear test: the upstream side of the discharge baffle eroded at an unacceptable rate; two of the main cathode tubes experienced oxidation, deformation, and failure; and the accelerator grid impingement current was more than an order of magnitude higher than that of the baseline mercury thruster. The charge exchange ion eroison was not quantified in this test. There were no measurable changes in the accelerator grid thickness or the accelerator grid hole diameters.

Rawlin, Vincent K.↗
NASA NTRS1988

Shuttle RCS primary thruster injector flow visualization

An image-transmitting fiber-optics scope with a dry gas purge of the optics head has been used to visually evaluate the condition of surplus thrusters in the Space Shuttle's Reaction Control System; it was subsequently applied to flight thrusters. The technique uses water for flow visualization, and obviates thruster disassembly. The innovative use of gas purging of a fiber-optics head allows the unobstructed and distortion-free viewing of the flow streams, and testing has shown the technique to be ideally suited to injector flow assessments following thruster exposure to extensive contamination.

Wells, Dennis L.↗
NASA NTRS1989

The effects of magnetic nozzle configurations on plasma thrusters

Plasma thrusters have been operated at power levels from 10kW to 0.1MW. 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 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 will be studied without applied fields and with applied magnetic nozzle fields. Unique computational studies will utilize existing codes which accurately include transport processes. Unique diagnostic studies will support the experimental studies to generate new data. Both computation and diagnostics will be 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 NTRS1989

A data acquisition and storage system for the ion auxiliary propulsion system cyclic thruster test

A nine-track tape drive interfaced to a standard personal computer was used to transport data from a remote test site to the NASA Lewis mainframe computer for analysis. The Cyclic Ground Test of the Ion Auxiliary Propulsion System (IAPS), which successfully achieved its goal of 2557 cycles and 7057 hr of thrusting beam on time generated several megabytes of test data over many months of continuous testing. A flight-like controller and power supply were used to control the thruster and acquire data. Thruster data was converted to RS232 format and transmitted to a personal computer, which stored the raw digital data on the nine-track tape. The tape format was such that with minor modifications, mainframe flight data analysis software could be used to analyze the Cyclic Ground Test data. The personal computer also converted the digital data to engineering units and displayed real time thruster parameters. Hardcopy data was printed at a rate dependent on thruster operating conditions. The tape drive provided a convenient means to transport the data to the mainframe for analysis, and avoided a development effort for new data analysis software for the Cyclic test. This paper describes the data system, interfacing and software requirements.

Hamley, John A.↗
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.↗