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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

5kW xenon ion thruster lifetest

This paper describes the results of the first lifetest of a high power ring-cusp ion thruster. A 30-cm laboratory model thruster was operated steady-state at a nominal beam power of 5kW on xenon propellant for approximately 900 hours. This test was conducted to identify life-limiting erosion mechanisms and thruster design 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 have been incorporated to solve problems identified by the lifetest, are also described.

Patterson, Michael J.↗
NASA NTRS1975

Optical properties of mercury ion thruster exhaust plumes Significance for candidate SEP science instruments

Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume emission rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter was considered incompatible with simultaneous thruster operation.

Goldstein, R.↗
NASA NTRS1984

High frequency plasma generator for ion thrusters

The results of a program to experimentally develop two new types of plasma generators for 30 cm electrostatic argon ion thrusters are presented. The two plasma generating methods selected for this study were by radio frequency induction (RFI), operating at an input power frequency of 1 MHz, and by electron cyclotron heating (ECH) at an operating frequency of 5.0 GHz. Both of these generators utilize multiline cusp permanent magnet configurations for plasma confinement and beam profile optimization. The program goals were to develop a plasma generator possessing the characteristics of high electrical efficiency (low eV/ion) and simplicity of operation while maintaining the reliability and durability of the conventional hollow cathode plasma sources. The RFI plasma generator has achieved minimum discharge losses of 120 eV/ion while the ECH generator has obtained 145 eV/ion, assuming a 90% ion optical transparency of the electrostatic acceleration system. Details of experimental tests with a variety of magnet configurations are presented.

Goede, H.↗
NASA NTRS1975

Solutions for discharge chamber sputtering and anode deposit spalling in small mercury ion thrusters

Proposed solutions to the problems of sputter erosion and sputtered material spalling in the discharge chamber of small mercury ion thrusters are presented. The accelerated life test evaluated three such proposed solutions: (1) the use of tantalum as a single low sputter yield material for the exposed surfaces of the discharge chamber components subject to sputtering, (2) the use of a severely roughened anode surface to improve the adhesion of the sputter-deposited coating, and (3) the use of a wire cloth anode surface in order to limit the size of any coating flakes which might spall from it. Because of the promising results obtained in the accelerated life test with anode surfaces roughened by grit-blasting, experiments were carried out to optimize the grit-blasting procedure. The experimental results and an optimal grit-blasting procedure are presented.

Power, J. L.↗
NASA NTRS1999

NSTAR Ion Thrusters and Power Processors

The purpose of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) project is to validate ion propulsion technology for use on future NASA deep space missions. This program, which was initiated in September 1995, focused on the development of two sets of flight quality ion thrusters, power processors, and controllers that provided the same performance as engineering model hardware and also met the dynamic and environmental requirements of the Deep Space 1 Project. One of the flight sets was used for primary propulsion for the Deep Space 1 spacecraft which was launched in October 1998.

Bond, T. A.↗
NASA NTRS2019

Near Field Probe Measurements in the Plume of a NEXT Ion Thruster

Due to the extremely long lifetime and ground testing complications, such as facility backpressure effects, modeling is a necessary tool for validating mission lifetime requirements. Both NASA's Glenn Research Center and Jet Propulsion Laboratory have developed lifetime validation models for the NEXT ion thruster. The largest uncertainties in these models are due to unknown plasma properties, such as plasma potential, that occur very near to the grid. Previous studies have made measurements in the mid to far field. Here, we make very near field measurements of the plasma properties in the ion beam within 0.1 thruster radii of the grid using various probes. Emissive and triple Langmuir probes are swept through the plume to spatially resolve plasma potential, electron temperature and ion density very near to the grid. The goal is to provide refined inputs to the lifetime modeling efforts and increase the accuracy of the models.

Arthur, Neil↗
NASA NTRS2019

Near Field Probe Measurements in the Plume of a NEXT Ion Thruster

Due to the extremely long lifetime and ground testing complications, such as facility backpressure effects, modeling is a necessary tool for validating mission lifetime requirements. Both NASA's Glenn Research Center and Jet Propulsion Laboratory have developed lifetime validation models for the NEXT ion thruster. The largest uncertaintiesin these models are due to unknown plasma properties, such as fields plasma potential, that occur very near to the grid. Previous studies have made measurements in the mid to far field. Here, we make very near field measurements of the plasma properties in the ion beam within 0.1 thruster radii of the grid using various probes. Emissive and triple Lang-muir probes are swept through the plume to spatially resolve plasma potential, electron temperature and ion density very near to the grid. The goal is to provide refined inputs to the lifetime modeling efforts and increase the accuracy of the models.

Arthur, Neil A.↗
NASA NTRS1973

An engineering model 30 cm ion thruster

Thruster development at Hughes Research Laboratories and NASA Lewis Research Center has brought the 30-cm mercury bombardment ion thruster to the state of an engineering model. This thruster has been designed to have sufficient internal strength for direct mounting on gimbals, to weigh 7.3 kg, to operate with a corrected overall efficiency of 71%, and to have 10,000 hours lifetime. Subassemblies, such as the ion optical system, isolators, etc., have been upgraded to meet launch qualification standards. This paper presents a summary of the design specifications and performance characteristics which define the interface between the thruster module and the remainder of the propulsion system.

Poeschel, R. L.↗
NASA NTRS1978

Pulse ignition characterization of mercury ion thruster hollow cathode using an improved pulse ignitor

An investigation of the high voltage pulse ignition characteristics of the 8 cm mercury ion thruster neutralizer cathode identified a low rate of voltage rise and long pulse duration as desirable factors for reliable cathode starting. Cathode starting breakdown voltages were measured over a range of mercury flow rates and tip heater powers for pulses with five different rates of voltage rise. Breakdown voltage requirements for the fastest rising pulse (2.5 to 3.0 kV/micro sec) were substantially higher (2 kV or more) than for the slowest rising pulse (0.3 to 0.5 kV/micro sec) for the same starting conditions. Also described is an improved, low impedance pulse ignitor circuit which reduces power losses and eliminates problems with control and packaging associated with earlier designs.

Wintucky, E. G.↗
NASA NTRS1978

Pulse ignition characterization of mercury ion thruster hollow cathode using an improved pulse ignitor

An investigation of the high voltage pulse ignition characteristics of the 8-cm mercury ion thruster neutralizer cathode identified a low rate of voltage rise and long pulse duration as desirable factors for reliable cathode starting. Cathode starting breakdown voltages were measured over a range of mercury flow rates and tip heater powers for pulses with five different rates of voltage rise. Breakdown voltage requirements for the fastest rising pulse (2.5 to 3.0 kV/microsec) were substantially higher (2 kV or more) than for the slowest rising pulse (0.3 to 0.5 kV/microsec) for the same starting conditions. The paper also describes an improved, low impedance pulse ignitor circuit which reduces power losses and eliminates problems with control and packaging associated with earlier designs.

Wintucky, E. G.↗
NASA NTRS1981

Simulation of charge exchange plasma propagation near an ion thruster propelled spacecraft

A model describing the charge exchange plasma and its propagation is discussed, along with a computer code based on the model. The geometry of an idealized spacecraft having an ion thruster is outlined, with attention given to the assumptions used in modeling the ion beam. Also presented is the distribution function describing charge exchange production. The barometric equation is used in relating the variation in plasma potential to the variation in plasma density. The numerical methods and approximations employed in the calculations are discussed, and comparisons are made between the computer simulation and experimental data. An analytical solution of a simple configuration is also used in verifying the model.

Robinson, R. S.↗
NASA NTRS2000

Performance of Titanium Optics on a NASA 30 Cm Ion Thruster

The results of performance tests with two titanium optics sets are presented and compared to those of molybdenum optics. All tests were conducted on a 30 cm ion thruster that was nearly identical to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) thruster design. Optics performance tests were conducted over a thruster input power range of 0.5 to 4.6 kW. Optics performance including impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, near-field beam current density profiles, beam divergence angles, and beam divergence thrust correction factors were determined throughout this power range. The impingement-limited total voltages for titanium optics were within 10 to 55 V of those for molybdenum optics. Electron backstreaming limit magnitude as a function of peak beam current density for both molybdenum and titanium optics were within a few volts of each other, indicating similar hot grid gaps for these two grid materials during steady-state operation. Beam divergence half-angles at 90 percent of the total beam current and thrust correction factors for both titanium optics sets were within one degree and one percent, respectively, of those for molybdenum optics. When thruster power was increased to 2.3 kW immediately following discharge ignition, the titanium screen grid came into contact with the accelerator grid within five minutes of ignition. Relative to molybdenum, titanium's larger thermal expansion and smaller thermal conductivity likely caused the screen grid to thermally expand more relative to the accelerator grid during startup.

Soulas, George C.↗
NASA NTRS1976

Application of ion thruster technology to a 30-cm multipole sputtering ion source

A 30-cm electron-bombardment ion source has been designed and fabricated for micromachining and sputtering applications. This source has a multipole magnetic field that employs permanent magnets between permeable pole pieces. An average ion current density of 1 ma/sq cm with 500-eV argon ions was selected as a design operating condition. The ion beam at this operating condition was uniform and well collimated, with an average variation of + or -5 percent over the center 20 cm of the beam at a distance up to 30 cm from the ion source.

Robinson, R. S.↗
NASA NTRS1975

Dynamic characteristics of a 30-centimeter mercury ion thruster

The present work reports on measurements of the fluctuations in the beam current, discharge current, neutralizer keeper current, and discharge voltage of a 30-cm ion thruster made with 60Hz laboratory-type power supplies. The intensities of the fluctuations (ratio of the root-mean-square magnitude to time-average quantity) were found to depend significantly on the beam and magnetic baffle currents. The shape of the frequency spectra of the discharge plasma fluctuations was related to the beam and magnetic baffle currents. The predominant peaks of the beam and discharge current spectra occurred at frequencies less than 30 kilohertz. This discharge chamber resonance could be attributable to ion-acoustic wave phenomena. Cross-correlations of the discharge and beam currents indicated that the dependence on the magnetic baffle current was strong. The measurements revealed that the discharge current fluctuations directly contribute to the beam current fluctuations and that the power supply characteristics can modify these fluctuations.

Serafini, J. S.↗
NASA NTRS1995

Assessment Of C60 As A Propellant Material For Ion Thrusters

Report presents analyses and data to support proposed use of C60 (buckminsterfullerene) as alternative to Xe, current propellent material of choice for use in ion thrusters. Concept of using C60 for this purpose described in "Electrostatic Propulsion Using C60 Molecules" (NPO-18526).

Rapp, Don↗
NASA NTRS2011

NEXT Propellant Management System Integration With Multiple Ion Thrusters

As a critical part of the NEXT test validation process, a multiple-string integration test was performed on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to flight. A test plan was developed for the sub-system integration test for verification of PMS and thruster system performance and functionality requirements. Propellant management system calibrations were checked during the single and multi-thruster testing. The low pressure assembly total flow rates to the thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS operated in 1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare very favorably indicating that flow rates to Thruster PM1R were similar in all cases. Characterizations of discharge losses, accelerator grid current, and neutralizer performance were performed as more operating thrusters were added to the PMS. There were no variations in these parameters as thrusters were throttled and single and multiple thruster operations were conducted. The propellant management system power consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 C. The total power consumed by the PMS was 10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, and multiple thrusters addressed, in part, the NEXT PMS and propulsion system performance and functionality requirements.

Sovey, James S.↗