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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1971

Design and development of a small structurally integrated ion thruster system

A 5 cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. With 2-D vectorable grids, thrust is 0.56 mlb at a beam voltage of 1200 V, total mass efficiency is 64%, and electrical efficiency is 46.8%. Nonvectoring thrust of 0.41 is demonstrated at very low specific impulse with dielectric coated grids. Structural integrity is demonstrated with dielectric coated grids for shock (30 G), sinusoidal (9 G), and random (19 G squared/Hz) accelerations. System envelope is 31.6 cm long by 14 cm flange B.C. with a mass of 8.5 kg including 6.2 Kg mercury propellant.

Hyman, J., Jr.↗
NASA NTRS1979

Sputtering in mercury ion thrusters

Ground-based tests of Hg ion thrusters have identified sputter erosion of thruster components as one of the main life limiting phenomena. Subsequent measurements have revealed that sputtering rates can be affected by background gases at pressures as low as 10 to the -10th torr. With the recent interest in thin film technology, sputtering in the presence of reactive gases has been studied in great detail. This paper presents the results of many of those studies and applies them to the sputtering of electric thrusters. A model, which assumes that chemisorption is the dominant mechanism, is applied to the sputtering rate measurements of the screen grid of a 30-cm thruster in the presence of nitrogen. The model utilizes inputs from a variety of experimental and analytical sources. The model of environmental effects on sputtering was applied to thruster conditions of low discharge voltage and a discussion of the comparison of theory and experiment is presented.

Mantenieks, M. A.↗
NASA NTRS1982

Parallel plate radiofrequency ion thruster

An 8-cm-diam. argon ion thruster is described. It is operated by applying 100 to 160 Mhz rf power across a thin plasma volume in a strongly divergent static magnetic field. No cathode or electron emitter is required to sustain a continuous wave plasma discharge over a broad range of propellant gas flow. Preliminary results indicate that a large fraction of the incident power is being reflected by impedance mismatching in the coupling structure. Resonance effects due to plasma thickness, magnetic field strength, and distribution are presented. Typical discharge losses obtained to date are 500 to 600 W per beam ampere at extracted beam currents up to 60 mA.

Nakanishi, S.↗
NASA NTRS1982

Development of a large inert gas ion thruster

A 30 cm inert gas electrostatic ion thruster has been developed, exhibiting excellent performance. In the development, the effective anode area was reduced by altering the magnetic field geometry to improve plasma containment, consistent with operational stability. The propellant introduction scheme has the effect of 'folding' the discharge chamber without the increased wall loss penalty associated with a longer chamber. These features contribute to a low discharge cost (eV/ion) versus mass utilization characteristic which remains relatively flat even to high mass utilizations.

Steiner, G.↗
NASA NTRS1971

Ion thruster accelerator system Patent

Development and characteristics of ion thruster accelerator with single glass coated grid to provide increased ion extraction capability and larger diameter accelerator system

Banks, B. A.↗
NASA NTRS1981

Performance of a magnetic multipole line-cusp argon ion thruster

The performance of a 17-cm diameter line-cusp ion thruster, using argon and xenon propellants, is investigated. Basic optimization of the thruster was accomplished without ion extraction and tests were conducted with small vacuum facilities and power supplies. The evaluation of the sensitivity of the performance to changes in discharge electrode configurations and the magnetic circuit is discussed. Final optimization results show that an argon propellant utilization efficiency of 0.9, at 260 W of discharge power per beam ampere, was obtained with a thruster having 20 rows of magnets and 10 tubular anodes. The average ion beam current densities were as high as 13 mA/sq cm. The results are considered to be the highest yet achieved with argon propellant.

Sovey, J. S.↗
NASA NTRS2006

Analytical Ion Thruster Discharge Performance Model

A particle and energy balance model of the plasma discharge in magnetic ring-cusp ion thrusters has been developed. The model follows the original work of Brophy in the development of global 0-D discharge models that utilize conservation of particles into and out of the thruster and conservation of energy into the discharge and out of the plasma in the form of charged particles to the walls and beam and plasma radiation. The present model is significantly expanded over Brophy's original work by including self-consistent calculations of the internal neutral pressure, electron temperature, primary electron density, electrostatic ion confinement (due to the ring-cusp fields), plasma potential, discharge stability, and time dependent behavior during recycling. The model only requires information on the thruster geometry, ion optics performance and electrical inputs such as discharge voltage and currents, etc. to produce accurate performance curves of discharge loss versus mass utilization efficiency. The model has been benchmarked against the NEXIS Laboratory Model (LM) and Development Model (DM) thrusters, and successfully predicts the thruster discharge loss as a function of mass utilization efficiency for a variety of thrusters. The discharge performance model will be presented and results showing ion thruster performance and stability given.

discharge chambers↗
NASA NTRS2019

Uncertainty Quantification of Modeled Electron Backstreaming Failure for the NEXT Ion Thruster

Excessive electron back streaming is one of the primary life-limiting failure modes for gridded ion thrusters. Physics-based modeling of the optics grid erosion and electron back streaming margins is augmented with statistical uncertainty quantification techniques to generate a life expectancy distribution of this particular failure mechanism for the NEXT ion thruster. Generation of distributions instead of a single point estimate provide a more comprehensive picture of thruster failure probabilities and life expectation for various mission applications.

Yim, John T.↗
NASA NTRS1978

Prediction of plasma properties in mercury ion thrusters

A simplified theoretical model was developed which obtains to first order the plasma properties in the discharge chamber of a mercury ion thruster from basic thruster design and controllable operating parameters. The basic operation and design of ion thrusters is discussed, and the important processes which influence the plasma properties are described in terms of the design and control parameters. The conservation for mass, charge and energy were applied to the ion production region, which was defined as the region of the discharge chamber having as its outer boundary the surface of revolution of the innermost field line to intersect the anode. Mass conservation and the equations describing the various processes involved with mass addition and removal from the ion production region are satisfied by a Maxwellian electron density spatial distribution in that region.

Longhurst, G. R.↗
NASA NTRS2020

Ion Thruster Produced Roll Torque

Disturbances to the ion engine’s thrust vector will cause a spacecraft to spin about its axis if left unmanaged. Spin about the yaw and pitch axes can be easily handled by a gimbal with enough authority. Spin about the roll axis, however, must be handled by additional thrusters or reaction wheels. In order to capitalize on the high efficiency of their thrusters, missions utilizing electric propulsion as primary propulsion generally include long periods of thrusting (several years). It is necessary to quantify and understand the ion thruster produced roll torque because it will define the amount of chemical propellant that must be carried or the lifetime and quantity of momentum wheels required for the mission. The roll torque produced by the NASA’s Evolutionary Xenon Thruster (NEXT) ion thruster is analyzed through a combination of theoretical calculations and magnetic field simulations. Experimental techniques for measuring roll torque and past flight data are also discussed.

Neil A Arthur↗
NASA NTRS1975

Accelerated life test of sputtering and anode deposit spalling in a small mercury ion thruster

Tantalum and molybdenum sputtered from discharge chamber components during operation of a 5 centimeter diameter mercury ion thruster adhered much more strongly to coarsely grit blasted anode surfaces than to standard surfaces. Spalling of the sputtered coating did occur from a coarse screen anode surface but only in flakes less than a mesh unit long. The results were obtained in a 200 hour accelerated life test conducted at an elevated discharge potential of 64.6 volts. The test approximately reproduced the major sputter erosion and deposition effects that occur under normal operation but at approximately 75 times the normal rate. No discharge chamber component suffered sufficient erosion in the test to threaten its structural integrity or further serviceability. The test indicated that the use of tantalum-surfaced discharge chamber components in conjunction with a fine wire screen anode surface should cure the problems of sputter erosion and sputtered deposits spalling in long term operation of small mercury ion thrusters.

Power, J. L.↗
NASA NTRS1975

Digital computer control of a 30-cm mercury ion thruster

The major objective was to define the exact role of an onboard spacecraft computer in the control of ion thrusters. An initial computer control system with accurate high speed capability was designed, programmed, and tested with the computer as the sole control element for an operating ion thruster. The command functions and a code format for a spacecraft digital control system were established. A second computer control system was constructed to operate with these functions and format. A throttle program sequence was established and tested. A two thruster array was tested with these computer control systems and the results reported.

Low, C. A., Jr.↗
NASA NTRS2004

Tailoring Ion-Thruster Grid Apertures for Greater Efficiency

A report proposes tailoring the diameters of the apertures in the accelerator grid of an ion thruster to reduce the open grid area through which un-ionized propellant gas can escape. The result would be a reduction in the loss of propellant gas and a corresponding increase in propellant efficiency. In a typical ion thruster, the plasma density decreases with radial distance from the centerline, and as a consequence, the diameters of ion beamlets decrease with increasing radial distance. According to the proposal, the apertures, through which the ion beamlets must pass, would be sized to match the diameters (with margin) of the beamlets. The decrease of the aperture diameters with radial distance would result in a significant reduction in the open grid area: In an example based on representative design parameters, the reduction could be as much as 30 percent. In this example, the transparency to un-ionized propellant would decrease from 0.24 to 0.17 and, as a result, the propellant efficiency would increase from 0.91 to 0.96.

Brophy, John↗
NASA NTRS1996

3-D Simulations of NSTAR Ion Thruster Plasma Environment

Full three-dimensional numerical models based on partical-in-cell with Monte Carlo collisions (PIC-MCC) simulations are developed to study ion thruster plasma interactions. 3-D simulation results for the NSTAR ion thruster plasma environment are presented.

NSTAR↗
NASA NTRS1983

Simplified power processing for ion-thruster subsystems

A design for a greatly simplified power-processing unit (SPPU) for the 8-cm diameter mercury-ion-thruster subsystem is discussed. This SPPU design will provide a tenfold reduction in parts count, a decrease in system mass and cost, and an increase in system reliability compared to the existing power-processing unit (PPU) used in the Hughes/NASA Lewis Research Center Ion Auxiliary Propulsion Subsystem. The simplifications achieved in this design will greatly increase the attractiveness of ion propulsion in near-term and future spacecraft propulsion applications. A description of a typical ion-thruster subsystem is given. An overview of the thruster/power-processor interface requirements is given. Simplified thruster power processing is discussed.

Wessel, F. J.↗
NASA NTRS2006

Structural Analysis of Pyrolytic Graphite Optics for the HiPEP Ion Thruster

The long lifetime requirements of interplanetary exploration missions is driving the need to develop long-life components for the electric propulsion thrusters that are being targeted for these missions. One of the primary life-limiting components of ion thrusters are the optics, which are continuously eroded during the operation of the thruster. Pyrolytic graphite optics are being considered for the High Power Electric Propulsion (HiPEP) ion thruster because of their very high resistance to erosion. This paper describes the structural analysis of the HiPEP pyrolytic graphite. A description of the development of the grid model, as well as the development of the effective properties and stress concentrations in the apertured area of the grids is included. An evaluation of the use of curved grids shows that the increased stiffness (compared to flat grids) prevents intergrid impact during launch, however, the residual stresses introduced by curving the grids pushes the resulting peak stresses beyond the critical stress. As a result, flat grids are recommended as the design solution. Thermally induced grid displacements during normal thruster operation are also presented.

Meckel, Nicole↗
NASA NTRS1982

Qualification test results of IAPS 8 cm ion thrusters

Two 8-cm-diameter 5-mN mercury ion thrusters with associated beam shields and gimbals have been fabricated and qualification-tested for the Ion Auxiliary Propulsion System (IAPS) flight test. Functional and performance tests of the thrusters, conducted before, during and after exposure to protoflight vibration and thermal vacuum conditions demonstrated no effects of the environmental exposure and little performance dispersion between the thrusters.

Collett, C.↗