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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 397 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Thermo-mechanical design aspects of mercury bombardment ion thrusters.

The mechanical design criteria are presented as background considerations for solving problems associated with the thermomechanical design of mercury ion bombardment thrusters. Various analytical procedures are used to aid in the development of thruster subassemblies and components in the fields of heat transfer, vibration, and stress analysis. Examples of these techniques which provide computer solutions to predict and control stress levels encountered during launch and operation of thruster systems are discussed. Computer models of specific examples are presented.

Schnelker, D. E.↗
NASA NTRS2008

Performance Evaluation of the Prototype Model NEXT Ion Thruster

The performance testing results of the first prototype model NEXT ion engine, PM1, are presented. The NEXT program has developed the next generation ion propulsion system to enhance and enable Discovery, New Frontiers, and Flagship-type NASA missions. The PM1 thruster exhibits operational behavior consistent with its predecessors, the engineering model thrusters, with substantial mass savings, enhanced thermal margins, and design improvements for environmental testing compliance. The dry mass of PM1 is 12.7 kg. Modifications made in the thruster design have resulted in improved performance and operating margins, as anticipated. PM1 beginning-of-life performance satisfies all of the electric propulsion thruster mission-derived technical requirements. It demonstrates a wide range of throttleability by processing input power levels from 0.5 to 6.9 kW. At 6.9 kW, the PM1 thruster demonstrates specific impulse of 4190 s, 237 mN of thrust, and a thrust efficiency of 0.71. The flat beam profile, flatness parameters vary from 0.66 at low-power to 0.88 at full-power, and advanced ion optics reduce localized accelerator grid erosion and increases margins for electron backstreaming, impingement-limited voltage, and screen grid ion transparency. The thruster throughput capability is predicted to exceed 750 kg of xenon, an equivalent of 36,500 hr of continuous operation at the full-power operating condition.

Herman, Daniel A.↗
NASA NTRS1995

Hall thruster ion beam characterization

The distribution of ion current density as a function of angle from the thruster axis was measured at nominal operating conditions of 300 Volts discharge voltage and 4.5 Amperes discharge current for two SPT-100 stationary plasma thrusters from Fakel Enterprises, a D-55 anode layer thruster from the Central Scientific Research Institute for Machine Building (TsNIIMASH), and a first, second, and third generation T-100 stationary plasma thruster from the Scientific Research Institute of Thermal Process (NIITP). The data showed that the current density distributions of these thrusters were similar. Some differences in peak ion current density were observed. Multiply charged ions were found to be a small fraction of the plasma plume for all of the thrusters. The effect of facility pressure on ion current density distribution was found to be nonnegligible at pressures above 2 x 10(exp -6) Torr. The ion current density distributions of a new SPT-100 and a 6000 hour wear tested SPT-100 exhibited no discernible difference. Ion current density measurements were also taken at off-nominal thruster operating conditions.

Manzella, David H.↗
NASA NTRS2004

Post-Test Analysis of the Deep Space One Spare Flight Thruster Ion Optics

The Deep Space 1 (DSl) spare flight thruster (FT2) was operated for 30,352 hours during the extended life test (ELT). The test was performed to validate the service life of the thruster, study known and identify unknown life limiting modes. Several of the known life limiting modes involve the ion optics system. These include loss of structural integrity for either the screen grid or accelerator grid due to sputter erosion from energetic ions striking the grid, sputter erosion enlargement of the accelerator grid apertures to the point where the accelerator grid power supply can no longer prevent electron backstreaming, unclearable shorting between the grids causes by flakes of sputtered material, and rouge hole formation due to flakes of material defocusing the ion beam. Grid gap decrease, which increases the probability of electron backstreaming and of arcing between the grids, was identified as an additional life limiting mechanism after the test. A combination of accelerator grid aperture enlargement and grid gap decrease resulted in the inability to prevent electron backstreaming at full power at 26,000 hours of the ELT. Through pits had eroded through the accelerator grid webbing and grooves had penetrated through 45% of the grid thickness in the center of the grid. The upstream surface of the screen grid eroded in a chamfered pattern around the holes in the central portion of the grid. Sputter deposited material, from the accelerator grid, adhered to the downstream surface of the screen grid and did not spall to form flakes. Although a small amount of sputter deposited material protruded into the screen grid apertures, no rouge holes were found after the ELT.

NASA Solar electric propulsion Technology Applicat↗
NASA NTRS1986

Improvement of ion thruster design

Two types of measurements were performed on ion thrustors equipped with SmCo magnets in either ring cusp or line cusp arrangements. Langmuir probes were used to measure plasma potential, electron density, and electron temperture in all regions inside the thruster. Loss fluxes to various surfaces were determined by measuring the currents to foils attached to or imbedded in the surface. Data were obtained for several sets of discharge voltages and currents. The loss currents were determined from current vs voltage characteristics observed on a transistor curve tracer oscilloscope. Both ion and electron currents were measured to all parts of the walls and to all parts of the cathode assembly using collecting plates. These measurement were also made for various parameter sets. In line cusp configuration the plasma density is essentially as predicted by existing calculations. In the ring cusp arrangement the interior of the plasma contains an inhomogeneous and relatively large magnetic field so the geometry is decidely two-dimensional and the models of Self (1967) and of Kino and Sham (1966) do not agree.

Carpenter, R. T.↗
NASA NTRS1972

Experimental investigation of a throttlable 15 cm hollow cathode ion thruster

The use of dished high perveance grids on a 15 cm modified SERT 2 thruster is shown to facilitate throttled operation over a beam current range from 60 to 600 mA. Effects of increasing the radial component of the magnetic field in the main discharge chamber and decreasing the dimensions of the cathode discharge region are examined and found to degrade performance to the extent that primary electrons are forced in toward the center-line of the thruster. Studies of the baffle aperture region of two thrusters indicate that the electric potential gradient vector is perpendicular to the local magnetic field lines when the thruster is operating properly. The correlation between the shape of the ion beam current density and that of the ion density at the screen grid within the thruster is shown to be 94%. Additional experimental studies on maximum propellant utilization, plasma ion production cost, neutral density in the cathode discharge region, double ion production in hollow cathode thrusters and thermal flow meter performance are discussed.

Wilbur, P. J.↗
NASA NTRS1973

Power processor for a 20CM ion thruster

A power processor breadboard for the JPL 20CM Ion Engine was designed, fabricated, and tested to determine compliance with the electrical specification. The power processor breadboard used the silicon-controlled rectifier (SCR) series resonant inverter as the basic power stage to process all the power to the ion engine. The breadboard power processor was integrated with the JPL 20CM ion engine and complete testing was performed. The integration tests were performed without any silicon-controlled rectifier failure. This demonstrated the ruggedness of the series resonant inverter in protecting the switching elements during arcing in the ion engine. A method of fault clearing the ion engine and returning back to normal operation without elaborate sequencing and timing control logic was evolved. In this method, the main vaporizer was turned off and the discharge current limit was reduced when an overload existed on the screen/accelerator supply. After the high voltage returned to normal, both the main vaporizer and the discharge were returned to normal.

Biess, J. J.↗
NASA NTRS1978

Effect of facility background gases on internal erosion of the 30-cm Hg ion thruster

Sputtering erosion of the upstream side of the molybdenum screen grid by discharge chamber ions in mercury bombardment thrusters was considered. Data which revealed that the screen grid erosion was very sensitive to the partial pressure of certain background gases in the space simulation vacuum facility were presented along with results of tests conducted to evaluate this effect. It is shown from estimates of the screen grid erosion in space that adequate lifetime for proposed missions exists.

Rawlin, V. K.↗