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At least 955 records · Page 53
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

5200 cycle of an 8-cm diameter Hg ion thruster

An accelerated cycle test was conducted in which an 8-cm Engineering Model Thruster (EMT) prototype successfully completed 5200 on-off cycles and a total of more than 1300 hours of thruster operation at a 4.5 mN thrust level. Cathode tip heater powers required for starting and keeper voltages remained well within acceptable limits. The discharge chamber utilization and electrical efficiency were nearly constant over the duration of the test. It is concluded that on-off cyclic operation by itself does not appreciably degrade starting capability or performance of the 8-cm EMT.

Mantenieks, M. A.↗
NASA NTRS1978

Sensitivity of 30-cm mercury bombardment ion thruster characteristics to accelerator grid design

The design of ion optics for bombardment thrusters strongly influences overall performance and lifetime. The operation of a 30-cm thruster with accelerator grid open area fractions ranging from 43 to 24 percent, was evaluated and compared with previously published experimental and theoretical results. Ion optics properties measured included the beam current extraction capability, the minimum accelerator grid voltage to prevent backstreaming, ion beamlet diameter as a function of radial position on the grid and accelerator grid hole diameter, and the high energy, high angle ion beam edge location. Discharge chamber properties evaluated were propellant utilization efficiency, minimum discharge power per beam amp, and minimum discharge voltage.

Rawlin, V. K.↗
NASA NTRS1978

Ion beam plume and efflux measurements of an 8-cm mercury ion thruster

Measurements of the ion beam plume and efflux constituents of an 8-cm mercury ion thruster have been carried out in the TRW 5 x 10 foot testing chamber. Charged components (ion beam plume) were measured with an array of movable position Faraday cups and retarding potential analyzers yielding both current density and particle energy determinations. Neutral components (ion beam efflux) were determined with a movable position ionization gauge. Measurements of the ion beam plume were performed for a thruster both with and without a sputter shield. Analysis of data in terms of normalized effluxes has been carried out and has been applied to an example calculation of efflux compatibility with a communications spacecraft.

Komatsu, G. K.↗
NASA NTRS1978

Diagnostic evaluations of a beam-shielded 8-cm mercury ion thruster

An engineering model thruster fitted with a remotely actuated graphite fiber polyimide composite beam shield was tested in a 3- by 6.5-meter vacuum facility for in-situ assessment of beam shield effects on thruster performance. Accelerator drain current neutralizer floating potential and ion beam floating potential increased slightly when the shield was moved into position. A target exposed to the low density regions of the ion beam was used to map the boundaries of energetic fringe ions capable of sputtering. The particle efflux was evaluated by measurement of film deposits on cold, heated, bare, and enclosed glass slides.

Nakanishi, S.↗
NASA NTRS1978

Evolution of the 1-mlb mercury ion thruster subsystem

A general description and review of the auxiliary Electric Propulsion program, which led to the present 1-mlb (4.5 mN) thruster system, is presented. The developmental history, performance, and major lifetests of each component of the system are traced over the past 10 years. Major components include the 8-cm diameter ion thruster, the power processor, and the propellant reservoir and distribution system.

Kerslake, W. R.↗
NASA NTRS1977

Digital simulation of a 30-cm diameter ion thruster control system

A hybrid model of a 30-cm diameter ion thruster and transistorized power processor was formulated and implemented on a digital computer. The control loops as presented are stable for normal values of loop gain when propellant line delay times are zero. Small oscillations are present in the loops when representative thruster delay times are included in the model. Discharge supply characteristics can affect the control loop oscillations. Recovery from runaway can be aggravated by delay time and, to a greater extent, by discharge supply characteristics.

Kudo, I.↗
NASA NTRS1979

Plasma particle trajectories around spacecraft propelled by ion thrusters

The thruster plasma is assumed to be described by a collimated energetic beam and a cloud of ionized thermal propellant produced by charge-exchange. A simple adiabatic model is used to describe the expansion of these neutral plasmas away from the source. As the pressure falls, shielding currents dissipate, and the geomagnetic field takes control of the particles. In low earth orbit, it is concluded that the vehicle easily outruns its thruster plasma. At geosynchronous altitude, the local electric fields around high voltage surfaces collect return current from the thermal plasma that appears to be limited only by the available space charge. Results appropriate to proposed electric propulsion missions and the solar power satellite are presented and operational considerations are discussed.

Liemohn, H. B.↗
NASA NTRS1978

Inert gas thrusters

Inert gas thrusters have continued to be of interest for space propulsion applications. Xenon is of interest in that its physical characteristics are well suited to propulsion. High atomic weight and low tankage fraction were major factors in this choice. If a large amount of propellant was required, so that cryogenic storage was practical, argon is a more economical alternative. Argon was also the preferred propellant for ground applications of thruster technology, such as sputter etching and deposition. Additional magnetic field measurements are reported. These measurements should be of use in magnetic field design. The diffusion of electrons through the magnetic field above multipole anodes was studied in detail. The data were consistent with Bohm diffusion across a magnetic field. The theory based on Bohm diffusion was simple and easily used for diffusion calculations. Limited startup data were obtained for multipole discharge chambers. These data were obtained with refractory cathodes, but should be useful in predicting the upper limits for starting with hollow cathodes.

Kaufman, H. R.↗
NASA NTRS1979

Sputtering in mercury ion 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 NTRS1978

Convective heat flux in a laser-heated thruster

An analysis is performed to estimate the convective heating to the wall in a laser-heated thruster on the basis of a solution of the laminar boundary-layer equations with variable transport properties. A local similiarity approximation is used, and it is assumed that the gas phase is in equilibrium. For the thruster described by Wu (1976), the temperature and pressure distributions along the nozzle are obtained from the core calculation. The similarity solutions and heat flux are obtained from the freestream conditions of the boundary layer, in order to determine if it is necessary to couple the boundary losses directly to the core calculation. In addition, the effects of mass injection on the convective heat transfer across the boundary layer with large density-viscosity product gradient are examined.

Wu, P. K. S.↗
NASA NTRS1979

Pulsed plasma thruster backflow characteristics

The exhaust plume of a millipound pulsed plasma thruster has been investigated using a unique vacuum facility with cryogenically cooled low back-scatter chamber walls. The mass flux distribution of the plume was determined, with emphasis on the region upstream of the thruster. Measurements of the plume material scattered from the chamber walls were used with articulated collimator-quartz crystal microbalance measurements of the total flux to estimate the upstream flux originating in the plume. This flux was found to be of order 10 to the minus 10 g/sq cm/ pulse and to vary approximately with the inverse square of the radius. Measurements indicate the directed plume is essentially axisymmetric with a slight elongation in the direction parallel to the interelectrode gap. This downstream flux decreases with radius along a Gaussian profile with 90% of the plume confined to a 40 deg half-angle cone.

Rudolph, L. K.↗
NASA NTRS1979

A model for nitrogen chemisorption in ion thrusters

A theoretical model describing the formation of nitrogen species subject to chemisorption on ion thruster discharge chamber surfaces is presented. Molecules, atoms, atomic ions and molecular ions are identified as the important species in the analysis. Current densities of the atomic and molecular ions predicted by the model are compared to current densities measured in the beam of a SERT II thruster. The predicted and measured values of these two current densities are shown to agree within about + or - 100%. The mechanisms involved in the erosion of a surface subjected to simultaneous nitrogen chemisorption and sputter erosion by high energy ions are also discussed.

Wilbur, P. J.↗
NASA NTRS1979

Pulsed plasma thruster contamination studies

The exhaust plume of the one millipound pulsed plasma thruster has a measurable backflow upstream of the nozzle exit plane which may deposit on and degrade the performance of exposed spacecraft surfaces. High speed photographs and Faraday cup measurements suggest that this backflow is predominantly an electrically neutral, relatively low energy vapor. Articulated collimator quartz crystal microbalance measurements of this backflow were made for a thruster with a radically modified nozzle and a flat plate backflow shield, to determine the backflow sensitivity to nozzle design changes. The results are compared with the original nozzle backflow and show a measurable reduction in the backflow directly upstream of the shield.

Rudolph, L. K.↗
NASA NTRS1979

Interactions between a spacecraft and an ion thruster produced environment

A model of the charge-exchange plasma flow around a spacecraft is used to determine charge-exchange plasma densities and arrival rates at various locations on the spacecraft. A general assessment of particle deposition and effects of such are presented as they relate to the spacecraft and particular science instruments. Various aspects of the interactions between the ion thruster produced environment and a Comet Rendezvous (CR) spacecraft are addressed. No instruments will be affected adversely when the thrusters are off, excluding those that have passive radiators which may collect mercury.

Carruth, M. R., Jr.↗
NASA NTRS1981

Propagation of charge-exchange plasma produced by an ion thruster

Under the proper conditions there is an end-effect of a long, cylindrical Langmuir probe which allows a significant increase in collected ion current when the probe is aligned with a flowing plasma. This effect was used to determine the charge-exchange plasma flow direction at various locations relative to the ion thruster. The ion current collected by the probe as a function of its angle with respect to the plasma flow allows determination of the plasma density and plasma flow velocity at the probe's location upstream of the ion thruster optics. The density values obtained from the ion current agreed to within a factor of two of density values obtained by typical voltage-current Langmuir probe characteristics.

Carruth, M. R., Jr.↗
NASA NTRS1981

Interactions between a spacecraft and an ion thruster produced environment

The electron bombardment ion thruster is a candidate propulsion device for several proposed interplanetary missions. A comet rendezvous mission is expected to be the first use of a Solar Electric Propulsion System (SEPS). Because this is the first use of such a system, it is necessary to understand the interactions between the spacecraft and the environment produced by the SEPS. A preliminary assessment of the interactions between a thruster produced environment and the Comet Rendezvous spacecraft, including its science instruments, is presented which concludes that compatibility between the SEPS and the spacecraft can be obtained.

Carruth, M. R., Jr.↗
NASA NTRS1981

Recent work on an RF ion thruster

An experimental investigation of an rf ion thruster using an immersed coupler in an argon discharge is reported. The conical coil, used to couple rf power into the discharge, is placed inside the discharge vessel. The discharge was self-sustained by 100-150 MHz rf power at low environmental pressures. The ion extraction was accomplished by conventional accelerated grid optics from an unoptimized 8 cm diameter ion thruster.

Lee, R. Q.↗
NASA NTRS1981

MPD thruster application study

Developmental considerations for the magneto-plasma-dynamic (MPD) thruster are defined. General characteristics of an MPD engine are compared to those of chemical propulsion and ion bombardment engines and performance criteria which are mission specific are examined. Requirements for thruster ground testing facilities are discussed and the utilization of the space shuttle for an orbital flight test is addressed.

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