Mercury Ion Thruster Research Articles

Model for computing volume-averaged plasma properties in electron-bombardment ion thrusters

This paper describes a model for computing volume-averaged values of the plasma properties inside electron-bombardment ion thrusters, using spatially varying Langmuir-probe measurements. The proposed model is intended to identify the correct plasma property averaging process that transforms the measured, spatially varying plasma parameters within a given thruster into their equivalent, spatially uniform values. This transformation technique allows thruster designs to be evaluated and compared on the basis of a single average value for each plasma parameter. In addition, it allows valid quantitative comparison with modern theoretical models of ion thruster discharge plasmas that are based on the characteristics of a spatially uniform plasma. Average values of the Maxwellian electron temperature and density, as well as the primary electron energy and density, are computed, compared, and interpreted for several different operational set points of two 30-cm diam mercury ion thrusters; a ring-cusp thruster and a J-series thruster, J-2. For low values of beam-ion-production cost, the superior performance of the ring-cusp thruster is correlated with a higher volume-averaged Maxwellian electron temperature and primary electron density compared to the J-series thruster. The volume-averaged plasma properties for both thrusters are compared with the analytical predictions of a detailed ion thruster performance model that is based on the characteristics of a spatially uniform plasma. Good quantitative and qualitative agreement is obtained between the volume-averaged and the ion thruster performance model calculations.

ETS-III ion engine flight operations in the extended mission period

An ion engine subsystem with two electron bombardment mercury ion thrusters of 5 cm anode diameter was flight tested on the Engineering Test Satellite-Ill during an extended mission from September 1983 to March 1985. Three tests were conducted: on/off cyclic tests for the both thrusters, continuous beam test for thruster 1, and continuous discharge test for thruster 2. The flight data were compared with that from the main mission (September 1982 to August 1983). With total operating hours and restarts of 66 h/47 cycles for the neutralize^ 64 h/53 cycles for the discharge, and 53 h/45 cycles for the beam on thruster 1 and 239 h/180 cycles for the neutralize!-, 356 h/188 cycles for the discharge, and 220 h/172 cycles for the beam on thruster 2, the ion production and extraction were found normal and unchanged. The hollow cathodes were found to be activated by repeating their operations. The operation of the entire spacecraft was ended with the depletion of the hydrazine used in the reaction control subsystem.

Experimental and computer model results on an ion thruster charge-exchange plasma

Ion thrusters can be used in a variety of primary and auxiliary space propulsion applications. A thruster produces a charge-exchange plasma that can interact with various systems on the spacecraft. The propagation of the charge-exchange plasma is crucial in determining the interaction of that plasma with the spacecraft. This paper compares experimental measurements with computer model predictions of the propagation of the chargeexchange plasma from a 30 cm mercury ion thruster. Plasma parameters compared include flow directions, ion densities, and directed energies. Good agreement is found in a region upstream of, and close to, the ion thruster optics. Outside of this region, the agreement is reasonable in view of the modeling difficulties.

Discharge-Chamber Sputtering in Mercury Ion Thrusters

Discharge-chamber sputter erosion has been investigated in two electron-bombardment mercury ion thrusters at 8-cm and 12-cm beam diameter (respectively, the SIT-8 and the SIT-12 thrusters). Sputter erosion is found to be concentrated primarily on the screen electrode and is due primarily to singly charged mercury ions, Hg + . A model of discharge-cha mber sputter-erosion has been developed that predicts erosion rates that are in good agreement with observed rates. A three fold reduction in the rate of total mass erosion has been achieved in both the SIT-8 and SIT-12 thrusters by biasing the thruster shell (including screen electrode) at keeper potential rather than at cathode potential. PUTTER erosion of discharge-cha mber surfaces has been identified as an important life-limiting process in electron- bombardment mercury ion thrusters.1'2 Present mercury thruster designs incorporate features that reduce the failure hazard posed by sputter erosion3; these measures appear to be adequate for small (8-cm-diameter) thrusters, but further improvement is needed in large (30-cm-diameter) thrusters intended for primary-propulsion application. Details of the discharge processes that cause sputter erosion have, however, remained largely unexplored. This paper describes the results of an investigation of discharge-cha mber sputter erosion and the development of several analytical and diagnostic tools that have been used to advantage in this investigation. In addition, a straightforwar d modification in discharge-cha mber design is described that has resulted in a three fold reduction in the rate of total mass erosion measured in operation of an 8-cm and a 12-cm mercury ion thruster. of the ion thruster fires downward into a frozen mercury collector, which was installed to avoid the experimental uncertainties that might result from backsputtering of nonvolatile material from a conventional solid-metal collector. Typical facility pressures (with an operating thruster) are in the range of 1 to 5 x 10 ~4 Pa. Several of the most important diagnostic measurements are described in greater detail in the following.

Projection ion lithography with aperture lenses

This paper reports on experiments in projection ion lithography in which multiple demagnified images have been etched directly into a substrate through an aperture lens array, without using a resist or proximity mask. The ion source is an 8-cm-beam-diam, Kaufman-type mercury ion thruster, which accelerates ions through a grid of perforated plates, and then neutralizes the resultant beamlets by injecting electrons at ground potential. The neutralized ion beam is allowed to pass through a suitable mask and thence onto an array of apertures of a few tens of microns in diameter. Ions impinging on the apertures are accelerated through a field produced by imposing a negative potential on a substrate placed just downstream of the aperture array. An image of the mask is etched in the substrate directly below each aperture. The substrate is a polished metal disk which has been sputter coated with a number of alternating layers of chromium and copper, each about 500 Å thick. Etching of these layers quickly produces an easily visible contour map which accurately depicts variations in the etch rate over the substrate surface. It is necessary to generate random motion of the substrate relative to the ion source to prevent projection of a superimposed image of the ion source grid system. To date, mask elements of the order of 2–3 mm in diameter have been used to produce multiple images demagnified by about 20–30× with resolution better than 10 μm. Suitable equipment is now being fabricated to permit exploration of the submicron region.

Cyclic Life-Test of an Ion Tfaruster Hollow Cathode

To assist in assessing the durability of the U.K. T4A and T5 10-cm mercury ion thrusters when used for NorthSouth stationkeeping, cyclic life-tests of thruster components have been carried out. This paper reports a cyclic discharge test of a hollow cathode in a diode configuration, together with thermal cycling tests of cathode heaters. The automatically controlled discharge test consisted of repetitive on and off periods with normal durations of 3 h and 40 min, respectively. The test was terminated voluntarily after 4200 h of discharge operation and 1343 starts. The rate of degradation of performance, as indicated by keeper and plasma potentials, was fully acceptable for the envisaged missions, as were the time, power, and energy required for discharge initiation. The thermal cycling tests demonstrated that the heater technology is satisfactory. It was concluded from a consideration of various aspects of the long-term operation of hollow cathodes that diode tests are adequately representative of thruster conditions.

Measurements of Plasma Fluctuations in an 8-cm Mercury Ion Thruster

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Dynamic Characteristics of a 30-cm Mercury Ion Thruster

Fluctuations of the discharge and beam plasmas of a 30-cm ion thruster have been measured using 60-Hz laboratory-type power supplies. The thruster was operated at beam currents of 2.0,1.5, and 1.0 A over a range of magnetic baffle currents, at constant discharge power and voltage. The time-varying properties of the discharge voltage and current, the ion beam current, and the neutralizer keeper current were measured. The intensities of the fluctuations (ratio of the rms magnitude to time-average quantity) were found to depend significantly on the beam and magnetic baffle currents. The predominant peaks of the beam and discharge current spectra occurred at frequencies less than 30kHz. Cross-correlations of the discharge and beam currents indicated that the dependence on the magnetic baffle current was strong.

Development of a 5-cm Flight-Qualified Mercury Ion Thruster

5 Space Shuttle Phase B System Study Final Report, Part II, Technical Summary, Rept. MDC E0308, June 1971, McDonnell Douglas Corp., St. Louis, Mo. 6 Orton, G. F. and Schweickert, T. F., Space Shuttle Auxiliary Propulsion System Design Study—Phase A, Requirements Definition, Rept. MDC E0603, Feb. 1972, McDonnell Douglas Corp., St, Louis, Mo. 7 Herr, P. N., private communication, April 1973, NASA Lewis Research Center, Cleveland, Ohio. 8 Kelly, P. J., Space Shuttle Auxiliary Propulsion System Design Study—Program Plan, Rept. MDC E0436, July 1971 (Revised Dec. 1971), McDonnell Douglas Corp., St. Louis, Mo. 9 Anglim, D. D., Bruns, A. E., Perryman, D. C., and Wieland D. L., Space Shuttle Auxiliary Propulsion System Design Study— Phase C and E Report, Storable Propellants RCS/OMS/APU Integration Study, Rept. MDC E0708, Dec. 1972, McDonnell Douglas Corp., St. Louis Mo.

Component Development for a 10-ciii Mercury Ion Thruster

BROAD range of component development, a part of which has already been reported in some detail,1'2 is in progress at the RAE in support of a 10-cm mercury ion thruster using a hollow cathode electron source.3 This paper presents the designs and describes the performance characteristics of the three major components developed for use in a pressurized mercury feed system. As the empirical approach to ion thruster investigations, although eventually effective, is hot thought to be economical of manpower or resources, an attempt has been made throughout this work to obtain a clear understanding of the physical principles involved in the operation of each component. This should enable future designs to be produced with confidence and will also allow assessments of the advantages of further development.

A Facility for Long-Term Ion Thruster System Testing

A space simulation facility was modified to provide the required orbital environment for operational testing of a NASA-furnished SERT II mercury ion thruster system for nearly a year of continuous operation. The test required that the thruster exhaust plume energy be absorbed without sputtering of material detrimental to the thruster. This was accomplished by developing a mercury collector and liner cooled by cryogenic nitrogen. A further requirement for electrical and thermal isolation of the collector and other items (at potentials as high as 5000 v in the presence of mercury vapor) was met by designing shielded insulators and vacuum passthroughs. Automatic safety features were incorporated to protect the test article, which eliminated the need for continuous staffing. The facility provided fail-safe control of the environment with only a very few minor malfunctions throughout the life test of the thruster system. The facility was subsequently cleaned and employed in other test programs; no contamination or degradation traceable to the long-term operation with mercury has been observed.

Plasma properties and performance of mercury ion thrusters

Electron bombardment mercury ion thrustor plasma properties and performance, computing ion beam current, discharge losses and propellant utilization efficiency

Research on hollow cathodes in mercury ion thrusters: WR Mickelsen and G Palumbo, Rep NASA-CR-109627, Mar 1970, 69 ages ( Sci Tech Aerospace Reps, 8 (12), 2286, N70-25703)

Research on hollow cathodes in mercury ion thrusters: WR Mickelsen and G Palumbo, Rep NASA-CR-109627, Mar 1970, 69 ages ( Sci Tech Aerospace Reps, 8 (12), 2286, N70-25703)

Automatic controls for facilities and Mercury ion thrusters in durability testing.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.