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The primary life-limiting mechanism for state-of-the-art ion thrusters is erosion of the accelerator grid by charge-exchange ions produced in the plume of the thruster. This paper reviews carbon-based grid development activities, specifically addressing grid manufacturing, optics performance, and launch and operational survivability.
Numerical results of particle-in-cell/Monte Carlo calculations of accelerator grid erosion in an ion thruster are presented. Specifically, it is shown that a three-dimensional model is required to account for the experimentally observed pitting of the accelerator grid between grid apertures. Some comparisons with earlier two-dimensional, axisymmetric model are made, and it is shown that, for identical operating conditions of the thruster, the wear-through time in the three-dimensional model is about two to three times higher than that obtained previously with the two-dimensional model, namely on the order of 10,000 hours for sample calculation.
Further numerical results are presented of earlier particle-in-cell/Monte Carlo calculations of accelerator grid erosion in an ion thruster. A comparison between numerical and experimental results suggests that the accelerator grid impingement is primarily due to ions created far downstream from the accelerator grid. In particular, for the same experimental conditions as those of Monheiser and Wilbur at Colorado State University, it is found that a downstream plasma density of 2 x 10 exp 14/cu m is required to give the same ratio of accelerator grid impingement current to beam current (5 percent). For this condition, a potential hill is found in the downstream region of 2.5 V.
Kaufman mercury-ion thrustor grid design, discussing radial slot configuration
Geometrically different sets of dished accelerator grids were tested on five 30-cm thrusters. The geometric variation of the grids included the grid-to-grid spacing, the screen and accelerator hole diameters and thicknesses, the screen and accelerator open area fractions, ratio of dish depth to the dish diameter, compensation, and aperture shape. In general, the data taken over a range of beam currents for each grid set included the minimum total accelerating voltage required to extract a given beam current and the minimum accelerator grid voltage required to prevent electron backstreaming.
Small hole accelerator grid ion optical systems have been tested as a possible means of improving 30-cm ion thruster performance. The effects of small hole grids on the critical aspects of thruster operation including discharge chamber performance, doubly-charged ion concentration, effluent beam characteristics, and plasma properties have been evaluated. In general, small hole accelerator grids are beneficial in improving thruster performance while maintaining low double ion ratios. However, extremely small accelerator aperture diameters tend to degrade beam divergence characteristics. A quantitative discussion of these advantages and disadvantages of small hole accelerator grids, as well as resulting variations in thruster operation characteristics, is presented.
Eighteen geometrically different sets of dished accelerator grids were tested on five 30-cm thrusters. The geometric variation of the grids included the grid-to-grid spacing, the screen and accelerator hole diameters and thicknesses, the screen and accelerator open area fractions, ratio of dish depth to dish diameter, compensation, and aperture shape. In general, the data taken over a range of beam currents for each grid set included the minimum total accelerating voltage required to extract a given beam current and the minimum accelerator grid voltage required to prevent electron backstreaming.
Increases in grid perveance through reductions in grid spacings without the attendant tendency toward arcing and shorting between grids at close spacing are shown to be achievable through substitution of higher-perveance dished grids for flat grids. Such dished grids can be spaced more closely than flat grids without exhibiting a propensity for shorting because their thermal warpage characteristics are more uniform and predictable.
Previous investigation under award NAG3-25 10 sought to determine the best method of LIF to determine the carbon density in a thruster plume. Initial reports from other groups were ambiguous as to the number of carbon clusters that might be present in the plume of a thruster. Carbon clusters would certainly affect the ability to LIF; if they were the dominant species, then perhaps the LIF method should target clusters. The results of quadrupole mass spectroscopy on sputtered carbon determined that minimal numbers of clusters were sputtered from graphite under impact from keV Krypton. There were some investigations in the keV range by other groups that hinted at clusters, but at the time the proposal was presented to NASA, there was no data from low-energy sputtering available. Thus, the proposal sought to develop a method to characterize the population only of atoms sputtered from a graphite target in a test cell. Most of the ground work had been established by the previous two years of investigation. The proposal covering 2003 sought to develop an anti-Stokes Raman shifting cell to generate VUW light and test this cell on two different laser systems, ArF and YAG- pumped dye. The second goal was to measure the lowest detectable amounts of carbon atoms by 156.1 nm and 165.7 nm LIF. If equipment was functioning properly, it was expected that these goals would be met easily during the timeframe of the proposal, and that is the reason only modest funding was requested. The PI was only funded at half- time by Glenn during the summer months. All other work time was paid for by Whitworth College. The college also funded a student, Charles Shawley, who worked on the project during the spring.
In classical gridded electrostatic ion thrusters, positively charged ions are generated from a plasma discharge of noble gas propellant and accelerated to provide thrust. To maintain overall charge balance on the propulsion system, a separate electron source is required to neutralize the ion beam as it exits the thruster. However, if high-electronegativity propellant gases (e.g., sulfur hexafluoride) are instead used, a plasma discharge can result consisting of both positively and negatively charged ions. Extracting such electronegative plasma species for thrust generation (e.g., with time-varying, bipolar ion optics) would eliminate the need for a separate neutralizer cathode subsystem. In addition for thrusters utilizing a RF plasma discharge, further simplification of the ion thruster power system may be possible by also using the RF power supply to bias the ion optics. Recently, the PEGASES (Plasma propulsion with Electronegative gases) thruster prototype successfully demonstrated proof-of-concept operations in alternatively accelerating positively and negatively charged ions from a RF discharge of a mixture of argon and sulfur hexafluoride.i In collaboration with NASA Marshall Space Flight Center (MSFC), the Georgia Institute of Technology High-Power Electric Propulsion Laboratory (HPEPL) is applying the lessons learned from PEGASES design and testing to develop a new thruster prototype. This prototype will incorporate design improvements and undergo gridless operational testing and diagnostics checkout at HPEPL in April 2014. Performance mapping with ion optics will be conducted at NASA MSFC starting in May 2014. The proposed paper discusses the design and preliminary performance testing of this electronegative gas plasma thruster prototype.
A substantial improvement in the performance of an 8-centimeter-diameter auxiliary propulsion thruster was achieved by reducing the diameter of the accelerator grid apertures. The accelerator grid hole geometry was defined by ion machining accelerator grids on an 8-centimeter thruster at thrust levels of 2.2, 4.4, and 6.7 millinewtons was (mN). A thruster with an ion machined accelerator grid was operated at a thrust of 4.4 mN for 1000 hours. The discharge propellant utilization was 92% at an eV/ion of 338. Thruster performance and accelerator grid hole geometry was documented as a function of thrust level. It was also determined that the small hole accelerator grid has a very low backstreaming voltage limit. In fact the thruster can be operated with the acclerator grid held at neutralizer tip potential.
A substantial improvement in the performance of an 8-centimeter-diameter auxiliary propulsion thruster has been achieved by reducing the diameter of the accelerator grid apertures. The accelerator grid hole geometry has been defined by ion machining accelerator grids on an 8-centimeter thruster at thrust levels of 2.2, 4.4, and 6.7 millinewtons (mN). A thruster with an ion machined accelerator grid has been operated at a thrust of 4.4 mN for 1000 hours. The discharge propellant utilization was 92% at an eV/ion of 338. Thruster performance and accelerator grid hole geometry has been documented as a function of thrust level. It was also determined that the small hole accelerator grid has a very low backstreaming voltage limit. In fact the thruster can be operated with the accelerator grid held at neutralizer tip potential.
In gridded electrostatic thrusters, ions are produced by electron bombardment in the discharge chamber. In most of these thrusters, a single, centrally located hollow cathode supplies the ionizing electrons. An applied magnetic field in the discharge chamber restricts the electrons leaving the hollow cathode to a very narrow channel. In this channel, the high electron current density ionizes both propellant gas flowing from the hollow cathode, and other neutrals from the main propellant flow from the plenum. The processes that occur just past the hollow cathode exit are very important. In recent engine tests, several cases of discharge cathode orifice place and keeper erosion have been reported. In this paper we present results from a new 1-D, variable area model of the plasma processes in the magnetized channel just downstream of the hollow cathode keeper. The model predicts plasma densities, and temperatures consistent with those reported in the literature for the NSTAR engine, and preliminary results from the model show a potential maximum just downstream of the cathode.
Several closely-space dished accelerator grid systems were fabricated and tested on a 30-cm diameter mercury bombardment thruster and they appear to be a solution to the stringent requirements imposed by the near-term, high-thrust, low specific impulse electric propulsion missions. The grids were simultaneously hydroformed and then simultaneously stress relieved. The ion extraction capability and discharge chamber performance were studied as the total accelerating voltage, the ratio of net-to-total voltage, grid spacing, and dish direction were varied.
Several closely-spaced dished accelerator grid systems have been fabricated and tested on a 30-cm diameter mercury bombardment thruster and they appear to be a solution to the stringent requirements imposed by the near-term, high-thrust, low specific impulse electric propulsion missions. The grids were simultaneously hydroformed and then simultaneously stress relieved. The ion extraction capability and discharge chamber performance were studied as the total accelerating voltage, the ratio of net-to-total voltage, grid spacing, and dish direction were varied.
A slightly modified 8-cm Hg ion thruster demonstrated significant increase in performance. Thrust was increased by almost a factor of five over that of the baseline thruster. Thruster operation with various three grid ion optics configurations; thruster performance as a function of accelerator grid open area, cathode baffle, and cathode orifice size; and a life test of 614 hours at a beam current of 250 mA (17.5 mN thrust) are discussed. Highest thruster efficiency was obtained with the smallest open area accelerator grid. The benefits in efficiency from the low neutral loss grids were mitigated, however, by the limitation such grids place on attainable ion beam current densities. The thruster components suffered negligible weight losses during a life test, which indicated that operation of the 8-cm thruster at extended levels of thrust and power is possible with no significant loss of lifetime.
Ion beamlets of predetermined configurations are formed by shaped apertures in the screen grid of an ion thruster having a double grid accelerator system. A plate is placed downstream from the screen grid holes and attached to the accelerator grid. When the ion thruster is operated holes having the configuration of the beamlets formed by the screen grid are sputtered through the plate at the accelerator grid.
NASA's Evolutionary Xenon Thruster (NEXT) is a 7-kW class gridded ion thruster-based propulsion system that was initially developed from 2002 to 2012 under NASAs In-Space Propulsion Technology Program to meet future science mission requirements. In 2015, a contract was awarded to Aerojet Rocketdyne, with subcontractor ZIN Technologies, to design, build and test two NEXT flight thrusters and two power processing units that would be available for use on future NASA science missions. Because an additional goal of this contract is to take steps towards offering NEXT as a commercialized system, it is called the NEXT-Commercial project, or NEXT-C. This paper reviews the capabilities of the NEXT-C system, status of the NEXT-C project, and the forward plan to build, test, and deliver flight hardware in support of future NASA and commercial applications. It also briefly addresses some of the potential applications that could utilize the hardware developed and built by the project.