Lyon B. King

Study of Interspacecraft Coulomb Forces and Implications for Formation Flying

In the course of exploiting spacecraft formations for use in separated interferometry (or other missions), it is possible that the separation distance between vehicles will be on the order of 10 m. The effects of spacecraft charging on the dynamics of very closely spaced formations are investigated. For certain high-Earth orbits, the ambient plasma conditions will conspire to produce significant spacecraft charging in an environment with a plasma Debye length of more than 100 m. For such conditions, it is shown that the potential exists to develop disruptive interspacecraft Coulomb forces and torques, with magnitude comparable to candidate formation-keeping thrusters over distances of tens of meters. Because of the unexpectedly large interaction forces, the concept of purposely charging spacecraft to affect formation-keeping Coulomb forces is also explored. Analytic methods are developed that show the existence of static equilibrium formations in Earth orbit using only inter-vehicle coulomb forces for one-, two-, and three-dimensional formations. Such Coulomb formations would be free of the risk of plume contamination due to thrusters firing in close proximity. Figures of merit for the proposed Coulomb control system are calculated in a manner analogous to traditional propulsion systems, and it is shown that required forces can be created with milliwatts of power, can be controlled on a millisecond timescale, and imply specific impulses that can be as high as 1013 seconds.

Energetics of Propellant Options for High-Power Hall Thrusters

Krypton, cadmium, iodine, cesium, mercury, and bismuth are compared with xenon in several areas of performance, including thrust, specific impulse, probability of ionization, maximum theoretical efficiency, and sputter yield. The lighter propellants such as krypton and cadmium are favorable for high-Isp, low-thrust applications, whereas heavier propellants such as mercury and bismuth are preferable for low-Isp, high-thrust missions. Calculations of the ionizing collision rate show that cesium had the highest ionization probability (and lowest ionization energy), krypton had the lowest probability (and highest ionization energy), with the other propellant falling between cesium and krypton. Sputter erosion calculations show that for a carbon surface, heavier atoms will sputter less at low ion energies (less than 2000 eV) than light atoms, and will sputter much less on a kilograms-per-kilogram basis.

Mass Spectral Measurements in the Plume of an SPT-100 Hall Thruster

Knowledgeoftheion speciesemittedbytheSPT-100Hallthrusterprovidesconsiderableinsighttobasicthruster performance. To gain detailed information concerning the composition of the plume plasma, a unique molecular beam mass spectrometer was constructed to interrogate the exhaust plume. An investigation using this technique yielded boththeionizationfraction ofthexenonpropellantandaqualitativeanalysisoftheminorityplasma species attributed to ground-test facility interactions. The plasma was found to consist of 89% Xe + , 11% Xe 2+ , and 0.2% Xe 3+ . The existence of parasitic facility gasses including nitrogen, oxygen, and water vapor was documented, as well as trace carbon ions caused by sputtering of graphite surfaces in proximity of the thruster.

Gridded retarding pressure sensor for ion and neutral particle analysis in flowing plasmas

Although classical electrostatic plasma probes can provide detailed information regarding electrons and ions, these devices rely on charge-carrying particles both for energy filtering and for detection. Neutral particles are transparent to such diagnostics. In light of this a new probe design was developed to provide a measurement of the neutral particle flux component of a flowing plasma. The simple design adopted for this probe utilizes a set of electrostatic retarding grids to remove the charged particles from the flow. The neutral particles are sensed within the tube of an off-the-shelf hot cathode ionization gauge. Additionally, by varying the potential on the retarding grids, this same probe configuration can be used to measure the ion energy distribution. Preliminary data from this probe were taken in the exhaust plume of a Hall-current accelerator designed for space propulsion. This probe was found to be a very simple, accurate diagnostic tool for such use. By using this probe in combination with ...

Progress on the Development of a Direct Evaporation Bismuth Hall Thruster

It has been demonstrated that using segmented anodes for thermal regulation of anodes is possible by controlling where the discharge current attaches to. Thrust, Isp and efficiency measurements were taken on a segmented Hall thruster in order to ascertain what effect moving the discharge current attachment has on thruster performance. Overall, very little change in thrust, specific impulse and efficiency were measured across the operating spectrum when running on xenon. Thermal measurements were also taken but it was found that anode power density needed to be substantially increased to achieve the temperatures necessary for operation on bismuth. The anode face area was subsequently reduced and using a unique dualpropellant distributor, this work reports on experiments to use a xenon discharge as a “jump start” mechanism to provide waste heat necessary to initiate direct bismuth evaporation. Using the shim electrodes and magnetic fields for temperature control, the thruster is operated entirely on bismuth after a xenon warm-up stage.

Near- and Far-Field Plume Studies of a One-Kilowatt Arcjet

To support studies of transport in arcjet plumes, axial and radial profiles of electron temperature, electron number density, stagnation pressure, and flow field were obtained over an extensive volume of the plume of a 1-kW arcjet operating on hydrogen. All experiments were performed in a 6 by 9 m vacuum chamber at a tank pressure of less than 4 x 10 ~4 torr during arcjet operation. Electron temperatures obtained spectroscopically 1.2 cm downstream of the exit plane ranged from 0.10 to 0.13 eV, while electron number densities determined ~2 cm downstream of the exit plane via langmuir probe varied between 0.3-1 x 10 12 cm~3. Far-field langmuir probe measurements showed that a rapid radial variation in electron number density exists, ranging from 0.5 to 5 x 10 9cm~3, and from 0.5 to 2 x 10 9 cm~3, 30 and 88 cm downstream of the exit plane, respectively. Electron temperatures at these axial locations show much less of an axial dependence, ranging between 0.070.20 eV at both axial positions. Finally, an impact pressure probe was used to measure the radial profiles of stagnation pressure 53 and 64 cm from the exit plane as well as flow angle. The impact pressure probe data compare favorably with stagnation pressures predicted by a source-flow code and suggests that the heavy particles diffuse less radially than do the electrons.

Near and far-field plume studies of a 1 kW arcjet

In order to begin the process of characterizing transport in arcjet plumes, profiles of electron temperature, number density, pressure, and flow field patterns were obtained over an extensive volume of the plnme of a 1 kW arcjet operating on hydrogen. Axial and radial measnrements made over a region of the plume that extends from the arcjet exit plane to over one meter downstream of it are reported. Emission spectroscopy measurements of electron temperature were made near the exit plane of the arcjet. All experiments were performed in a 6 m by 9 m vacuum chamber with a pumping speed of over 100,000 l/s on hydrogen. Tank pressure was typically maintained to less than 3.8 x lor4 Torr during arcjet operation. In addition to the plume study, arcjet performance measurements are reportcd. 2 re rP kT, = electron temperature, eV V = probe bias, V Z, = ion charge vi,, = thrust efficiency y = ratio.of specific heats AD = Debye length, em A, = electron mean free path, cm A, = ion mean free path, cm aP = plasma potential, V x p = dimensionless probe potential 0 = flow angle = nozzle exit radius, m = radius from plume centerline, m = Langmuir probe collector electrode diameter, em r

Development of a Vaporizing Liquid Bismuth Anode for Hall Thrusters

Bismuth metal vapor Hall thrusters may have superior performance and economic characteristics when compared to xenon. From increased efficiency to reduced propellant and testing costs, bismuth seems to have a bright future. Of paramount importance when developing a practical bismuth device is the mechanism by which the propellant flow is controlled. This paper reports on an effort to use waste heat from the thruster to control the evaporation of a reservoir of liquid bismuth maintained within the discharge chamber. Research done thus far indicates that mass flow control can be achieved via a segmented anode configuration that serves as a thermostat to control input power into the bismuth reservoir. Thermal modeling has indicated that sufficient thermal gradients can be maintained between anode segments. Laboratory testing on xenon development thrusters validates the scheme to control reservoir temperature through discharge current sharing.

Ion-Energy Diagnostics in the Plasma Exhaust Plume of a Hall Thruster

Of primary concernwiththeintegration of Hall thrusterson conventional satellitedesigns isthepossibledamaging effect of high-energy exhaust ions impinging upon spacecraft surfaces. This paper reports on measurements of plasma ion-energy distributions within the plume of an SPT-100 Hall thruster using a custom-designed molecularbeam mass spectrometer. With this instrument ion energy was measured over a complete 360-deg circumference about the thruster at a radius of 0.5 m from the exit plane and over a total inclusive arc of 260 deg at 1.0-m radius. These data uncovered the existence of high-energy ions departing the thruster at angles exceeding 90 deg from the thrust vector and continuing well into the backe ow region of the plume. Through an analysis of the energy structure, the evidence of charge-exchange collisions occurring between plume ions and background neutrals was documented; such collisions produced anomalous distributions of ions having voltages greater than that applied to the thruster discharge. Nomenclature Ac = area of collector, m 2 d = analyzer plate separation, m Ei = ion energy, J or eV e = elementary charge, C f (ui) = ion velocity distribution, s/m f (Vi) = ion voltage distribution, s/m GCEM = gain of electron multiplier K45 = spectrometer constant k = Boltzmann constant, J/K L = analyzer interslit distance, m mi = mass of ion, kg ni = ion density, m i 3 q = ion integer charge state r = radial distance from thruster, m ui = ion velocity, m/s Vb = main distribution voltage, V Vi = ion voltage, V Vm = most-probable voltage, V Vp = repelling plate voltage, V w = analyzer slit width, m x = spatial coordinate, m y = spatial coordinate, m h = angle, deg s i = distribution half-width at e i 1 point, V

Ionic and neutral particle transport property measurements in the plume of an SPT-100

In order to understand the interaction between the SPT-100 thruster and the spacecraft accurate knowledge of exhaust plume properties must be obtained. This investigation utilized a suite of inter-related diagnostic probes in an effort to provide detailed knowledge of the exhaust plume on a molecular kinetic level. Probes used for the investigation included a Retarding Potential Analyzer, (RPA), a planar Faraday probe, both total and radiant heat flux probes, as well as a newly developed Neutral Particle Flux probe. Combining knowledge gained from these probes it was possible to quantify the transport of mass, energy, and charge within the plume at 0.5m and 1.0m radius from thruster exit. Among the phenomena uncovered were a high energy ionic annulus surrounding a narrow low energy core within the plume at 0.5m. This core was still incompletely filled in at 1.0m radius. Additionally, a population of energetic charge exchange neutrals was found along with the corresponding low energy charge exchange ions due to plume/background gas interactions.