Frank S. Gulczinski

Performance Characteristics of a 5 kW Laboratory Hall Thruster

Abstract : The University of Michigan and United States Air Force Research Laboratory have jointly developed a 5 kW class Hall effect thruster. This thruster was developed to investigate, with a variety of diagnostics, a thruster similar to that specified by IHPRPT goals. The configuration of this thruster is adjustable so that diagnostic access to the interior of the thruster can be provided as necessary, and to allow for the exploration of various thruster geometries. At nominal conditions, the thruster was designed to operate at 5 kW with a predicted specific impulse of 2200 s. The actual operating parameters at 5 kW were 2326 s specific impulse, with 246 mN of thrust at an efficiency of 57%. These conditions are comparable to those of thrusters under commercial development, making the information learned from the study of this thruster applicable to the understanding of its commercial counterparts.

Analysis of Hall-effect thrusters and ion engines for orbit transfer missions

Analytical methods were combined with actual thruster data to create a model used to predict the performance of systems based on two types of electric propulsion thrusters, Hall-effect thrusters and ion engines, for several orbit transfer missions. Two missions were trip time constrained: a LEO-GEO transfer and a LEO constellation transfer. Hall thrusters were able to deliver greater payload due to their higher overall specific power. For the power limited orbit topping mission, the choice of thruster is dependent on the user’s need. Ion engines can deliver the greatest payload due to their higher specific impulse, but they do so at the cost of higher trip time. Study of reusable electric orbit transfer vehicle systems indicates that they can offer payload mass gains over chemical systems, but that these gains are less than those offered by other electric propulsion transfer scenarios due to the necessity of carrying propellant for return trips. Additionally, solar array degradation leads to increased trip time for subsequent reusable transfers. * Research Aerospace Engineer, Member AIAA ** Group Leader, USAF Electric Propulsion Lab, Member AIAA This paper is declared a work of the US Government and is not subject to copyright protection in the United States. INTRODUCTION: The US Air Force has recently completed several studies to investigate the potential advantages of advanced space propulsion for several orbit transfer scenarios. The first study investigated advanced propulsion concepts for expendable orbit transfer vehicles and concluded that the potential launch vehicle downsizing that resulted from the use of high specific impulse thrusters provided significant cost savings over base line chemical launch vehicle/upper stage systems. The second study looked at reusable advanced upper stages and preliminary indications are that while there remains the potential for launch vehicle downsizing, it is significantly reduced compared to expendable systems. This difference was largely due to the added propellant required to perform the round trip mission from low-earth orbit to geostationary orbit. Both studies pointed out advantages for advanced electric propulsion systems based on xenon propellant. The objective of this paper is to analyze the tradeoffs between Hall-effect thrusters and ion engines as a high power propulsion system for orbit transfer missions. Both the Hall-effect thruster and the gridded ion engine are classified as electrostatic thrusters and operate on heavy noble gases, primarily xenon. These electric propulsion devices are capable of specific impulses ranging from approximately 1500 to 4000 seconds, compared to chemical systems which typically operate in the range of 300 to 400 seconds. Electric propulsion is a type of rocket propulsion for space vehicles and satellites which utilizes electric and/or magnetic processes to accelerate a propellant at a much higher specific impulse than attainable using classical chemical propulsion. The concomitant reduction in required propellant mass results in increased payload mass capability. The method of analysis used in this study is based on the model developed by Messerole. It has been modified to reflect the most current information on thruster development levels and

Investigation of Propellant Inefficiencies in a Pulsed Plasma Thruster

Abstract : A Pulsed Plasma Thruster (PPT) benefits from the inherent engineering simplicity and reduced tankage fraction gained by storing the propellant as a solid. The solid is converted to the gaseous state during an electric surface discharge. Previous research has concluded that the bulk of the propellant expands gas-dynamically from the chamber at low directed velocity, with possibly as little as 10% ionized and efficiently accelerated to high velocity using electromagnetic forces. The two velocity components result in a low propellant utilization efficiency. Critical to improving the PPT efficiency is preventing the formation of the low-velocity propellant and/or developing a means of accelerating it electromagnetically. In the present work measurements are made of the solid propellant conversion to the gaseous state with the intent of better understanding the formation process. By better understanding the propellant conversion it is hoped that future PPTs can be designed with significantly increased propellant efficiencies.

Propellant Charring in Pulsed Plasma Thrusters

The Teflon ablation in a micro-pulsed plasma thruster is studied with the aim of understanding the charring phenomenon. Microscopic analysis of the charred areas shows that it contains mainly carbon. It is concluded that the carbon char is formed as a result of carbon flux returned from the plasma. A simplified model of the current layer near the Teflon surface is developed. The current density and the Teflon surface temperature have peaks near the electrodes that explain preferential ablation of these areas, such as was observed experimentally. Comparison of the temperature field and the ablation rate distribution with photographs of the Teflon surface shows that the area with minimum surface temperature and ablation rate corresponds to the charring area. This finding suggests that the charring may be related to a temperature effect.

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

Near-field ion energy and species measurements of a 5 kW laboratory Hall thruster

The Plasmadynamics and Electric Propulsion Laboratory (PEPL) has used its Molecular Beam Mass Spectrometer (MBMS) to determine the ion energy distribution of the P5 5 kW laboratory Hall thruster. A skimmer was used to obtain a sample of the plasma in the near-field region, 10 cm downstream of the thruster exit plane. The thruster was operated at several discharge conditions and was rotated with respect to the sampling skimmer in order to determine ion energy profiles at various plume angles. These measurements were compared to far-field ion energy measurements taken 75 cm from the discharge plane in order to examine the evolution of the ion energy profile and the effects of testing environment on the results. Both ion energy measurements and time-of-flight mass spectroscopy revealed evidence of singly, doubly, triply, and quadruply charged xenon ions within the plume. Ion energy distributions were used to determine that the thruster’s magnetic field was oriented such that the plume has an overall inward focus. It was seen that sampling the plasma closer to the thruster results in distributions that have undergone fewer changes as a result of collisions with other ions in the plume and with background neutrals within the test facility. Concerns regarding these facility effects have led to the development of a new Miniaturized Ion Energy Analyzer (MIEA) for future experimental work.

Resonance-probe measurements of plasma densities in electric-propulsion plumes

We are developing a resonance-probe plasma diagnostic that uses a microwave network analyzer for use in electric-propulsion research. To show the feasibility of our resonance-probe implementation, we have measured plasma densities in the plume of the 5-kW-class P5 Hall-effect thruster and compared them to measurements made with a Langmuir probe. Our preliminary work in this area indicates that the resonance-probe technique shows considerable promise. The resonance-probe technique should prove to be a useful tool to support electric-propulsion research since it able to provide high temporal-resolution electron density measurements. Nomenclature B 0 background magnetic-flux density, T c speed of light in a vacuum, 2.998 × 10 8 m/s k Boltzmanns constant, 1.38 × 10 −23 J/K l RP resonance-probe tip length, m m e electron mass, 9.109 × 10 −31 kg m i ion mass, kg n e electron plasma density, m −3 q charge magnitude, 1.602 × 10 −19 C T e electron temperature, K v p propagation (phase) velocity, m/s ε 0 free space permittivity, 8.85 × 10 −12 F/m λ wavelength of excitation frequency, m ω ce angular electron-cyclotron frequency, rad/s ω pe angular electron-plasma frequency, rad/s ω uh angular upper-hybrid frequency, rad/s