Ronald A. Spores

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.

Propellant Losses Because of Particulate Emission in a Pulsed Plasma Thruster

Abstract : Propellant inefficiency material in particulate form is characterized in a laboratory pulsed plasma thruster (PPT) operating at 1 Hz with a 204 discharge energy (20 W). Exhaust deposits are collected and analyzed using a combination of a scanning electron microscope with energy dispersive x-ray analysis and microscopic imaging. Teflon(trademark) particulates are observed with characteristic dianietens ranging from over 100 micrometers down to less than 1 micrometer.

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.

Overview of the USAF Electric Propulsion Program

Abstract : An overview of current electric propulsion research and development efforts within the United States Air Force is presented. The Air Force supports electric propulsion primarily through the Air Force Office of Scientific Research (AFOSR), the Air Force Research Laboratory (AFRL) and the AFOSR European Office of Aerospace Research and Development (EOARD). Overall direction for the programs comes from Air Force Space Command (AFSPC), with AFRL mission analysis used to define specific technological advances needed to meet AFSPC priorities. AFOSR funds basic research in electric propulsion throughout the country in both academia and industry. The AFRL Propulsion Directorate conducts electric propulsion efforts in basic research, engineering development, and space flight experiments. EOARD supports research at foreign laboratories that feeds directly into AFSR and AFRL research programs. Current research efforts fall into 3 main categories defined loosely by the thruster power level. All three agencies are conducting research at the low-power regime (P 30 kW) is realizing increased emphasis.

Effect of Propellant Temperature on Efficiency in the 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 and accelerated by an electric discharge across the propellant face. Previous research has concluded that as little as 10% of the consumed propellant is converted to plasma and efficiently accelerated. The remaining propellant is consumed in the form of late-time vaporization and particulate emission, creating minimal thrust. Critical to improving the PPT performance is improving the propellant utilization. The present work demonstrates one possible method of increasing the PPT propellant efficiency. By measuring the PPT thrust, propellant consumption, and propellant temperature while varying the power level, duration of the experimental run, and total propellant mass, a correlation is established between decreased propropellant temperature and increased propellant efficiency. The method is demonstrated by performance measurements at 60 W and S W, which show a 25% increase in thrust efficiency, while the propellant temperature decreases from 135 to 42 deg C. Larger increases in the efficiency may be realized on-orbit where operating temperatures are commonly subzero. The dependence of propellant consumption on temperature also creates systematic errors in laboratory measurements with short experimental runs, and orbit analyses where the PPT performance measured at one power level is linearly scaled to the power available on the spacecraft.

PPT Research at AFRL: Material Probes to Measure the Magnetic Field Distribution in a Pulsed Plasma Thruster

Abstract : The focus of the PPT basic research program at AFRL has now shifted to understanding the sources of the low energy efficiency. Based on previous research modifications such as changing the electrode geometry, discharge frequency, and discharge energy may all result in moderate increases to the energy efficiency. What is required from a basic research standpoint is a diagnostic capability that can acquire information with sufficient accuracy to enable PPT designers to understand why certain influences increase performance - and then design PPTs which maximize these effects. To model a fluid description of the PPT plasma, the critical measurements are magnetic field and density. Temperature, composition and charge state also become critical as the models become more detailed. This paper describes a magnetic field probe array used at AFRL to map the magnetic fields in a laboratory model PPT. The paper focuses on determining to what extent the probe perturbs the plasma, the measurement limitations. Also discussed are options towards making this critical measurement with increased accuracy.

An investigation of electrostatic probe perturbations on the operational characteristics of a Hall thruster and on the measurement of local plasma parameters

A high speed reciprocating probe system was constructed and used to investigate the perturbations of electrostatic probes on Hall thruster operations and on the measurement of local plasma parameters. Two regimes were investigated: inside the discharge chamber where the probe actively burned and at the exit plane where the probe experienced significant particle flux without burning. Experiments showed that in the interior of the thruster, significant probe material ablation occurred and severe perturbations to thruster operation and plasma measurements could not be avoided, even with residence times as short as 0.5 s. At the exit plane of the thruster, no material ablation was observed, however, significant variations of the local measurements of electron temperature and number density were observed. Measurement variations were successfully avoided by the use of the high-speed probe, which kept the probe residence time to less than 0.5 s. ∗ Research Aerospace Engineer, AFRL Electric Propulsion Laboratory, Member AIAA ξ Group Leader, AFRL Electric Propulsion Laboratory, Member AIAA ℘ Chief, Spacecraft Propulsion Branch, Member AIAA NOMENCLATURE A j Current collection area of electrode j A Power loading area Cp Specific heat e Electron charge f Particle flux Ie Electron current Ii Ion current Ji Ion current density K Kinetic energy kb Boltzmann constant k th Thermal conductivity l Probe length me Electron mass mi Ion mass ne Electron number density P Power to probe r Probe radius Te Electron temperature Ti Ion temperature t Melting time Tmax Melting temperature V j Potential of electrode j V f Floating potential Vp Plasma potential ∆V Probe bias v Particle velocity λd Debye length ρ Material density χdj Nondimensional potential difference between electrode j and electrode 1

MICROSECOND TIMESCALE SURFACE TEMPERATURE MEASUREMENTS IN MICRO-PULSED PLASMA THRUSTERS

A time-resolved surface temperature diagnostic for ablation-controlled arcs is in development at the Air Force Research Laboratory at Edwards AFB. The diagnostic draws on heritage from the experimental dynamic crack propagation community which has used photovoltaic infrared detectors to measure temperature rise in materials in the process of fracture. The microsecond time-scales involved in the fracture process suggest that such detectors may be applicable to the ablation-controlled discharges in pulsed plasma thrusters as a direct measurement of surface temperature during and after the arc. HgCdTe detectors are evaluated for use on the surface of an AFRL micro-pulsed plasma thruster. Evaluation of the diagnostic focuses on development of valid calibration methodology and application of the detector in the presence of a plasma. Current calibration techniques are reviewed with physical limitations discussed. An estimate for the wavelength dependent emissivity of Teflon is determined and measurements are performed in real-time on the surface of a thruster. Teflon vapor pressure is calculated for post-pulse temperature measurements.