Gregory G. Spanjers

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.

AFRL MicroPPT Development for Small Spacecraft Propulsion

Abstract : A class of miniaturized pulsed plasma thrusters (PPT), known as MicroPPTs, is currently in development at the Air Force Research Laboratory. Operating by means of a surface discharge across a Teflon (TM) propellant fuel bar only a few millimeters in diameter, the MicroPPT delivers a thrust-to-power ratio of 5-10 Micro-N-s/J. In the near term, these thrusters can provide propulsive attitude control on 150-kilogram-class spacecraft using one-fifth the dry mass of conventional torque rods and reaction wheels. Eventually these thrusters are designed for primary and attitude control propulsion on future 25-kilogram-class spacecraft. Efforts to characterize MicroPPT performance and the thruster plume are underway. To this end, a modified torsional thrust stand has been developed for the purpose of accurately measuring the low-level thrust generated by the MicroPPT. A Herriott Cell interferometer is introduced to establish the plume electron and neutral densities. Comparison of the measured electron density with modeling predictions shows close agreement. Additionally, a Pockels cell has been developed to provide a zero-impedance MicroPPT breakdown voltage measurement, and an intensified CCD array has been used to characterize the divergence of both the thruster plume and the late-time particulate emission. A synopsis is presented of the status of the thrusters development, including lifetime, thermal, and environmental testing. (2 refs.)

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.

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.

Electromagnetic Effects in the Near-Field Plume Exhaust of a Micro-Pulsed-Plasma Thruster

A model is presented of the near-field plasma-plume of a pulsed plasma thruster (PPT). As a working example, a micro-PPT developed at the U.S. Air Force Research Laboratory is considered. This is a miniaturized design of the axisymmetric PPT with a thrust in the 10-µ N range that utilizes TeflonTM as a propellant. The plasma plume is simulated using a hybrid fluid‐particle-in-cell direct simulation Monte Carlo approach. The plasma plume model is combined with Teflon ablation and plasma generation models that provide boundary conditions for the plume. This approach provides a consistent description of the plasma flow from the surface into the near plume. The magnetic field diffusion into the plume region is also considered, and plasma acceleration by the electromagnetic mechanism is studied. Teflon ablation and plasma generation analyses show that the Teflon surface temperature and plasma parameters are strongly nonuniform in the radial direction. The plasma density near the propellant surface peaks at about 10 24 m −3 , whereas the electron temperature peaks at about 4 eV near the electrodes. The plume simulation shows that a region with high density is developed at a few millimeters from the thruster exit plane at the axis. This high-density region exists during the entire pulse, but the plasma density decreases from about 2 × × 10 22 m −3 at the beginning of the pulse down to 0.3 × × 10 22 m −3 at 5 µs. The velocity phase is centered at about 20 km/s in the axial direction. At later stages of the pulse, there are two ion populations with positive and negative radial velocity. Electron and neutral densities predicted by the plume model are compared with near-field measurements using a two-color interferometer, and good agreement is obtained.

Optimization Issues for a Micropulsed Plasma Thruster

Several issues related to the design of a micropulsed plasma thruster (µPPT) are considered. It is concluded that the choice of the optimal energy level for a given µPPT geometry is very important. If the discharge energy is small, propellant charring would limit the operational time of the thruster. It is found that the charring phenomenon is associated with nonuniformity (in the radial direction between the electrodes) in the propellant ablation rate. On the other hand, higher energy leads to discharge constriction on the positive electrode and causes azimuthal nonuniformity. Reasoning leading to such nonuniformity is considered, and general suggestions for optimal energy and thruster size selections are presented.

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.