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Polzin, Kurt A.

Publications and source records attributed to Polzin, Kurt A..

At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Iodine Hall Thruster Propellant Feed System for a CubeSat

The components required for an in-space iodine vapor-fed Hall effect thruster propellant management system are described. A laboratory apparatus was assembled and used to produce iodine vapor and control the flow through the application of heating to the propellant reservoir and through the adjustment of the opening in a proportional flow control valve. Changing of the reservoir temperature altered the flowrate on the timescale of minutes while adjustment of the proportional flow control valve changed the flowrate immediately without an overshoot or undershoot in flowrate with the requisite recovery time associated with thermal control systems. The flowrates tested spanned a range from 0-1.5 mg/s of iodine, which is sufficient to feed a 200-W Hall effect thruster.

Polzin, Kurt A.↗
NASA NTRS2014

The Iodine Satellite (iSAT) Hall Thruster Demonstration Mission Concept and Development

The use of iodine propellant for Hall thrusters has been studied and proposed by multiple organizations due to the potential mission benefits over xenon. In 2013, NASA Marshall Space Flight Center competitively selected a project for the maturation of an iodine flight operational feed system through the Technology Investment Program. Multiple partnerships and collaborations have allowed the team to expand the scope to include additional mission concept development and risk reduction to support a flight system demonstration, the iodine Satellite (iSAT). The iSAT project was initiated and is progressing towards a technology demonstration mission preliminary design review. The current status of the mission concept development and risk reduction efforts in support of this project is presented.

Dankanich, John W.↗
NASA NTRS2014

Design and Preliminary Testing Plan of Electronegative Ion Thruster

Electronegative ion thrusters are a new iteration of existing gridded ion thruster technology differentiated by their ability to produce and accelerate both positive and negative ions. The primary motivations for electronegative ion thruster development include the elimination of lifetime-limiting cathodes from a thruster system and the ability to generate appreciable thrust through the acceleration of both positive or negative-charged ions. Proof-of-concept testing of the PEGASES (Plasma Propulsion with Electronegative GASES) thruster demonstrated the production of positively and negatively-charged ions (argon and sulfur hexafluoride, respectively) in an RF discharge and the subsequent acceleration of each charge species through the application of a time-varying electric field to a pair of metallic grids similar to those found in gridded ion thrusters. Leveraging the knowledge gained through experiments with the PEGASES I and II prototypes, the MINT (Marshall's Ion-ioN Thruster) is being developed to provide a platform for additional electronegative thruster proof-of-concept validation testing including direct thrust measurements. The design criteria used in designing the MINT are outlined and the planned tests that will be used to characterize the performance of the prototype are described.

Schloeder, Natalie R.↗
NASA NTRS2014

Pulsed Inductive Plasma Acceleration: Performance Optimization Criteria

Optimization criteria for pulsed inductive plasma acceleration are developed using an acceleration model consisting of a set of coupled circuit equations describing the time-varying current in the thruster and a one-dimensional momentum equation. The model is nondimensionalized, resulting in the identification of several scaling parameters that are varied to optimize the performance of the thruster. The analysis reveals the benefits of underdamped current waveforms and leads to a performance optimization criterion that requires the matching of the natural period of the discharge and the acceleration timescale imposed by the inertia of the working gas. In addition, the performance increases when a greater fraction of the propellant is initially located nearer to the inductive acceleration coil. While the dimensionless model uses a constant temperature formulation in calculating performance, the scaling parameters that yield the optimum performance are shown to be relatively invariant if a self-consistent description of energy in the plasma is instead used.

Polzin, Kurt A.↗
NASA NTRS2013

Development And Testing Of The Inertial Electrostatic Confinement Diffusion Thruster

The Inertial Electrostatic Confinement (IEC) diffusion thruster is an experiment in active development that takes advantage of physical phenomenon that occurs during operation of an IEC device. The IEC device has been proposed as a fusion reactor design that relies on traditional electrostatic ion acceleration and is typically arranged in a spherical geometry. The design incorporates two radially-symmetric spherical electrodes. Often the inner electrode utilizes a grid of wire shaped in a sphere with a radius 15 to 50 percent of the radius of the outer electrode. The inner electrode traditionally has 90 percent or more transparency to allow particles (ions) to pass to the center of the spheres and collide/recombine in the dense plasma core at r=0. When operating the IEC, an unsteady plasma leak is typically observed passing out one of the gaps in the lattice grid of the inner electrode. The IED diffusion thruster is based upon the idea that this plasma leak can be used for propulsive purposes. The IEC diffusion thruster utilizes the radial symmetry found in the IEC device. A cylindrical configuration is employed here as it will produce a dense core of plasma the length of the cylindrical grid while promoting the plasma leak to exhaust through an electromagnetic nozzle at one end of the apparatus. A proof-of-concept IEC diffusion thruster is operational and under testing using argon as propellant (Figure 1).

Becnel, Mark D.↗
NASA NTRS2013

Standard Practices for Usage of Inductive Magnetic Field Probes with Application to Electric Propulsion Testing

Inductive magnetic field probes (also known as B-dot probes and sometimes as B-probes or magnetic probes) are often employed to perform field measurements in electric propulsion applications where there are time-varying fields. Magnetic field probes provide the means to measure these magnetic fields and can even be used to measure the plasma current density indirectly through the application of Ampere's law. Measurements of this type can yield either global information related to a thruster and its performance or detailed, local data related to the specific physical processes occurring in the plasma. Results of the development of a standard for B-dot probe measurements are presented, condensing the available literature on the subject into an accessible set of rules, guidelines, and techniques to standardize the performance and presentation of future measurements.

Polzin, Kurt A.↗
NASA NTRS2013

Standard Practices for Usage of Inductive Magnetic Field Probes with Application to Electric Propulsion Testing

Inductive magnetic field probes (also known as B-dot probes and sometimes as B-probes or magnetic probes) are useful for performing measurements in electric space thrusters and various plasma accelerator applications where a time-varying magnetic field is present. Magnetic field probes have proven to be a mainstay in diagnosing plasma thrusters where changes occur rapidly with respect to time, providing the means to measure the magnetic fields produced by time-varying currents and even an indirect measure of the plasma current density through the application of Ampère's law. Examples of applications where this measurement technique has been employed include pulsed plasma thrusters and quasi-steady magnetoplasmadynamic thrusters. The Electric Propulsion Technical Committee (EPTC) of the American Institute of Aeronautics and Astronautics (AIAA) was asked to assemble a Committee on Standards (CoS) for Electric Propulsion Testing. The assembled CoS was tasked with developing Standards and Recommended Practices for various diagnostic techniques used in the evaluation of plasma thrusters. These include measurements that can yield either global information related to a thruster and its performance or detailed, local data related to the specific physical processes occurring in the plasma. This paper presents a summary of the standard, describing the preferred methods for fabrication, calibration, and usage of inductive magnetic field probes for use in diagnosing plasma thrusters. Inductive magnetic field probes (also called B-dot probes throughout this document) are commonly used in electric propulsion (EP) research and testing to measure unsteady magnetic fields produced by time-varying currents. The B-dot probe is relatively simple in construction, and requires minimal cost, making it a low-cost technique that is readily accessible to most researchers. While relatively simple, the design of a B-dot probe is not trivial and there are many opportunities for errors in probe construction, calibration, and usage, and in the post-processing of data that is produced by the probe. There are typically several ways in which each of these steps can be approached, and different applications may require more or less vigorous attention to various issues.

Polzin, Kurt A.↗
NASA NTRS2013

Flow Control and Measurement in Electric Propulsion Systems: Towards an AIAA Reference Standard

Accurate control and measurement of propellant flow to a thruster is one of the most basic and fundamental requirements for operation of electric propulsion systems, whether they be in the laboratory or on flight spacecraft. Hence, it is important for the electric propulsion community to have a common understanding of typical methods for flow control and measurement. This paper addresses the topic of propellant flow primarily for the gaseous propellant systems which have dominated laboratory research and flight application over the last few decades, although other types of systems are also briefly discussed. While most flight systems have employed a type of pressure-fed flow restrictor for flow control, both thermal-based and pressure-based mass flow controllers are routinely used in laboratories. Fundamentals and theory of operation of these types of controllers are presented, along with sources of uncertainty associated with their use. Methods of calibration and recommendations for calibration processes are presented. Finally, details of uncertainty calculations are presented for some common calibration methods and for the linear fits to calibration data that are commonly used.

thrust measurement↗
NASA NTRS2013

Inductive Pulsed Plasma Thruster Development and Testing at NASA-MSFC

THE inductive pulsed plasma thruster (IPPT) is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. In the present work, we present a summary of the IPPT research and development conducted at NASA's Marshall Space Flight Center (MSFC). As a higher-power, still relatively low readiness level system, there are many issues associated with the eventual deployment and use of the IPPT as a primary propulsion system on spacecraft that remain to be addressed. The present program aimed to fabricate and test hardware to explore how these issues could be addressed. The following specific areas were addressed within the program and will be discussed within this paper. a) Conical theta-pinch IPPT geometry thruster configuration. b) Repetition-rate multi-kW thruster pulsing. c) Long-lifetime pulsed gas valve. d) Fast pulsed gas valve driver and controller. e) High-voltage, repetitive capacitor charging power processing unit. During the course of testing, a number of specific tests were conducted, including several that, to our knowledge, have either never been previously conducted (such as multi-KW repetition-rate operation) or have not been performed since the early 1990s (direct IPPT thrust measurements).2 Conical theta-pinch IPPT thrust stand measurements are presented in Fig. 1 while various time-integrated and time

Polzin, Kurt A.↗
NASA NTRS2013

Thrust Stand Measurements of a Conical Inductive Pulsed Plasma Thruster

Inductive Pulsed Plasma Thrusters (iPPT) spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current Propellant is accelerated and expelled at a high exhaust velocity (O(10 -- 100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, inductive pulsed plasma thrusters can suffer from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, inductive pulsed plasma thrusters can suffer from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. A conical coil geometry may offer higher propellant utilization efficiency over that of a at inductive coil, however an increase in propellant utilization may be met with a decrease in axial electromagnetic acceleration, and in turn, a decrease in the total axially-directed kinetic energy imparted to the propellant.

Hallock, Ashley K.↗
NASA NTRS2012

Optical multi-species gas monitoring sensor and system

The system includes at least one light source generating light energy having a corresponding wavelength. The system's sensor is based on an optical interferometer that receives light energy from each light source. The interferometer includes a free-space optical path disposed in an environment of interest. The system's sensor includes an optical device disposed in the optical path that causes light energy of a first selected wavelength to continue traversing the optical path whereas light energy of at least one second selected wavelength is directed away from the optical path. The interferometer generates an interference between the light energy of the first selected wavelength so-traversing the optical path with the light energy at the corresponding wavelength incident on the optical interferometer. A first optical detector detects the interference. At least one second detector detects the light energy at the at least one second selected wavelength directed away from the optical path.

Polzin, Kurt A.↗
NASA NTRS2012

Thrust Stand Measurements of a Conical Inductive Pulsed Plasma Thruster

Inductive Pulsed Plasma Thrusters (iPPT) are spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current. Propellant is accelerated and expelled at a high exhaust velocity (O(10 .. 100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, inductive pulsed plasma thrusters can suffer from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. A conical coil geometry may o er higher propellant utilization efficiency over that of a at inductive coil, however an increase in propellant utilization may be met with a decrease in axial electromagnetic acceleration, and in turn, a decrease in the total axially-directed kinetic energy imparted to the propellant.

Hallock, Ashley K.↗
NASA NTRS2012

Inductive Pulsed Plasma Thruster Model with Time-Evolution of Energy and State Properties

A model for pulsed inductive plasma acceleration is presented that consists of a set of circuit equations coupled to both a one-dimensional equation of motion and an equation governing the partitioning of energy. The latter two equations are obtained for the plasma current sheet by treating it as a single element of finite volume and integrating the governing equations over that volume. The integrated terms are replaced where necessary by physically-equivalent quantities that are calculated through the solution of other parts of the governing equation set. The model improves upon previous one-dimensional performance models by permitting the time-evolution of the energy and state properties of the plasma, the latter allowing for the tailoring of the model to different gases that may be chosen as propellants. The time evolution of the various energy modes in the system and the associated plasma properties, calculated for argon propellant, are presented to demonstrate the efficacy of the model. The model produces a result where efficiency is maximized at a given value of the electrodynamic scaling term known as the dynamic impedance parameter. Qualitatively and quantitatively, the model compares favorably with performance measured for two separate inductive pulsed plasma thrusters, with disagreements attributable to simplifying assumptions employed in the generation of the model solution.

Polzin, Kurt A.↗
NASA NTRS2012

Thrust Stand Measurements of a Conical Pulsed Inductive Plasma Thruster

Pulsed inductive plasma thrusters [1-3] are spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current. Propellant is accelerated and expelled at a high exhaust velocity (O(10-100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, pulsed inductive plasma thrusters can su er from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. The Microwave Assisted Discharge Inductive Plasma Accelerator (MAD-IPA)[4], shown in Fig. 1 is a pulsed inductive plasma thruster that is able to operate at lower pulse energies by partially ionizing propellant with an electron cyclotron resonance (ECR) discharge inside a conical inductive coil whose geometry serves to potentially increase propellant and plasma plume containment relative to at coil geometries. The ECR plasma is created with the use of permanent mag- nets arranged to produce a thin resonance region along the inner surface of the coil, restricting plasma formation and, in turn, current sheet formation to areas of high magnetic coupling to the driving coil.

Hallock, Ashley K.↗
NASA NTRS2012

Effect of Inductive Coil Geometry on the Operating Characteristics of an Inductive Pulsed Plasma Thruster

Operational characteristics of two separate inductive thrusters with conical theta pinch coils of different cone angles are explored through thrust stand measurements and time- integrated, unfiltered photography. Trends in impulse bit measurements indicate that, in the present experimental configuration, the thruster with the inductive coil possessing a smaller cone angle produced larger values of thrust, in apparent contradiction to results of a previous thruster acceleration model. Areas of greater light intensity in photographs of thruster operation are assumed to qualitatively represent locations of increased current density. Light intensity is generally greater in images of the thruster with the smaller cone angle when compared to those of the thruster with the larger half cone angle for the same operating conditions. The intensity generally decreases in both thrusters for decreasing mass flow rate and capacitor voltage. The location of brightest light intensity shifts upstream for decreasing mass flow rate of propellant and downstream for decreasing applied voltage. Recognizing that there typically exists an optimum ratio of applied electric field to gas pressure with respect to breakdown efficiency, this result may indicate that the optimum ratio was not achieved uniformly over the coil face, leading to non-uniform and incomplete current sheet formation in violation of the model assumption of immediate formation where all the injected propellant is contained in a magnetically-impermeable current sheet.

Hallock, Ashley K.↗
NASA NTRS2012

Comparison of Computed and Measured Performance of a Pulsed Inductive Thruster Operating on Argon Propellant

Pulsed inductive plasma accelerators are electrodeless space propulsion devices where a capacitor is charged to an initial voltage and then discharged through a coil as a high-current pulse that inductively couples energy into the propellant. The field produced by this pulse ionizes the propellant, producing a plasma near the face of the coil. Once a plasma is formed if can be accelerated and expelled at a high exhaust velocity by the Lorentz force arising from the interaction of an induced plasma current and the magnetic field. A recent review of the developmental history of planar-geometry pulsed inductive thrusters, where the coil take the shape of a flat spiral, can be found in Ref. [1]. Two concepts that have employed this geometry are the Pulsed Inductive Thruster (PIT)[2, 3] and the Faraday Accelerator with Radio-frequency Assisted Discharge (FARAD)[4]. There exists a 1-D pulsed inductive acceleration model that employs a set of circuit equations coupled to a one-dimensional momentum equation. The model was originally developed and used by Lovberg and Dailey[2, 3] and has since been nondimensionalized and used by Polzin et al.[5, 6] to define a set of scaling parameters and gain general insight into their effect on thruster performance. The circuit presented in Fig. 1 provides a description of the electrical coupling between the current flowing in the thruster I1 and the plasma current I2. Recently, the model was upgraded to include an equation governing the deposition of energy into various modes present in a pulsed inductive thruster system (acceleration, magnetic flux generation, resistive heating, etc.)[7]. An MHD description of the plasma energy density evolution was tailored to the thruster geometry by assuming only one-dimensional motion and averaging the plasma properties over the spatial dimensions of the current sheet to obtain an equation for the time-evolution of the total energy. The equation set governing the dynamics of the coupled electrodynamic-current sheet system is composed of first-order, coupled ordinary differential equations that can be easily solved numerically without having to resort to much more complex 2-D finite element plasma simulations.

Polzin, Kurt A.↗