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NASA NTRS2012

Magnetic Field Effects on Plasma Plumes

Here, we will discuss our numerical studies of plasma jets and loops, of basic interest for plasma propulsion and plasma astrophysics. Space plasma propulsion systems require strong guiding magnetic fields known as magnetic nozzles to control plasma flow and produce thrust. Propulsion methods currently being developed that require magnetic nozzles include the VAriable Specific Impulse Magnetoplasma Rocket (VASIMR) [1] and magnetoplasmadynamic thrusters. Magnetic nozzles are functionally similar to de Laval nozzles, but are inherently more complex due to electromagnetic field interactions. The two crucial physical phenomenon are thrust production and plasma detachment. Thrust production encompasses the energy conversion within the nozzle and momentum transfer to a spacecraft. Plasma detachment through magnetic reconnection addresses the problem of the fluid separating efficiently from the magnetic field lines to produce maximum thrust. Plasma jets similar to those of VASIMR will be studied with particular interest in dual jet configurations, which begin as a plasma loops between two nozzles. This research strives to fulfill a need for computational study of these systems and should culminate with a greater understanding of the crucial physics of magnetic nozzles with dual jet plasma thrusters, as well as astrophysics problems such as magnetic reconnection and dynamics of coronal loops.[2] To study this problem a novel, hybrid kinetic theory and single fluid magnetohydrodynamic (MHD) solver known as the Magneto-Gas Kinetic Method is used.[3] The solver is comprised of a "hydrodynamic" portion based on the Gas Kinetic Method and a "magnetic" portion that accounts for the electromagnetic behaviour of the fluid through source terms based on the resistive MHD equations. This method is being further developed to include additional physics such as the Hall effect. Here, we will discuss the current level of code development, as well as numerical simulation results

Ebersohn, F.↗

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 NTRS2014

Anomalous Thrust Production from an RF Test Device Measured on a Low-Thrust Torsion Pendulum

This paper describes the eight-day August 2013 test campaign designed to investigate and demonstrate viability of using classical magnetoplasmadynamics to obtain a propulsive momentum transfer via the quantum vacuum virtual plasma. This paper will not address the physics of the quantum vacuum plasma thruster, but instead will describe the test integration, test operations, and the results obtained from the test campaign. Approximately 30-50 micro-Newtons of thrust were recorded from an electric propulsion test article consisting primarily of a radio frequency (RF) resonant cavity excited at approximately 935 megahertz. Testing was performed on a low-thrust torsion pendulum that is capable of detecting force at a single-digit micronewton level, within a stainless steel vacuum chamber with the door closed but at ambient atmospheric pressure. Several different test configurations were used, including two different test articles as well as a reversal of the test article orientation. In addition, the test article was replaced by an RF load to verify that the force was not being generated by effects not associated with the test article. The two test articles were designed by Cannae LLC of Doylestown, Pennsylvania. The torsion pendulum was designed, built, and operated by Eagleworks Laboratories at the NASA Johnson Space Center of Houston, Texas. Approximately six days of test integration were required, followed by two days of test operations, during which, technical issues were discovered and resolved. Integration of the two test articles and their supporting equipment was performed in an iterative fashion between the test bench and the vacuum chamber. In other words, the test article was tested on the bench, then moved to the chamber, then moved back as needed to resolve issues. Manual frequency control was required throughout the test. Thrust was observed on both test articles, even though one of the test articles was designed with the expectation that it would not produce thrust. Specifically, one test article contained internal physical modifications that were designed to produce thrust, while the other did not (with the latter being referred to as the "null" test article). Test data gathered includes torsion pendulum displacement measurements which are used to calculate generated force, still imagery in the visible spectrum to document the physical configuration, still imagery in the infrared spectrum to characterize the thermal environment, and video imagery. Post-test data includes static and animated graphics produced during RF resonant cavity characterization using the COMSOL Multiphysics® software application. Excerpts from all of the above are included and discussed in this paper. Lessons learned from test integration and operations include identification of the need to replace manual control of the resonant cavity target frequency with an automated frequency control capability. Future test plans include the development of an automatic frequency control circuit. Test results indicate that the RF resonant cavity thruster design, which is unique as an electric propulsion device, is producing a force that is not attributable to any classical electromagnetic phenomenon and therefore is potentially demonstrating an interaction with the quantum vacuum virtual plasma. Future test plans include independent verification and validation at other test facilities.

Brady, David↗

NASA NTRS2014

Anomalous Thrust Production from an RF Test Device Measured on a Low-Thrust Torsion Pendulum

This paper describes the test campaigns designed to investigate and demonstrate viability of using classical magnetoplasmadynamics to obtain a propulsive momentum transfer via the quantum vacuum virtual plasma. This paper will not address the physics of the quantum vacuum plasma thruster (QVPT), but instead will describe the recent test campaign. In addition, it contains a brief description of the supporting radio frequency (RF) field analysis, lessons learned, and potential applications of the technology to space exploration missions. During the first (Cannae) portion of the campaign, approximately 40 micronewtons of thrust were observed in an RF resonant cavity test article excited at approximately 935 megahertz and 28 watts. During the subsequent (tapered cavity) portion of the campaign, approximately 91 micronewtons of thrust were observed in an RF resonant cavity test article excited at approximately 1933 megahertz and 17 watts. Testing was performed on a low-thrust torsion pendulum that is capable of detecting force at a single-digit micronewton level. Test campaign results indicate that the RF resonant cavity thruster design, which is unique as an electric propulsion device, is producing a force that is not attributable to any classical electromagnetic phenomenon and therefore is potentially demonstrating an interaction with the quantum vacuum virtual plasma.

Brady, David A.↗

NASA NTRS2018

Acceleration Modes and Transitions in Pulsed Plasma Accelerators

Pulsed plasma accelerators typically operate by storing energy in a capacitor bank and then discharging this energy through a gas, ionizing and accelerating it through the Lorentz body force. Two plasma accelerator types employing this general scheme have typically been studied: the gas-fed pulsed plasma thruster and the quasi-steady magnetoplasmadynamic (MPD) accelerator. The gas-fed pulsed plasma accelerator is generally represented as a completely transient device discharging in approximately 1-10 microseconds. When the capacitor bank is discharged through the gas, a current sheet forms at the breech of the thruster and propagates forward under a j (current density) by B (magnetic field) body force, entraining propellant it encounters. This process is sometimes referred to as detonation-mode acceleration because the current sheet representation approximates that of a strong shock propagating through the gas. Acceleration of the initial current sheet ceases when either the current sheet reaches the end of the device and is ejected or when the current in the circuit reverses, striking a new current sheet at the breech and depriving the initial sheet of additional acceleration. In the quasi-steady MPD accelerator, the pulse is lengthened to approximately 1 millisecond or longer and maintained at an approximately constant level during discharge. The time over which the transient phenomena experienced during startup typically occur is short relative to the overall discharge time, which is now long enough for the plasma to assume a relatively steady-state configuration. The ionized gas flows through a stationary current channel in a manner that is sometimes referred to as the deflagration-mode of operation. The plasma experiences electromagnetic acceleration as it flows through the current channel towards the exit of the device. A device that had a short pulse length but appeared to operate in a plasma acceleration regime different from the gas-fed pulsed plasma accelerators was developed by Cheng, et al. The Coaxial High ENerGy (CHENG) thruster operated on the 10-microseconds timescales of pulsed plasma thrusters, but claimed high thrust density, high efficiency and low electrode erosion rates, which are more consistent with the deflagration mode of acceleration. Separate work on gas-fed pulsed plasma thrusters (PPTs) by Ziemer, et al. identified two separate regimes of performance. The regime at higher mass bits (termed Mode I in that work) possessed relatively constant thrust efficiency (ratio of jet kinetic energy to input electrical energy) as a function of mass bit. In the second regime at very low mass bits (termed Mode II), the efficiency increased with decreasing mass bit. Work by Poehlmann et al. and by Sitaraman and Raja sought to understand the performance of the CHENG thruster and the Mode I / Mode II performance in PPTs by modeling the acceleration using the Hugoniot Relation, with the detonation and deflagration modes representing two distinct sets of solutions to the relevant conservation laws. These works studied the proposal that, depending upon the values of the various controllable parameters, the accelerator would operate in either the detonation or deflagration mode. In the present work, we propose a variation on the explanation for the differences in performance between the various pulsed plasma accelerators. Instead of treating the accelerator as if it were only operating in one mode or the other during a pulse, we model the initial stage of the discharge in all cases as an accelerating current sheet (detonation mode). If the current sheet reaches the exit of the accelerator before the discharge is completed, the acceleration mode transitions to the deflagration mode type found in the quasi-steady MPD thrusters. This modeling method is used to demonstrate that standard gas-fed pulsed plasma accelerators, the CHENG thruster, and the quasi-steady MPD accelerator are variations of the same device, with the overall acceleration of the plasma depending upon the behavior of the plasma discharge during initial transient phase and the relative lengths of the detonation and deflagration modes of operation.

Polzin, Kurt A.↗

NASA NTRS1996

Evaluation of Cryogenic Power Conditioning Subsystems for Electric Propulsion Spacecraft

The power requirement of vehciles desgined to transport cargo supporting a piloted expedition to Mars is in the range of megawatts. Therefore, it is imperative that the megawatt - class power processing unit designed for high-power nuclear electric propulsion vehicles using turboalternators and advanced magnetoplasmadynamic thrusters shouls be such that the overall system efficienty is as high as possible with minimum system specific mass.

cryogenic↗

NASA NTRS1994

SP-100 Dynamic Power and Lithium-Propellant MPD Nuclear Electric Propulsion Technology Requirements

The objective of this study was to evaluate the requirements (including system integration, design, test requirements, and schedule) for the propulsion and power conversion systems of a nuclear electric propulsion (NEP) vehicle using an SP-100 reactor with a dynamic power conversion system, Li-propellant magnetoplasmadynamic (MPD) thrusters, Li-propellant storage and feed systems, and the power conditioning electronics required to convert the power output from the power system to the form (voltage, current) needed by the thrusters. Potassium-Rankine power conversion systems have the potential for the greatest mission benefit in terms of minimum mass and volume (as compared to Brayton or Stirling power conversion systems), but they require the most development.

Murray, Kathy H.↗

NASA NTRSDate not supplied

Hybrid NEP-Chemical Vehicle and Propulsion Technology Study for Crewed Mars Missions

Nuclear Electric Propulsion (NEP) is a suite of technologies with the potential to enable crewed opposition-class missions to Mars. Each subsystem comprising an NEP system has multiple technology options that present tradeoffs in vehicle sizing and technology development risk. Multiple vehicle concepts have been studied in the past and are currently being studied at NASA. Three models have been developed to illustrate the impact of key technology parameters on overall sizing for a crewed Mars transportation vehicle. Vehicle mass required to close the mission as a function of NEP system alpha (kWe/kg) and specific impulse (Isp) is estimated with a detailed trajectory model. A power system mass model estimates alpha as a function of power and radiator mass assumption. A power conversion system thermodynamic model predicts the radiator area required to close a Brayton cycle power conversion system. Combined with the architecture mass required for mission closure, the radiator area and alpha model provide insight into how technology development may impact the mission. Our focus is on parametric sweeps of the whole design space rather than any particular point design. Results from the mission model comparing electric propulsion technology (Hall, Ion, Magnetoplasmadynamic) and both 2039 and 2042 opposition-class mission launch windows are included. The results of the study illustrate the technology parameters that can result in mission closure and illustrate where the architecture is sensitive to variations in technology performance or requirements.

Matthew Duchek↗

NASA NTRSDate not supplied

Evaluating the Performance of Hybrid Vehicles between LEO, Cislunar Space, and Mars

Nuclear power and electric propulsion technologies can enable a broad range of existing and future space mission concepts. This paper evaluates how these two technology classes, when combined into systems with various performance levels (and paired with a chemical propulsion system) affect vehicle sizing and mission feasibility for round-trip, opposition-class, crewed missions to Mars. The technology trade-space evaluated in this paper encompasses: various nuclear power system mass-efficiency levels (or power-specific mass), a range of output power levels available to the electric propulsion system, two electric propulsion thruster technologies (Hall and Magnetoplasmadynamic) at various characteristic specific impulse levels, and a range of propellant tank mass sizing coefficient. Each combination of the technologies is evaluated for the 2039 and the 2042 Mars mission opportunities for a range of imposed total mission duration constraints. The results show that synergies between technology performance levels exist and that increasing the performance of some parameters might have undesired consequences for the trajectory and vehicle. It was found that increasing the mass-efficiency of the nuclear power system is broadly beneficial, resulting in lower vehicle masses and enabling shorter mission durations. However, increasing the specific impulse of a thruster or increasing the available power level do not always yield a net benefit to the vehicle or mission; since increasing the specific impulse or power level (holding power system mass-efficiency constant) tends to decrease the vehicle thrust-to-mass ratio. Furthermore, the results show that mission feasibility is highly dependent on the mission duration; the required technology performance level to enable a mission is relaxed as the mission duration increases. Given the optimally sized vehicles resulting from the trade-space for a crewed mission to Mars, this paper also evaluates the impact of aggregating the Mars mission vehicle in Low-Earth Orbit and transferring it to the Mars mission departure node (using the on-board power and propulsion systems). The trade-space for this portion of the analysis includes the option of in-space vehicle refueling at the Mars mission departure node. The results show that on-orbit refueling can relax the required technology performance levels to achieve lower launch masses and orbit-raising durations at the cost of increased architecture complexity and risk.

Nuclear Electric Propulsion↗

NASA NTRS2003

Crewed Mission to Callisto Using Advanced Plasma Propulsion Systems

This paper describes the engineering of several vehicles designed for a crewed mission to the Jovian satellite Callisto. Each subsystem is discussed in detail. Mission and trajectory analysis for each mission concept is described. Crew support components are also described. Vehicles were developed using both fission powered magneto plasma dynamic (MPD) thrusters and magnetized target fusion (MTF) propulsion systems. Conclusions were drawn regarding the usefulness of these propulsion systems for crewed exploration of the outer solar system.

Callisto↗