MPD Thruster Research Articles

Exploring efficiency in next generation high temperature superconducting-enhanced applied field magnetoplasmadynamic thrusters: A combined numerical and experimental study

The integration of superconductors into space propulsion has emerged as a promising avenue. Superconductors have demonstrated significant effectiveness in space propulsion devices in terms of performance. This study introduces a low-power High-Temperature superconducting Applied Field MPD thruster, utilizing High-Temperature superconductors to propel plasma out of the thruster to generate thrust. Replacing traditional copper magnets with High-Temperature superconducting magnets has not only led to substantial improvements but has also addressed challenges such as the increased weight of traditional copper coils, the energy-intensive nature of copper magnets consuming high power, and the extremely low temperatures required by conventional superconductors, necessitating the use of helium cooling tanks, which can add to the overall weight of the thruster.The research combines numerical simulations with detailed experimental validation to evaluate performance. As predicted by the simulation model, experimental results demonstrate a thrust of 220 mN at 15 mg/s and 15.2 kW, and 180 mN at 5 mg/s and 14.8 kW, with a specific impulse of 2760s, while maintaining a discharge current of 200 A. Additionally, the discharge voltage in the experimental setup increases with magnetic field strength but decreases with mass flow rate. Plasma fall voltage also increases with both the applied magnet field strength set by the High-Temperature superconducting magnet and the mass flow rate. Meanwhile, the fractional electrode fall voltage decreases with the applied field. Discharge voltage, plasma fall voltage, and electrode fall voltage are analyzed as factors influencing the increase or decrease in thrust, specific impulse, and efficiency of the High-Temperature superconducting-enhanced Applied Field Magnetoplasmadynamic thruster.Furthermore, the study concludes with a comprehensive examination of the High-Temperature superconducting-enhanced Applied Field Magnetoplasmadynamic thruster, offering valuable insights for researchers. Notably, the achieved efficiencies for Applied field Magnetoplasmadynamic thruster at low input power surpass those of the same configuration with traditional copper magnet, highlighting the potential benefits of utilizing High Temperature superconducting magnets for enhanced performance.

準定常作動2cm級自己誘起磁場型MPDスラスタの推力特性評価

In order to understand the characteristics of the Self-field MPD thruster with a discharge chamber diameter of 2cm, a ground experiment system was constructed and the thrust characteristics were obtained. For the power supply system, a capacitor bank using an electric double-layer capacitor and IGBTs was adopted, and it was confirmed that rectangular discharge current pulses could be output at arbitrary times. In the operation check of the experimental system, the discharge current, discharge voltage, and ion saturation current remained constant during 2ms, confirming the quasi-steady operation of the small MPD thruster for 2ms. Thrust measurement experiments were conducted using argon as the propellant, and a thrust force of 0.9N was confirmed at a maximum discharge current of 2.2kA. The measured thrust was proportional to the square of the discharge current. Therefore, it was confirmed that electromagnetic acceleration was dominant. As the highest propulsive characteristics, a specific impulse of 1800s and a propulsive efficiency of 2.3% were confirmed.

Development of plasma beam irradiation facility using applied-field MPD thruster to study plasma-surface interactions

A plasma beam irradiation facility was developed based on the applied-field magnetoplasmadynamic (AF-MPD) thruster concept for studying plasma-surface interactions. The AF-MPD thruster was chosen because it can produce a plasma beam with high plasma density in continuous-wave mode. Two types of AF-MPD thruster were developed and used in this study: a type I source with a wide thruster channel was used for a heat flux test with Ar or Xe gas, while a type II source with a narrow thruster channel was used for an ion flux test with H2 or He gas. The plasma initially showed the characteristics of abnormal glow discharges and then a transition to arc occurred when the plasma current exceeded a threshold value. It was found that a cathode made of thoriated tungsten significantly lowered the threshold current for the transition from abnormal glow to arc. The maximum heat flux provided by our facility was measured to be 7 MW m−2 using a custom-made heat flux sensor, while the maximum hydrogen ion flux was measured to be 1 × 1023 m−2 s−1 using a Langmuir probe. The electron temperature ranged between (4–5) eV, while the electron density at the plasma plume (downstream) ranged between (1–4) × 1018 m−3.

Cathode temperature measurement of a hydrogen self-field MPD thruster during 1 ms quasi-steady operation

In this study, a 2D two-color pyrometer was developed to measure the cathode surface temperature distribution of a high-power hydrogen magnetoplasmadynamic thruster. The developed measurement system consists of an object lens, a beam splitter, bandpass filters, and two cameras. Wavelengths of 950 and 980 nm were selected to achieve a high signal-to-noise ratio by reducing the effects of plasma radiation. The cathode temperature distribution was measured in a discharge current range of 6–13 kA during approximately 1 ms quasi-steady operation. A spot-like high-temperature region over 3000 K and an overall low temperature were observed at low discharge currents. The overall cathode tip region was heated to approximately 2800 K at high discharge currents. The temperature distribution with high temperature only near the cathode tip was measured during quasi-steady operation, the duration of which is shorter than the duration of thermal conduction. These results show that the discharge current distribution near the cathode is stable during the time scale of sub-milliseconds. We confirmed that this temperature distribution remains constant during a 1 ms duty cycle through cathode temperature measurements at each operation time with a 0.2 ms exposure time. The measurement error was approximately 10% of the calculated temperature; a cross-check was performed utilizing near-infrared spectrum measurements.

Heat and particle load test facility using an applied-field MPD thruster for studying fusion divertor technology

A test facility for high heat and particle loads relevant to a fusion divertor has been developed using an applied-field magnetoplasmadynamic (AF-MPD) thruster. The AF-MPD thruster can provide high heat and particle fluxes to a target at a relatively low pressure with a relatively strong magnetic field. Several diagnostics for heat and particle fluxes such as a calorimeter and Langmuir probe were utilized to confirm that a heat flux of 4.0 MW m−2 and particle flux of 4.2 × 1022 m−2 s−1 were achieved using a 0.17 T permanent magnet. The electron temperature and electron density were estimated to be 4 eV and 1013 cm3, respectively, using optical emission spectroscopy based on the collisional-radiative model.

Detonation–deflagration one-dimensional mode transition analysis for parallel-plate plasma accelerators

Parallel-plate pulsed plasma thrusters (PPTs) and quasi-steady MPD thrusters are largely distinguished by the discrete current sheet propagation (also referred as detonation acceleration mode) and stationary steady-phase current patterns (deflagration acceleration mode), respectively. Experimental and numerical research attempts in the past have focused on showing that the two modes of operations could be explained using the same continuum process. Present work proposes a simple unified analytical energy balance model to capture this mode transition in pulsed electromagnetic thrusters. The model is tested for a high-energy PPT developed at NASA-Glenn Research Center, which operates predominantly in the quasi-steady mode. The influence of thruster geometry and circuit parameters on the mode of operation and electrical energy efficiency is analyzed. Based on the analysis of the energy transfer efficiency, it is found that operating the thruster in the quasi-steady mode improves the transfer efficiency by approximately $$10\%$$. It is concluded that the effective electrode length, pulse timing, circuit parameters, and electrode geometry could be optimized to operate the thruster in the quasi-steady/deflagration mode, hence improving the electrical energy transfer efficiency.

Numerical Analysis on the Plasma Behavior of a Hydrogen MPD Thruster at the Critical Current

Numerical Analysis on the Plasma Behavior of a Hydrogen MPD Thruster at the Critical Current

Increasing the effective voltage in applied-field MPD thrusters

Increasing the effective voltage is proposed to improve applied-field magnetoplasmadynamic thrusters (AF-MPDTs) efficiency. An analytical model for both effective voltage and thrust is presented and discussed on the basis of the generalized Ohm’s law. According to such a model, the effective voltage contributes to plasma acceleration and measures can be taken to affect it, including adjusting the fraction of cathode gas injection, applied magnetic field strength and discharge current. In order to verify the model, experiments were conducted on the conditions of stable operation of a 10 kW AF-MPDT. The results showed that discharge current affected thruster efficiency mainly by enhancing the back electromotive effect in plasma acceleration processes. The thrust efficiency was enhanced by an average of 9.1% by raising discharge current from 90 A to 180 A. The fraction of cathode gas injection mainly affected the resistive and Hall effect, and the thrust efficiency was enhanced by an average of 7.3% by increasing the fraction of cathode gas injection from 0.1 to 0.9. The applied magnetic field influenced the thruster performance by enhancing both the resistive and Hall effect and the back electromotive effect. The thrust efficiency was increased by an average of 3.0% by raising the applied magnetic field strength from 0.05 T to 0.17 T. It was found that the plasma acceleration of AF-MPDT was dominated by effective voltage, and increasing the effective voltage was an effective and comprehensive way to improve thruster performance.

The Effect of Anode Configuration on Hydrogen MPD Thruster Performance: A Numerical Study

The Effect of Anode Configuration on Hydrogen MPD Thruster Performance: A Numerical Study

Analytical investigation of anode heat flux in a magnetoplasmadynamic thruster

An analytical approach has been developed to analyze the anode thermal process in a self-field MPD thruster. Effects of both non-neutral and quasi-neutral plasma regions near the anode surface have been included. The heat flux calculation, which requires the contributions of current density and voltage drop, has been determined by deriving the analytical correlations between these parameters. The solution process has been formulated based on an iterative method to satisfy the condition of current density continuity at the mentioned regions interface. The anode of Princeton benchmark thruster has been investigated for ratios of the squared discharge current to mass flow rate of 16, 36, and 64 (kA)2·s/g. Generally, the calculated heat flux, current density, and voltage drop have shown a good agreement with the corresponding experimental data. Furthermore, the modeling has corroborated that the heat flux becomes maximum at the anode mid-plane where the total voltage drop is lowest. Also, it has been discussed that the second law of thermodynamics has obliged the current density to have a peak-shaped profile. Ultimately, a simple linearized model has been proposed to explain the spotty current effects on the non-neutral plasma structure.

A new mathematical model for studying fully ionized plasma flows in MPD thrusters

The paper describes a novel mathematical model to study the physics of fully ionized plasma flow through MPD thrusters otherwise known as magnetoplasmadynamic thrusters. The developed model consists of a set of differential equations obtained by coupling the Navier–Stokes equations with Maxwell’s equations. The model developed was tested for various 1D and 1.5D cases. These simulations have been carried out for a special case of geometries called Tikhonov fitted geometries for alleviating the instability phenomenon commonly known as onset phenomenon in self-field MPD thrusters. The work was also extended to applied field MPD thrusters. A new equation for magnetic field was introduced to incorporate a virtual cathode. All the computations were carried out using a higher order integration scheme. The results obtained were found to be in good agreement with the previous work done on similar models.

Effects of a magnetic Laval nozzle on an MPD thruster for various gas propellants

Effects of a magnetic Laval nozzle on an MPD thruster for various gas propellants

A phenomenological performance model for applied-field MPD thrusters

A theoretical model for the performance prediction of applied-field magnetoplasmadynamic thrusters (MPDTs) is presented. MPD thrusters have been long regarded as leading candidates for near-term, thrust-demanding missions due to their substantial thrust density and specific impulse even at moderate power levels (50–200kW). However, the complicated physics behind the acceleration mechanism as well as the challenging on-ground testing have delayed their development and optimization leading to a slow but constant decline of interest in such a technology. Despite several theoretical efforts in the last few decades, no complete and definitive understanding of the scaling relations governing their performance is yet available. In this work, a simple phenomenological model for both the thrust and the terminal voltage is presented and discussed. The validity of the model is then assessed through a systematic comparison with the experimental data available in the literature. It was found that the suggested model can actually capture most of the characteristic features of this class of thrusters within a 20% error for a wide range of operational conditions.

Numerical simulation of non-equilibrium plasma flow in a cylindrical MPD thruster using a high-order flux-difference splitting method

A two-dimensional axisymmetric computational algorithm is developed to simulate the plasma flow field in a self-field MPD thruster, in order to determine the flow behavior and the electromagnetic characteristics distribution. The convective flux vector is computed by using Roe׳s scheme in combination with Powell׳s eigensystem technique, and a new modified MUSCL technique called OMUSCL2 is employed to obtain the stable high-accuracy solution. Madrane–Tadmor entropy correction is used to prevent unrealistic expansion shocks near the electrodes tips. To accurately capture the physics of plasma in the system, different physical–chemical sub-models including multi-level non-equilibrium ionization model, generalized Ohm׳s law for partially ionized plasma, micro-instabilities effects, two-temperature model, and a real equation of state are considered. Numerical results of plasma flow simulation in a cylindrical lab-scale thruster, with mass flow rate of 6g/s and total discharge current of 8kA, are presented and comparison with experimental data shows good agreement between the predicted and measured contours of enclosed current and electric potential. The estimated thrust is 16.34N which exhibits less than 5% difference compared with measured value. Furthermore, this simulation properly predicts the experimentally observed argon jet structure.

Plasma-induced magnetic field in a MPD thruster and interaction with a magnetic nozzle

Plasma-induced magnetic field in a MPD thruster and interaction with a magnetic nozzle

MHD Simulation and Thermal Design of an MPD Thruster

MHD Simulation and Thermal Design of an MPD Thruster

Numerical Analysis of a Flow in an MPD Thruster Using One-Fluid Model

Numerical Analysis of a Flow in an MPD Thruster Using One-Fluid Model

Effect of magnetic Laval nozzle on a MPD Thruster

Effect of magnetic Laval nozzle on a MPD Thruster

A thrust formula for an MPD thruster with applied-magnetic field

In this paper a thrust formula for applied field MPD will be presented. The swirling velocity will be derived from the magnetic stress tensor and its conversion into axial energy into the magnetic nozzle will be analytically treated. The theoretical prediction of both swirling velocity and thrust will be compared to the measurements showing a reasonable agreement.

Numerical investigations of a repetitively pulsed MPD thruster

AbstractPerformance of a repetitively pulsed self‐field MPD thruster consisting of a flared anode and a short cathode was numerically investigated. For a time‐averaged input power of 100 kW and argon propellant, thrust and thrust efficiency were surveyed for various discharge frequencies under the condition of continuous propellant injection and a constant discharge duty ratio of 0.5. The results suggest that thrust efficiency of repetitively pulsed discharge operation can be improved up to about 16% with increasing discharge frequency owing to a relatively high time‐averaged electromagnetic thrust and an increase in aerodynamic thrust, and can surpass that of continuous discharge operation above 50 kHz. It is also shown that the frozen flow loss can be reduced by increasing discharge frequency, because the input power needed for reionization of the propellant can be suppressed. © 2012 Institute of Electrical Engineers of Japan. Published by John Wiley & Sons, Inc.