Applied-field Magnetoplasmadynamic Thrusters Research Articles

Observation of fast Alfvén wave in KAERI plasma beam irradiation facility using an applied-field MPD thruster

Electromagnetic fluctuations between 300 and 500 kHz were observed in both ion saturation current and magnetic field measurements in the KAERI plasma beam irradiation facility where the applied-field magnetoplasmadynamic (MPD) thruster is used as a plasma source. In this study, the plasmas were ignited with deuterium gas and the plasma current (input power) was varied from 55 A (2 kW) to 155 A (6 kW). The electron temperature and density measured downstream of the MPD thruster plasma were ∼5 eV and (2–4) × 1012 cm−3, respectively. To identify the fluctuations, the dispersion relation and the phase velocity were experimentally obtained using the Beall method and were compared to the theoretical values under warm plasma conditions. It was revealed that the observed fluctuations were fast Alfvén waves with a circularly polarized magnetic field. In addition to the fast Alfvén waves, we also observed electrostatic fluctuations at a higher frequency that was similar to the lower-hybrid frequency.

Exploring the mechanism of axial ion energy acquisition and loss in an magnetoplasmadynamic thruster plume

In order to investigate the mechanism of axial ion energy acquisition and loss in an applied-field magnetoplasmadynamic thruster (AF-MPDT) plume plasma, a 10 kW level MPDT was tested using a Gauss meter, a Langmuir probe, a Faraday probe, a retarding potential analyzer (RPA) and a thrust target stand. Axial ion energy (Eiz) and axial Lorentz force density (Fz) distributions are obtained and interconnected. In the R direction, a local peak in Eiz exists as a result of the presence of maximum Fz in proximity. This localized Fz surge is instigated by the diamagnetic current, a consequence of the abrupt density reduction near the beam edge. In the Z direction, Eiz first increases to a maximum and subsequently decreases. The increase is due to the dominant Fz in ion acceleration in the vicinity of the thruster exit. Further downstream, the Fz decreased and the effect of charge-exchange collisions became apparent, leading to the decrease of Eiz.

Modelling the Quench Behavior of an NI HTS Applied-Field Module for a Magnetoplasmadynamic Thruster Undergoing a 1kW Discharge

Recent advances in commercial miniaturised cryocoolers and high-temperature superconductors (HTS) have revived the discussion of using HTS electromagnets to enhance the thrust and efficiency of electric thrusters for space applications. <p xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">An HTS applied-field magnetoplasmadynamic (AF-MPD) thruster is currently being developed. While the thruster is operating, there will be a large time-variable heat load on the cryogenic environment. The operation of a low-power cryocooler and energised HTS coils (operating at 70 K) adjacent to streams of hot plasma and large electrical discharges (on the order of 1 kW) represents a significant thermal management problem. The electromagnetic and thermal behaviour of non-insulated (NI) coils under these conditions, and how resilient they are to quenching during thruster operation, is not well understood. <p xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">In this paper, a model is formulated to study the transient electromagnetic and thermal behaviour of an HTS-AF-MPD thruster with NI coils. The thruster and a conduction cooled cryogenic design are coupled via surface-to-surface radiation heat transfer. The model predicts current flow within the HTS, copper stabiliser and between turns, with the contact resistivity being a key input variable. Critical current is determined locally using temperature, magnetic field, and field angle in combination with a measured data set for a specific conductor. <p xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">This model reveals conditions where the cryocooler can passively compensate for large instantaneous heat loads on the coils, demonstrating quench resistance.

Plume deflection using an asymmetrical magnetic nozzle for an applied-field magnetoplasmadynamic thruster—An experimental demonstration and numerical analysis

The plume deflection of an applied-field magnetoplasmadynamic thruster is demonstrated using an asymmetrical magnetic nozzle. The two-dimensional plume diagnoses are performed downstream the deflected plume, where the plasma density, electron temperature, plasma potential, ion current density and the external current are measured by the Langmuir single probe, Faraday probe and a water-cooled Hall probe. A maximum plume deflecting angle of 12°is achieved. And a high-electric-potential region named the deflection arm is found on the anti-deflection side of the thruster axis. To obtain more kinetic insights, a fully-kinetic planar particle-in-cell simulation is also conducted under similar operating conditions, and the general distributions of the properties are reproduced qualitatively. Related closely to the deflection arm region, the E×B-induced Lorentz force is found to be one of the dominant deflecting force under ion non-magnetization condition, besides the Lorentz force induced by the diamagnetic current. As for the other contributors, the inertia-induced and the stress-induced Lorentz forces are rather insignificant.

Analytical investigation of outflow currents in applied field magnetoplasmadynamic thrusters

An applied field magnetoplasmadynamic thruster (AF-MPDT) is one of the most promising electric engines for the space missions with high thrust. However, the acceleration mechanisms in AF-MPDTs are complicated because of the application of the magnetic nozzle, which limits the performance improvement of the thrusters. The discharge currents may flow out of the discharge chamber because of the magnetic nozzle, which was neglected in previous research studies about the acceleration mechanisms in AF-MPDTs. A model based on the MHD equations has been established to investigate the contribution of outflow currents to the plasma acceleration. The accuracy of the model has been proved by the comparison of the calculated thrust and the experimental data. The performance of the argon-fed AF-MPDT is measured by a target thrust stand. The dominant acceleration mechanism of the AF-MPDTs is the plasma expansion in the magnetic nozzle when the magnetic fields are less than 0.2 T. The primary thermal energy source of the AF-MPDTs is the Joule heating when the magnetic fields are lower than 0.1 T. The viscous heating due to the outflow currents will be dominant when the magnetic fields are higher. An AF-MPDT without a discharge chamber is proposed to improve the performance. Compared with the conventional AF-MPDT, the new structure may provide an average 15% increase in performance at the same condition. • The influences of outflow currents on plasma acceleration are studied. • Higher outflow currents are beneficial for plasma acceleration in AF-MPDTs. • The plasma expansion is the dominant acceleration mechanism in AF-MPDTs. • An AF-MPDT without a discharge chamber is proposed for performance improvement.

The azimuthal currents in the ion-driven magnetic nozzle

Ion-driven magnetic nozzles (Ti > Te) are designed as intrinsic parts of cutting-edge propulsive technologies such as variable specific impulse magnetoplasma rockets (VASIMRs) and applied-field magnetoplasmadynamic thrusters. Employing a two-dimensional axisymmetric particle-in-cell (PIC) code, in the ion-driven magnetic nozzle, the compositions and distributions of azimuthal currents in different axial regions are investigated under various inlet ion temperatures Ti0 and found to differ dramatically from that in the electron-driven magnetic nozzles. Previously reported to be all paramagnetic and vanishing under a high magnetic field, the azimuthal currents resulting from the E × B drift are shown to turn diamagnetic and sustain a considerable magnitude when Ti0 is considered. The previously reported profile of diamagnetic drift current is altered by the introduction of inlet ion temperature, and the paramagnetic part is significantly suppressed. Moreover, a wide range of paramagnetic currents appear downstream due to the inward detachment of ions, which can also be reduced by increasing inlet ion temperature. Albeit considered in this paper, the azimuthal currents resulting from grad-B and curvature drift are still negligible in all cases of interest. The magnitude of diamagnetic azimuthal currents increases with amplifying Ti0, indicating a clear physical image of energy transformation from ion thermal energy to the directed kinetic energy through electromagnetic processes in the magnetic nozzle. Additionally, the magnetic inductive strength also has noticeable impacts on the azimuthal currents, the current magnitude tends to decrease as the magnetic field increases, and over-increment of it may result in larger divergence angles and lower nozzle efficiency.

Development and testing of a low-current applied-field magnetoplasmadynamic thruster with a rectangular discharge channel

This study explores the possibility of miniaturizing magnetoplasmadynamic thrusters (MPDTs) to significantly lower power and discharge current levels compared to most conventional MPDTs. A design alternative for MPDTs using a discharge channel with a rectangular cross-section is presented that enables the implementation of strong external magnetic fields to increase the applied-field Lorentz force. The thruster concept uses heaterless calcium aluminate electride (C12A7:e-) hollow cathodes as the electron source. A prototype of the concept intended for the low-amp current range generates thrust in the low millinewton range with a specific impulse ranging between 400 s and 1200 s at power levels below 500 W but shows high thermal power losses to the anode. A further miniaturized version of the concept intended for the sub-amp current range is thermally more sustainable but requires high mass flow rates to achieve a stable discharge, limiting the achievable specific impulse.

High-temperature superconductor-based power and propulsion system architectures as enablers for high power missions

The increasing competitiveness of electric propulsion systems (EPS) for primary spacecraft propulsion has paved the way for higher payload mass fractions by offering significantly higher specific impulses than chemical systems. Concurrently, High-Temperature Superconductors (HTS) have reached an unprecedented level of industrial maturity in recent years, and considering their low masses, compactness, and high current densities, they offer the potential to act as a disruptive technology in several spaceflight applications such as power management systems, re-entry and radiation shielding as investigated in the EU-funded MEESST project, as well as EPS. In the latter case, efforts are already ongoing to develop an HTS-enhanced Applied-Field Magnetoplasmadynamic (AF-MPD) thrusters for high power mission applications. The Tsiolkovsky equation infers that the payload mass fraction increases indefinitely with increased Specific Impulse (Isp), however, in the case of electric propulsion, the dependence of thrust on the available power complicates this issue when transfer time is a primary driver. Here, the Tsiolkovsky equation becomes inadequate and considering a non-dimensional version of the Tsiolkovsky equation in terms of the mission ΔV and transfer time, the EPS thrust efficiency, and the specific mass of the power system becomes necessary. This paper first discusses the recent advances in HTS and their suitability for spaceflight, before reviewing The development of power system technologies is reviewed and a conceptual power system architecture incorporating HTS is presented. These technologies are analysed using a non-dimensional Tsiolkovsky approach and their impacts on the overall payload mass fraction are assessed. For high-power missions (>100 kW), the use of HTS is shown to have a highly beneficial impact on the mass of the power system. Correspondingly, this enables higher payload mass fractions achievable at increased specific impulse operation, thus strengthening the case for high-power, high-Isp EPS technologies such as AF-MPDT.

Collisional radiative model for high-ionization-rate equilibrium argon plasma plume

An extensive collisional radiative model for the neutral argon and singly ionized argon plasma is developed to estimate the key properties of the high-ionization-rate equilibrium plasma plume. Major thirty-one argon atomic levels and ninety-five singly ionized argon levels are considered. The processes taken into account are radiative decay, electron impact ionization, excitation and deexcitation between all the levels, fast ion and atom impact excitation, and three body recombination. The model in conjunction with optical emission spectroscopy could be used to evaluate the electron temperature Te, the electron density ne, atom density ng and to investigate the dynamics of excitation population density. The local plasma spectrum in the plume region of an applied-field magnetoplasmadynamic thruster was measured with an optical immersion probe. The fundamental physical processes leading to W-shaped compound light variation in plume zone are explained based on the above plasma parameter distributions. A satisfactory agreement with the results measured by a Langmuir probe verifies the model's applicability to argon plasma plume.

Anode power deposition in applied field magnetoplasmadynamic thruster

Applied field magnetoplasmadynamic thrusters (AF-MPDTs) are one of the potential thrusters for high power space propulsion. However, the high anode power deposition limits the improvement in the thruster’s efficiency. According to previous research studies, the anode voltage drop is supposed to be the main factor of the anode power deposition in AF-MPDTs without considering the effect of the centrifugal force. Actually, the centrifugal force cannot be neglected when the magnetic field is high, where the anode voltage drop is not the only dominant factor of the anode power deposition. Due to the significant centrifugal force, there might be no obvious shortage of charge carriers near the anode surface in AF-MPDTs with propellant injection between the cathode and anode. The electrons will be compressed to the anode surface instead of the cathode. In order to investigate the reason for the anode power deposition in AF-MPDTs, a model considering the effect of the centrifugal force is established. The calculated anode power deposition shows good agreement with the corresponding experimental data. In addition, when the magnetic field is high, the viscous heating cannot be neglected, which is as significant as the anode sheath heating and Joule heating. Finally, in order to reduce the anode power deposition, a magnetic field distribution with a high value near the cathode and a low value near the anode is proposed. According to the evaluation by the model, the anode power deposition can be reduced by 30% via the new magnetic field when the magnetic field exceeds 0.1 T.

Simulation of the plume of a magnetically enhanced plasma thruster with SPIS

A three-dimensional fully kinetic particle-in-cell (PIC) simulation strategy has been implemented to simulate the acceleration stage of a magnetically enhanced plasma thruster (MEPT). The study has been performed with the open-source code Spacecraft Plasma Interaction Software (SPIS). The tool has been copiously modified to simulate properly the dynamics of a magnetized plasma plume. A cross-validation of the methodology has been done with Starfish, a two-dimensional open-source PIC software. Two configurations have been compared: (i) in the absence of a magnetic field and (ii) in the presence of a magnetic field generated by a coil with maximum intensity of 300 G at the thruster outlet. The results show a reduction of the plume divergence angle, an increase of ion speed and an increase of the specific impulse in the presence of the magnetic nozzle. The simulations presented in this study are representative of the operative conditions of a 50 W MEPT. Nonetheless, the methodology adopted can be extended to handle the magnetized plasma plume of several other types of thrusters such as electron cyclotron resonance and applied field magnetoplasmadynamic thrusters.

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.

Fabrication and Testing of Applied-Field Magneto Plasma Dynamic Thruster with Argon Propellant

The experimental setup of applied-field Magneto Plasma Dynamic Thruster (MPDT) is a form of electrically powered spacecraft propulsion system. The thrusters are used for deep space mission, which consist of long duration, low thrust acceleration and less propellant storage. MPDT uses the force on the charged particle by an electromagnetic field to generate thrust. The technology takes high discharge current for two electrodes and solenoid coil as applied-field. In the present investigation, the high discharge current is replaced by a limited discharge current, it is reduced by the use of capacitors bank to discharge the current for two electrodes. The solenoid coils are replaced by permanent magnet as applied-field to generate the thrust. The basic requirements for MPDT are the power circuit made by capacitors bank of electrolytic type. The electrolytic capacitors are smaller in size and produces limited voltage, which will support limited current flowing between the two electrodes. The material used for anode and cathode are copper and tungsten respectively. A permanent magnet is fixed over the anode coaxially and cathode is placed in center. A propellant will be injected in between the two electrodes. The propellant gas used in this thruster is argon. The performance of the thruster is tested in vacuum condition and the thrust value is measured by a load cell. The prototype of MPDT is tested and measurement of discharge current, thrust, specific impulse and efficiency are carried out. The thrust for various discharge current is measured and the efficient operating conditions of MPDT can be obtained.

High-spatial-resolution image reconstruction-based method for measuring electron temperature and density of the very near field of an applied-field magnetoplasmadynamic thruster

A method based on image reconstruction is proposed for measuring the electron temperature () and density () distribution of the axisymmetric low-temperature argon plasma plume of applied-field magnetoplasmadynamic thrusters (AF-MPDTs). The method uses a complementary metal–oxide–semiconductor camera with narrowband filters of 460 and 500 nm to obtain the raw projection images in specific wavelength ranges. An image reconstruction method is developed to construct relative emission intensity profiles. The electron temperature distribution can be obtained by combining the relative emission intensity ratio of two profiles with a simple argon kinetic model. The electron density distribution can be calculated by constructing a function of relative emission intensity containing and . To verify the feasibility of this method, we diagnosed a very near field plume of a 15 kW AF-MPDT and obtained the distribution of parameters with a spatial resolution of 0.23 mm. The results show that the maximum and are located on the centerline near the exit plane of the thruster and decrease with both radial and axial distances. The second peak appears in the radial direction of both the and distributions, while their zones do not overlap. A comparative analysis of measurement results obtained by this method and the Langmuir probe method is performed to prove its accuracy. Moreover, this measurement method has the advantages of high spatial resolution and the ability to diagnose high-density and magnetized plasma. This method can be also applied to other optically thin axisymmetric low-temperature argon plasmas.

Cathode erosion site distributions in an applied-field magnetoplasmadynamic thruster

Erosion can influence cathode life, and is thus considered to be one of the main factors limiting the application of applied-field magnetoplasmadynamic thrusters. In this paper, erosion sites on graphite cathodes are studied so as to identify the influence of applied magnetic field and the ratio of propellant mass flow rate supplied from cathode and anode. The experiment results show that the application of applied magnetic field can significantly reduce the erosion rate of the cathode compared to that without magnetic field. The erosion sites on the cathode vary with the relative position of the convergent-divergent magnetic field, and are mainly distributed in the divergent part of the field. The erosion sites on the cathodes are found to be related to the propellant supply. The decreasing anode mass flow rate enlarges the range of erosion. These results are much helpful for the analysis of cathode erosion site location since they provide evidences of erosion mechanisms and point out the directions for further research.

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.

A 100 KW Class Applied-field Magnetoplasmadynamic Thruster.

Applied-field magnetoplasmadynamic thrusters (AF-MPD thrusters) are hybrid accelerators in which electromagnetic and gas dynamic processes accelerate plasma to high speed; they have considerable potential for future space applications with the significant advantages of high specific impulse and thrust density. In this paper, we present a series of protocols for designing and manufacturing a 100 kW class of AF-MPD thruster with water-cooling structures, a 130 V maximum discharge voltage, a 800 A maximum discharge current, and a 0.25 T maximum strength of magnetic field. A hollow tantalum tungsten cathode acts as the only propellant inlet to inhibit the radial discharge, and it is positioned axially at the rear of the anode in order to relieve anode starvation. A cylindrical divergent copper anode is employed to decrease anode power deposition, where the length has been reduced to decrease the wall-plasma connecting area. Experiments utilized a vacuum system that can achieve a working vacuum of 0.01 Pa for a total propellant mass flow rate lower than 40 mg/s and a target thrust stand. The thruster tests were carried out to measure the effects of the working parameters such as propellant flow rates, the discharge current, and the strength of applied magnetic field on the performance and to allow appropriate analysis. The thruster could be operated continuously for significant periods of time with little erosion on the hollow cathode surface. The maximum power of the thruster is 100 kW, and the performance of this water-cooled configuration is comparable with that of thrusters reported in the literature.

Performance investigation of an argon fueled magnetoplasmadynamic thruster with applied magnetic field

In the present paper, a density-based central-upwind magnetohydrodynamic (MHD) code has been used to get insight into the acceleration mechanism of the applied-field magnetoplasmadynamic (MPD) thrusters. The magnetic field is axially applied by an external coil surrounding the anode and interacts with the discharge current and the induced azimuthal current to produce thrust. In the present work, the numerical modeling of applied-field magnetoplasmadynamic thrusters is performed with a separate magnetostatic code to produce external magnetic field from permanent magnets, and the density-based method is used to compute the resulting flow field from the MHD equations. The numerical model is applied to the NASA Lewis Research Center 100-kW magnetoplasmadynamic (MPD) thruster which is experimentally and numerically well documented to demonstrate its capability to capture the main characteristics of plasma acceleration and thrust production in such a device. The code is then used to investigate the thruster performance operating in the applied magnetic field strength range 10–100 mT at discharge currents of 750–2000 A with a constant mass flow rate of 0.1 g/s. The effect of the applied magnetic field inside and outside of the thruster is investigated and reported.

Target thrust measurement for applied-field magnetoplasmadynamic thruster

In this paper, we present a flat target thrust stand which is designed to measure the thrust of a steady-state applied-field magnetoplasmadynamic thruster (AF-MPDT). In our experiments we varied target-thruster distances and target size to analyze their influence on the target thrust measurement results. The obtained thrust-distance curves increase to local maximum and then decreases with the increasing distance, which means that the plume of the AF-MPDT can still accelerate outside the thruster exit. The peak positions are related to the target sizes: larger targets can make the peak positions further from the thruster and decrease the measurement errors. To further improve the reliability of measurement results, a thermal equilibrium assumption combined with Knudsen’s cosine law is adapted to analyze the error caused by the back stream of plume particles. Under the assumption, the error caused by particle backflow is no more than 3.6% and the largest difference between the measured thrust and the theoretical thrust is 14%. Moreover, it was verified that target thrust measurement can disturb the working of the AF-MPD thruster, and the influence on the thrust measurement result is no more than 1% in our experiment.

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.