Magnetoplasmadynamic Thruster Research Articles

Experimental investigation of electric propulsion systems using C12A7 electride hollow cathodes

The development and experimental investigation of two low-power electric propulsion concepts using compact heaterless C12A7 electride (C12A7:e-) hollow cathodes is presented. The first concept represents an electrothermal thruster, in which a cathode discharge is used to heat a gas that is subsequently accelerated in a nozzle-shaped anode. The second propulsion system is an attempt to develop a sub-500 W magnetoplasmadynamic thruster (MPDT) that uses a rectangular discharge channel that allows to increase the applied magnetic field and thus lower the necessary discharge current. Extensive parameter studies with both concepts were conducted, and the thrust and discharge properties of different geometric and operational configurations were determined. This work is a follow-up publication of a previous paper (Gondol and Tajmar in CEAS Space J 14:65–77, 2021).

Thrust estimation for HTS-magnet based Magneto Plasma Dynamic Thrusters (MPDT)

At present time, electric propulsion is being considered for outer deep space missions. Magneto Plasma Dynamic Thrusters are capable of accelerating semi- neutral plasma gas (Xenon, Iodine, Argon, Ammonia, or Lithium) to a high exhaust velocity using combination of RF power & high magnetic fields. High Temperature Superconducting (HTS) coils can generate high magnetic field due to high operating current density at temperatures below its critical temperature(Tc). These superconducting coils can be cooled by using the cold of outer dark space. The use of 2nd Generation (2G) HTS coils can increase thrust, by increasing high magnetic field with very low power consumption. Thus, specific impulse due to the combination of RF plasma and HTS based superconducting high magnetic field is increased drastically. In this paper, the thrust calculation are carried out for MPDT for varying magnetic fields produced by superconducting coils. The estimation of exhaust velocity and thrust generated by plasma for an MPDT is of great challenge and is required to determine for calculating the life of mission as well as the distance of travel. These estimations are required to support the indigenous development of MPDT.

Simulation of magnetoplasmadynamic process with applied magnetic field

The magnetoplasmadynamic thruster is a typical representative of the high-power electric propulsion device, and the magnetoplasmadynamics process is its core operating mechanism. In order to understand the influence of applied magnetic field on its operating characteristics, the particle-in-cell particle simulation method combined with the scale model based on the self-similarity criterion is used to simulate the operating process of magnetoplasmadynamic thruster with applied magnetic field. The reliability of the model and method are verified by comparing with the experimental results. The plasma characteristic parameter distribution of the thruster during ignition is analyzed, and the influence of external magnetic field and cathode current on the thruster performance are discussed. The research results show that the construction of the discharge arc between the cathode and anode is a key step for thruster ignition and efficient operation. A low-intensity magnetic field is not conducive to the construction of a stable discharge arc, while the plasma beam is concentrated near the axis and the main thrust generation mechanism is the self-field acceleration. The discharge arc between cathode and anode is stable by applying a high magnetic field, and the main mechanism of thrust generation is vortex acceleration, which causes the thrust and specific impulse to increase linearly with the strength of the external magnetic field. The efficiency of the thruster increases with cathode current and the applied magnetic field intensity increasing. The discharge voltage increases with the augment of cathode current, but first decreases and then increases with applied magnetic field intensity increasing.

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.

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.

Integrated study on the comprehensive magnetic-field configuration performance in the 150\xa0kW superconducting magnetoplasmadynamic thruster

Higher magnetic fields are always favoured in the magnetoplasmadynamic thruster (MPDT) due to its superior control of the plasma profile and acceleration process. This paper introduces the world's first integrated study on the 150 kW level AF-MPDT equipped with a superconductive coil. A completely new way of using superconducting magnet technology to confine plasma with high energy and extremely high temperatures is proposed. Using the PIC method of microscopic particle simulation, the plasma magnetic nozzle effect and performance of the MPDT under different magnetic-field conditions were studied. The integrated experiment used demonstrated that, in conjunction with the superconducting coil, greater homogeneity and a stronger magnetic field not only caused more even cathode ablation and improved its lifespan but also improved the performance of the MPDT (maximum thrust was 4 N at 150 kW, 0.56 T). Maximum thrust efficiency reached 76.6% and the specific impulse reached 5714 s.

A plasma equilibrium model for rapid estimation of SF-MPDT performance

In this work, a zero-dimensional plasma model for self-field magnetoplasmadynamic thrusters (SF-MPDTs) is proposed, which is based on the ion-number balance equation and energy balance equation, and can calculate the average electron temperature and the average ion temperature inside the discharge chamber conveniently. At the same time, the model can also predict the thruster performance, and the thruster performance predicted by the model is compared with experimental results, which proved the reliability of the model.

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.

Revolutionizing Spaceflight: A Study on Electric Propulsion and Air-Breathing MPDs

Modern liquid-fuel rocket propulsion harbors a number of great limitations. Among those is the weight of fuel, which makes up more than 90% of the mass of the SpaceX Falcon 9 (NASA, 2018). Electric propulsion has been used for decades as an alternative to liquid-fuel rockets due to low propellant requirements and high specific impulse. Although electric thrusters have strictly been used in non-atmospheric conditions, recent innovations attempt to expand its use to airspace. This quasi-experimental study focuses on the creation of an air-breathing magnetoplasmadynamic (MPD) thruster, with attempts being made to maximize the efficiency of the engine. Immense safety concerns prevented testing from occurring after the engine was built. However, the estimated performance of the built MPD is compared to a multitude of existing forms of electric propulsion, from Hall-effect thrusters to electrodynamic tethers. The concluding evidence suggests that air-breathing MPDs are not currently viable, high-power photon thrusters being of greater use in atmospheric conditions. Further research focusing on decreasing atmospheric breakdown voltage and increasing mirror reflectance of photon thrusters is suggested.

Design and demonstration of a micro air-fed magnetoplasmadynamic thruster for small satellites

This paper developed and tested a micro air-fed magnetoplasmadynamic thruster for small satellites. The thruster is a type of miniaturized electric propulsion system designed base on conventional full scale magnetoplasmadynamic thruster that operates at hundreds of kilowatts power. The thruster is designed and tested using normal air as the propellant under the pulse operation mode on a calibrated micro-force measurement thruster stand. The experiments revealed that the thruster can generate a 34.534 μNs impulse bit with an average power input of 1.857 ± 0.0679 W and thrust to power ratio of 8.266 μN/W, and the specific impulse is calculated to be 2319 s with the thruster efficiency of 9.402%, which is quite competitive comparing with other solid-state and liquid-fed pulse-mode thrusters. This paper presents the design and test results for the thruster under a low power level as well as an analysis of its problems and limitations with corresponding future research and optimization directions noted at the end.

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.

MPD Thruster Technology and Its Limitations

It was realized earlier that chemical propulsion systems utilize fuel very inefficiently, which greatly limits their lifespan. Electric propulsion is into existence to overcome this limitation of chemical propulsion. The magnetoplasmadynamic (MPD) thruster is presently the most powerful form of electromagnetic propulsion. It is the thruster’s ability to efficiently convert MW of electric power into thrust which gives this technology a potential to perform several orbital as well as deep space missions. MPD thruster offers distinct advantages over conventional types of propulsion for several mission applications with its high specific impulse and exhaust velocities. However, MPD thruster has limitations which limits its operational efficiency and lifetime. In this paper, the thruster limitations are reviewed with respect to three operational limits i.e., the onset phenomenon, cathode lifetime, and thruster overfed limits. The dependence and effects of the operational limits on each other is compared using different empirical models to derive a scaling factor that has been found for each geometrical arrangement; a limiting value exists beyond which the operation becomes highly unsteady and electrode erosion occurs. Along with reviewing and proposing methods to overcome power limitations for MPD thrusters, the relation between exit velocity and ratio of electrode’s radius is also verified using Maecker’s formula.

Magnetoplasmadynamic thruster with an applied field based on the second generation high-temperature superconductors

The present work reports the results of the magnetoplasmadynamic (MPD) thruster with an applied field (AF) based on the second-generation high-temperature superconductor (2G HTS) tape functioning research. Achieved thrust, specific impulse, and efficiency were investigated. Experiments were performed on the 25-kW magnetoplasmadynamic thruster model with two types of 10-mm and 17-mm cathodes. The investigation tests were performed with the goal of studying the efficiency of using 2G HTS tape as electromagnetic coils creating applied field for MPD thruster. The used as the applied magnetic field for thruster 2G HTS coil creates magnetic field up to 1 Tesla. The results showed that the thrust and specific impulse with HTSC applied field increases up to 300% and efficiency increases up to 700%. The best achieved results are following: thrust up to 850 mN, specific impulse up to 3840 s, and efficiency up to 54%. Created MPD thruster with HTSC applied field has a resource of 5400 s work total with all the cathodes, more than 150 launches, and single continuous launch up to 140 seconds. The experiments performed show the high efficiency of using the 2G HTS tape as the material for the electromagnetic coils creating applied field for magnetoplasmadynamic thrusters. Strong applied field up to 1 T electromagnetic coils were maid of the 2G HTS tape by SuperOx own production facilities.

Discharge characteristics of two-stage micro-cathode arc MPD thrusters with a permanent magnet and a pulsed magnetic field

Small, lightweight low-power micro-cathode arc thrusters (µCATs) with micronewton thrust are well suited to the altitude control of small satellites like CubeSats. For some applications (orbit raising, maneuvering) their thrust level needs to be improved. A possible approach for this could be the two-staged thruster—a micro-cathode thruster with a magnetoplasmadynamic (MPD) stage and an external magnetic field. In this article, we discuss some discharge features that such a two-stage µCAT-MPD experiences in each configuration of the magnetic field—formed with either a permanent magnet, or a pulsed magnetic coil. We found that in both configurations of the magnetic field, the thrust can be enhanced significantly (up to factor of 10) after some threshold voltage is applied to the second stage. The pulsed magnetic coil ensures better controllability of the magnetic field; however, it causes an undesirable time delay between the plasma generation moments in both stages, which consequently results in a moderate thrust increase. The permanent magnet provides a stable thrust increase; however, it cannot be switched off, which seems to be impractical for its use in micro-satellites. In both magnetic field configurations, the emissive electromagnetic noise level was found to be low-frequency (within tens of kHz) and quite moderate in amplitude, and mechanical noise was found to be two orders of magnitude lower than the thrust generated in the normal working regime.

Design and experimental results of a laser-ignited solid-propellant-fed magnetoplasmadynamic thruster.

A laser-ignited solid-propellant-fed magnetoplasmadynamic thruster (LISMPDT) is developed and tested. In this kind of thruster, a solid propellant is placed in a hollow cathode tube. When it works, the propellant is ablated and partially ionized by the laser; then, the plasma flow enters the chamber and is accelerated by the electromagnetic interaction. Unlike the typical magnetoplasmadynamic thrusters, almost any solid material can be used as a propellant in the LISMPDT; no tanks, valves, or piping systems are required for the propulsion system. Moreover, as the particles involved in the discharge process are produced from the laser ablation process, it would lead to a directed initial velocity and higher ionization. Meanwhile, a part of the energy serving the thruster could also be derived from the laser quasi as an add-on to the discharge energy with the separation of the ignition power and the ignition power unit from the thruster system. Experiments on the discharge characteristics and thrust performances were conducted and analyzed. The results validated the feasibility of the LISMPDT, and the maximal specific impulse reached ∼780 s, while the maximal thrust efficiency was expected to be about 9% under different operating conditions. It was also found that increasing the initial discharge voltage would effectively improve the propulsion performance while a longer laser pulse width led to reduced performance.

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.

Pulse-width variation of power supply for evaluating quasi-steady state of magneto-plasma-dynamic thruster operation

Magnetoplasmadynamic (MPD) thrusters are operated in a quasisteady state with about 1.0 ms pulse created by a pulse forming network (PFN). However, there is still no precedent to verify the operation time quantitatively. The nonsteady region of the pulse can lead to an error of the thrust performance against that of steady state operation. In addition, the propellant gas outside the discharge chamber can be consumed since the exhaust velocity exceeds the estimated velocity. This paper shows the first step in quantitative evaluation of the quasisteadiness of an MPD thruster operation. First, we developed a new power supply that outputs a flat-topped and less nonsteady region pulse with a variable pulse width. Compared with that of a PFN, the nonsteady region “tr + tf” decreased from 0.532 to 0.110 ms. By implementing the circuit shorter and adjusting the gate resistance, the surge voltage in the experiment was suppressed to 309 V, which is less than 2% error of that in the PSIM simulation, 305 V. Second, we operated an MPD thruster using the new power supply for discharge and the external magnetic field. As a result, we obtained operation time characteristics by sweeping the operation time from 0.3 to 5.0 ms. The current waveforms are in a range of 620 ± 70 A. We confirmed the consistency of the thrust, 0.32 ± 0.03 N. From the correlation between the input energy and the impulse, it is possible to discuss quasisteadiness of the MPD thruster operation using the determination coefficient and the offset.