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It has been urgently required to develop ionic liquid electrospray thruster (ILET) for rapidly expanding micro-/nanosatellites. In this work, ILETs with high-density array emitters are fabricated t...

An electrospray thruster designed for nanosatellites is introduced in this paper. The thruster has a compact size featuring a passive propellant transport method through the use of porous materials...

ESA’s Next Generation Gravity Mission (NGGM) is a candidate Mission of Opportunity for ESA-NASA cooperation in the frame of the MAss change and Geosciences International Constellation (MAGIC) . The mission aims at enabling long-term monitoring of the temporal variations of Earth’s gravity field at relatively high temporal (down to 3 days) and increased spatial resolutions (up to 100 km) at longer time. Such variations carry information about mass change induced by the water cycle and the related mass exchange among atmosphere, oceans, cryosphere, and land, and will complete our picture of global and climate change with otherwise partial or unavailable data. Over the last 15 years, numerous system and technology activities have been initiated by the Earth Observation Programmes (EOP) Directorate of the European Space Agency with the aim of advancing the maturity of the NGGM system and the key subsystems: particular attention was devoted to the design of the fine attitude control system, enabled by proportional thruster like variable specific impulse electrostatic thrusters based on Field Emission Electric Propulsion (FEEP), in which a liquid propellant is electrostatically extracted and accelerated to high exhaust velocity. The core element of this propulsion technology is a passively-fed, porous tungsten crown emitter, consisting of 28 sharp needles. This emitter technology has been developed and qualified over more than a decade at FOTEC and the Austrian Institute of Technology under ten ESA/EOP contracts since 2005 targeting the NGGM needs and it has recently been adapted for use as the main propulsion system in commercial nano- and small-satellites. This paper summarizes the development efforts of the last decade and provides an assessment of the performance of this thruster technology, after extensive simulation and testing campaigns.

This paper describes a metal plasma thruster (MPT) that is suited to nano- and microsatellites. The MPT uses solid metal propellant (hence requires no liquids, gases, flow valves, or flow controls)...

International audience; This article provides information about the sizing and standardization of a Faraday cup (FC) used as a plasma diagnostic. This instrument is used to accurately map the ion beam profile produced by an electric propulsion (EP) device. A FC is a cylindrical probe that uses an electrode, termed collector, to measure the current. Several studies have shown the relevance of adding an extra electrode, called collimator, to define the collection area and to minimize interactions with the ambient plasma. Both the electrodes are encapsulated into an isolated metallic housing that prevents ambient plasma from disturbing the measurements. In this case study, a field-emission-electric propulsion (FEEP) thruster is used. The FEEP technology uses electrostatic fields to extract liquid metal (indium) ions from a sharp surface and accelerates them to high velocities, providing thrust. The FEEP model used in this study is the ENPULSION NANO thruster from the Austrian company Enpulsion. We present results focusing on the sizing of a FC in terms of cup length, aperture diameter, and collection solid angle as well as on the material exposure to the ion beam. For a far-field ion beam study of a FEEP indium based electric thruster, our study outcomes show that a FC optimum sizing is a 50 mm long collector cup and a 7 mm wide inlet aperture. Moreover, shielding the repeller/collimator from direct exposure to the ion beam seems to greatly minimize perturbation during ion current acquisition. Finally, to only measure the ion current, a negative potential should be applied to the collector and repeller, where the latter is more negative. This study contributes to the effort on diagnostic standardization for EP device characterization. The goal is to enable repetitive and reliable determination of thruster parameters and performances.

Accurate, direct measurement of thrust or impulse is one of the most critical elements of electric thruster characterization, and one of the most difficult measurements to make. This paper summarizes recommended practices for the design, calibration, and operation of pendulum thrust stands, which are widely recognized as the best approach for measuring μN- to mN-level thrust and μNs-level impulse bits. The fundamentals of pendulum thrust stand operation are reviewed, along with the implementation of hanging pendulum, inverted pendulum, and torsional balance configurations. Methods of calibration and recommendations for calibration processes are presented. Sources of error are identified and methods for data processing and uncertainty analysis are discussed. This review is intended to be the first step toward a recommended practices document to help the community produce high quality thrust measurements.

A highly miniaturized Field Emission Electric Propulsion (FEEP) system is currently under development at TU Dresden, called NanoFEEP [ 1 ]. The highly miniaturized thruster heads are very compact and have a volume of less than 3 cm3 and a weight of less than 6 g each. One thruster is able to generate continuous thrust of up to 8 μN with short term peaks of up to 22 μN. The very compact design and low power consumption (heating power demand between 50 and 150 mW) are achieved by using Gallium as metal propellant with its low melting point of approximately 30 °C. This makes it possible to implement an electric propulsion system consisting of four thruster heads, two neutralizers and the necessary electronics on a 1U CubeSat with its strong limitation in space, weight and available power. Even formation flying of 1U CubeSats using an electric propulsion system is possible with this system, which is shown by the example of a currently planned cooperation project between Wuerzburg University, Zentrum fuer Telematik and TU Dresden. It is planned to use the NanoFEEP electric propulsion system on the UWE (University Wuerzburg Experimental) 1U CubeSat platform [ 2 ] to demonstrate orbit and two axis attitude control with our electric propulsion system NanoFEEP. We present the latest performance characteristics of the NanoFEEP thrusters and the highly miniaturized electronics. Additionally, the concept and the current status of a novel cold neutralizer chip using Carbon Nano Tubes (CNTs) is presented.

The characteristics of the propellant film on the tip surface of the field-emission electric propulsion thruster have a close relationship with the work state of the thruster. In this paper, the distributions of both the liquid film thickness and the fluid velocity along the needle tip have been calculated by building the liquid mechanics equations and a cone coordinate system suitable for the tip. The results show that the film thickness is tens to hundreds of nanometers and the fluid velocity along the needle tip is from tens of µm/s to a few mm/s.

Electric propulsion presents an excellent alternative to conventional chemical counterparts as higher exhaust velocities from a plasma-based thruster allow for more efficient mass utilization and expanded space mission capabilities. Understanding the detailed time and space dependant physical phenomena within the main thruster could potentially enhance the capabilities of the electric thruster. Hence, in the present work a two-stage Hall thruster is investigated and the potential feasibility of coupling the two stages to produce the ion-beam is evaluated. The ionization and acceleration events are investigated via multiphysics simulations. The key trends ranging from the discharge inception, amplification and drift to formation of the ion beam are discussed and characterized.

This paper proposes a coupled analysis method to deal with ion generation and ion drift in the air. Electric discharge is utilized in various products, and it is very important to clarify not only how much ion is generated but also where ion drifts in the air. The proposed method, which modeled interactions between electric field and ion behavior, including ion generation, ion drift, and charging of a dielectric body, has succeeded in calculating charge removal of a metallic target, which was tens of millimeters away from a pin electrode as a source of generated ion. The calculated results show that the amount of generated ion is drastically changed because electric field is affected by electric charge of ion in the air and charging of a dielectric body. This paper also reports a comparison of measured and calculated results of charge removal time of a metallic target, which were found to be in good agreement.

Faraday probes are a common plasma diagnostic used to determine the local ion charge flux of electric propulsion plumes. Standard practices, guidelines, and recommendations are provided for experimental methods and analysis techniques that aim to standardize community practices, to mitigate test environment effects, and to reduce systematic measurement error in order to improve plume predictions in the space environment. The approaches are applicable to time-averaged plasma properties in the near-field and far-field of electric propulsion plumes, with emphasis on Hall effect thrusters and gridded ion thrusters. Considerations for other electric propulsion technologies are provided, including electrosprays, arcjets, and electromagnetic thruster concepts. These test strategies are expected to increase the quality of comparisons between different thrusters and vacuum environments, thereby broadening the applicability of ground-based measurements and enhancing the fidelity for on-orbit predictions and modelin...

The back-sputtering deposition on thruster surface caused by ion bombardment on chamber wall material affects the performance of thrusters during the ground based electric propulsion endurance tests. In order to decrease the back-sputtering deposition, most of vacuum chambers applied in electric propulsion experiments are equipped with anti-sputtering targets. In this paper, a three-dimensional model of plume experimental system (PES) including double layer anti-sputtering target is established. Simulation cases are made to simulate the plasma environment and sputtering effects when an ion thruster is working. The particle in cell (PIC) method and direct simulation Monte Carlo (DSMC) method is used to calculate the velocity and position of particles. Yamamura's model is used to simulate the sputtering process. The distribution of sputtered anti-sputtering target material is presented. The results show that the double layer anti-sputtering target can significantly reduce the deposition on thruster surface. The back-sputtering deposition rates on thruster exit surface for different cases are compared. The chevrons on the secondary target are rearranged to improve its performance. The position of secondary target has relation with the ion beam divergence angle, and the radius of the vacuum chamber. The back-sputtering deposition rate is lower when the secondary target covers the entire ion beam.

Ion thruster plumes are simulated under a framework that was recently developed to take advantage of the large scalability of unstructured grids using adaptive mesh refinement. Momentum and charge-exchange collisions occurring between neutral and ion species, as well as the induced electric field due to ions, have been performed using multiple adaptive mesh refinement meshes to study the backflow of an ion thruster to a spacecraft surface. Furthermore, the ions and neutral species are directly coupled, which is found to have a 15% effect on the neutral species velocity profiles. The backflow for ions is found to occur when charge-exchange and momentum-exchange collisions are present, and then it is strongly enhanced when the induced electric field is considered. The ion energy distributions in the backflow region are obtained, and it is found that the inclusion of the electric field modeling is the most important factor in determining its shape. The plume backflow structure is also examined for a triple-t...

The article is devoted to the research of low-power (300 W) radio-frequency ion thruster designed at the Moscow Aviation Institute. The main results of experimental research of the thruster using the testfacility power supplies and the power processing unit of their own design are presented. The dependence of the working fluid ionization cost on its mass flow rate at the constant ion beam current was investigated experimentally. The influence of the shape and material of the discharge chamber on the integral characteristics of the thruster was studied. The recommendations on the optimization of the thruster primary performance were developed based on the results of experimental studies.

A novel approach to the design and fabrication of field-emission tips with self-aligned gates intended for electric propulsion micro-thruster applications is presented. Their micro-electromechanical systems fabrication process is derived from the recent proliferation of research toward developing field emitter arrays, which are used primarily for field-emission flat-panel display applications. An array of micron-sized tips for electric field enhancement via wet isotropic etching of silicon, using silicon nitride as a hard mask is fabricated. The wet etching is accomplished using a combination of hydrofluoric, nitric, and acetic acids. The tips were then coated with a metal layer to enhance wetting by indium, the proposed propellant. Next, a layer of SU-8 photoresist was applied by spin coating and patterned to serve as a dielectric spacer. A second layer of metal was then applied to serve as a gate electrode. In addition, the results of electrostatic simulations of the prototype is described.

A point-like silver ion source with thin film of RbAg 4 I 5 solid electrolyte was developed for application in field emission electric propulsion systems. The solid electrolyte produced by mechano-chemical synthesis was deposited on the apex of a silver tip using pulsed laser deposition. The running tests at 195 °C working temperature and 10 kV accelerating potential have demonstrated that the source can continuously produce positive ion current of several hundred pA during a few days with a peak value of 25–50 nA. Using TOF-SIMS analysis it has been shown that the ion current consists of Ag + with a small contribution of Rb + , lower than 0.5 at.%. We suggest that Ag + emission occurs mainly by means of field-assisted ion evaporation and/or ion desorption, and the solid electrolyte plays a role of the transport system, in which the emitted Ag + ions are continually replaced by mobile Ag + ions delivered from the silver reservoir by direct mass transfer under the influence of an external electric field.

In this study, the authors analysed an anomalous force observed in asymmetric capacitors, working in high electric potentials. From a lot of experimental measurements performed in their asymmetric capacitor, they detected real variations of the device inertia. An empirical formula of the developed force according to Clausius–Mossotti relation explained the experimental results with good agreement, suggesting that the anomalous force can be related directly to the electric dipoles inside the capacitor dielectric. They could also explain why that force always points toward the small electrode when it is subjected to a convergent electric field. Such simple electrical propulsion systems could allow in the future the substitution of the fuel propulsion technology.

One of the most recently developed modes of propulsion is electric propulsion. The commonly used chemical propulsion systems have the advantage of a high Specific Impulse as compared to that of ion propulsion systems. However, owing to the efficacy of ion propulsion systems, it is considered the future of space exploration.Electro thermal thrusters produce thrust by using electrical fields to force hot plasma out of the nozzle with certain exit velocity. The plasma’s exit velocity and the system’s thrust capacity, as of now, are insufficient for space travel to be conducted within a reasonable time. I intend to study the possibility of improving the thruster’s performance by using an aerospike nozzle as an exit nozzle which meets the conditions required for the thruster to function appropriately. I shall be studying the plasma plume exit velocity variation with respect to the nozzles used. Also, a thermal analysis will be conducted in order to find the correct material for the nozzle.

The Front Surface Acceleration (FSA) obtained in Laser Ion Source (LIS) systems is one of the most interesting methods to produce accelerated protons and ions. We implemented a LIS to study the electric field responsible for the ion acceleration mechanisms. A high impedance resistive probe was used to map the electric potential inside the chamber, near the target. We detected the time resolved profiles of the electric potential moving the probe from 4.7 cm to 6.2 cm with respect to the main target axis. The corresponding electric field depends on the distance x as 1 / x α with α ∼ 1.8 . We suggest that the electric field strength stems from the contribution of an electrostatic and an induced field.

The IFM Nano Thruster, which consists of a porous tungsten crown ion-emitter to provide thrust in the range of 10 µN to 1 mN at 2000–5000 s ISP, has been tested on a µN thrust balance that has previously been verified at the ESA Propulsion Laboratories. A comprehensive assessment of the thrust performance has been complemented by measurements of the plume divergence using a novel segmented collector setup that allows to assess the ion current distribution within a 2D-section of the plume in a high time resolution. From the plasma probe measurements, an upper bound of 0.93 for the thrust coefficient has been derived, while direct thrust measurements have shown a thrust coefficient between 0.8 and 0.9 over the entire thrust range. From the test results, a complete thruster performance map could be established linking emitter current and voltage with thrust and Isp.

Electric propulsion is now a successful method for primary and secondary propulsion of deep space long-duration missions and for geosynchronous satellite attitude control. The Plasma Physics Laboratory of UnB has been developing a permanent magnet Hall thruster (PHALL) for the UNIESPACO program, part of the Brazilian space activities program (PNAE) since 2004. The idea of using an array of permanent magnets, instead of an electromagnet, to produce a radial magnetic field inside the plasma channel of the thruster is very significant. It allows the development of a Hall thruster with power consumption low enough to be used in small- and medium-size satellites. The PHALL project consists on plasma source design, construction and characterization of the Hall-type propulsion engine using several plasma diagnostics sensors. PHALL is based on a plasma source in which a Hall current is generated inside a cylindrical channel with an axial electric field produced by a ring anode and a radial magnetic field produced by permanent magnets. In this work, a brief description of the plasma propulsion engine, its diagnostics instrumentation and measured plasma parameters with a focus for possible applications of PHALL on orbit transfer maneuvering for future Brazilian geostationary satellite space missions is shown. More specifically, we will show plasma density and temperature space profiles inside and outside the thruster channel, ion temperature measurements based on Doppler broadening of spectral lines and ion energy measurements. Based on the measured plasma parameters we construct an aptitude figure of the PHALL. It contains the specific impulse, total thrust, propellant flow rate and power consumption necessary for orbit raising of satellites. Based on previous studies of geosynchronous satellite orbit positioning we perform numerical simulations of satellite orbit raising from an altitude of 700–36,000 km using a PHALL operating in the 100–500 mN thrust range. To perform these calculations integration techniques were used. The main simulation parameters were orbit raising time, fuel mass, total satellite mass, thrust and exhaust velocity. We conclude by comparing our results with results obtained with known space missions performed with Hall thrusters.

To extend the lifetime of commercial and scientific satellites in low Earth orbit (LEO) and below (100–250 km of altitude) recent years showed an increased activity in the field of air-breathing electric propulsion as well as beamed-energy propulsion systems. However, preliminary studies showed that the propellant flow necessary for electrostatic propulsion at these altitudes exceeds the mass intake possible within reasonable limits, and that electrode erosion due to oxygen flow might limit the lifetime of eventual thruster systems. The pulsed plasma thruster (PPT), however, can be successfully operated with smaller mass intake and at relatively low power. This makes it an interesting candidate for air-breathing application in LEO and its feasibility is investigated within this paper. An analysis of such an air-breathing PPT system shows that for altitudes between 150 and 250 km, drag compensation is at least partially feasible assuming a thrust-to-power ratio of 30 mN/kW and a specific impulse of 5000 s. Further, to avoid electrode erosion, inductively heated electrothermal plasma generator technology is discussed to derive a possible propulsion system that can handle gaseous propellant without unfavorable side effects. Current technology can be used to create an estimated 4.4 mN of thrust per 1 mg/s of mass flow rate, which is sufficient to compensate the drag for small satellites in altitudes between 150 and 250 km.

Stationary structures formed on the surface of a liquid under the action of an electric field have been studied. With the use of the equations of electrohydrodynamics, it has been shown that cones are formed in a dynamic regime when the current of surface ions, which is induced by the tangential component of the electric field, dominates over the conductivity current in the bulk of a cone and the surface current induced by the flow of the liquid. The electric field strength both inside and outside the cone, as well as the velocity field inside the cone, has been determined. It has been shown that the angle of the cone depends on the current flowing on its surface. The characteristic size of the cone has been determined.

The Hall effect thruster is an electric propulsion device for space applications that efficiently reduces the propellant mass requirements on a spacecraft in comparison with chemical rockets. To date, the Hall effect thruster technology relies on the thermionic cathodes that consume up to 10% of the total propellant used in the system to neutralize the ion beam of the thruster; however, such propellant usage does not contribute to thrust generation. An array of thin-film, carbon nanotube field emitters that emit electrons through field emission can potentially neutralize the ion beam without consuming any propellant. This paper examines the effects of 40 min exposure of carbon nanotube field emitter arrays to the plasma environment in the exit plane of a Hall effect thruster. The physical structures that enable field emission appear largely unaffected by placement in the plasma as well as operation in the plasma. This indicates that a refined design of this carbon nanotube field emitter array may potentially provide an alternative to the thermionic cathode used on contemporary Hall effect thrusters and verifies that no fundamental incompatibilities exist between these two technologies.

The institute of aerospace engineering at TU Dresden is developing a highly miniaturized field emission electric propulsion (FEEP) system for Nano-satellites, called CubeSats. The presented propulsion system NanoFEEP is composed of the actual field emission thrusters and a carbon nanotubes silicon chip used as a cold electron source for neutralizing the electric charging behavior of the spacecraft. The NanoFEEP thrusters use Gallium as metal propellant and provide a continuous controllable thrust of up to 20 micro-newton. This thrust level is sufficient to control the attitude of the miniature CubeSat satellites and even perform orbit maneuvers. Two satellites missions will demonstrate these capabilities in the near future. One mission will show the feasibility of formation flying of CubeSats. The other mission will demonstrate the possibility of space debris avoidance with our propulsion system by deorbiting the satellite after mission completed.

Performance measurements have been obtained of a novel propulsion concept called the Halo thruster under development within the University of Surrey. The Halo thruster, a type of cusped-field thruster with close similarity to the cylindrical Hall thruster, is motivated by the need for low-power and low-cost electric propulsion for the small satellite sector. Two versions of the device are investigated in this study: a design using permanent magnets at high magnetic-field strength and a design using electromagnets with moderate field strength. While operating at 200 W discharge power, which is of particular interest to power-limited small satellite platforms, the permanent-magnet design achieved a maximum thrust efficiency of 8% at a specific impulse of approximately 900 s using a krypton propellant. By comparison, the electromagnet design achieved a maximum thrust efficiency of 28% at a specific impulse of approximately 1500 s at 200 W using a xenon propellant. For higher levels of power (tested up to 800 W), the performance of the electromagnetic design saturated at approximately 25% thrust efficiency using krypton and 30% using xenon. The thrust efficiency of the permanent-magnet design appeared to increase monotonically up to 600 W reaching a maximum value of 14%.Read More: https://arc.aiaa.org/doi/abs/10.2514/1.B36091

An inexpensive optically isolated nanoammeter, which permits digital logging of dc currents from an electrospray colloid thruster operated in vacuum has been designed for using with micro-Newton electric propulsion technology. The two-stage optically isolated system is intended to safely measure and record a current generated at high voltage, facilitating data acquisition with a grounded logging device such as a personal computer. It operates by converting the current into a frequency in a battery-powered, floated, high-voltage stage. Agilent Technologies fiber-optic kits were selected for their low cost and reliability. The fiber-optic cable passes between the high voltage and the grounded receiver stage via grommeted holes drilled in the metal housing. The use this optically isolated system permits accurate monitoring of micro-Newton thruster currents directly rather than through inaccurate proxies, giving physical insight into the behavior of the system at high accuracy and low cost.

Electromagnetic simulation is performed on two compact electrode designs to obtain higher electric field distribution between high voltage and ground electrode, which ensures high bacterial inactivation in liquid food. Electric field simulations are performed initially for test voltages of 1kV to 5kV to understand the nature of electric field distribution in the inactivation area. Then the applied voltage is gradually increased to induce transmembrane potential on the cell membrane. The microorganism modeled for field analysis is Staphylococcus aureus. For both the round edged and torus tube designs, the transmembrane potential of 0.5V to 1V is achieved with the spherical cell, modeled between the high voltage and the ground electrode, which shows pore formation possibility. Dimensions of the electrodes are maintained in mm and cm suitable for laboratory scale, continuous pulsed electric field treatments. In both the designs, emphasis is on the simulation to achieve higher electric field application in the inactivation area, which can be realized through satisfying electroporation phenomenon. From the observed results, it is understood that greater electric field application is achievable even using small efficient electrode designs, which in turn assures i) A greater bacterial inactivation in the liquid food and ii) A compact pulsed electric field experimental prototype.

Measuring electric field has a very wide range of applications in the power system, but for the ion field under the HVDC transmission lines, the current measurements are insufficient. In this paper, effects of interventional measurement, theoretical analysis, measuring programs are researched. Sensors in a Poisson field sense an electric field Ein, which consists of three parts. E0 is the origin electric field, E1 indicates the field generated by charges attached on the sensor, E2 represents the influence of the redistribution of ion flow after the sensor is put into the field. Based on multi-physics simulation, this paper figures out the effects of sensor's intervention through analysing the movement of charged particles and found that when the charging process is completed, there is a spherical region around the sensor where the charge density is zero, which is called the “empty zone”. This paper figures out the relationship between inside field and outside field. The electric field directly measured by the sensor kernel module can be approximately regarded as the sum of the space field which we aim to measure and the self-generated electric field of the charge accumulated on the metal shell of sensor's shell. A structure of two-MEMS-device is promoted to deal with the effect of the ion flow, and it could applied in the ion field or nominal electric field without ion flow and ignore the angular offsets of placing, according to the simulations.

For better integration of plasma thrusters into spacecraft and satellites, potential damaging effects of impinging high energy ions on the spacecraft surfaces should be taken into consideration. For analyzing the plume plasma in this regard, retarding potential analyzers (RPA) are used in electric propulsion as one of the most fundamental and widely used diagnostics tools. RPA determines the ion energy distribution of plume plasma which is in the downstream of a thruster. This paper reports on a successful design process of an RPA to be used with BURFIT-80 ion thruster. An in-house developed hybrid Particle-In-Cell Direct Simulation Monte Carlo (PIC-DSMC) code, which is previously applied for ion thruster grid region is implemented to simulate the flow in the RPA. Results are used to validate the effective operation of the RPA. Potential distributions in the RPA are investigated and the calculated currents collected by the collector is compared with an existing experimental study.

High current liquid metal ion sources are well known and found their first application as field emission electric propulsion thrusters in space technology. The aim of this work is the adaption of such kind of sources in broad ion beam technology. Surface patterning based on self-organized nano-structures on, e.g., semiconductor materials formed by heavy mono- or polyatomic ion irradiation from liquid metal (alloy) ion sources (LMAISs) is a very promising technique. LMAISs are nearly the only type of sources delivering polyatomic ions from about half of the periodic table elements. To overcome the lack of only very small treated areas by applying a focused ion beam equipped with such sources, the technology taken from space propulsion systems was transferred into a large single-end ion implanter. The main component is an ion beam injector based on high current LMAISs combined with suited ion optics allocating ion currents in the μA range in a nearly parallel beam of a few mm in diameter. Different types of LMAIS (needle, porous emitter, and capillary) are presented and characterized. The ion beam injector design is specified as well as the implementation of this module into a 200 kV high current ion implanter operating at the HZDR Ion Beam Center. Finally, the obtained results of large area surface modification of Ge using polyatomic Bi2+ ions at room temperature from a GaBi capillary LMAIS will be presented and discussed.

Ion acceleration through a slowly diverging magnetic nozzle between a ring anode and a hollow cathode set on the axis of symmetry has been realized. Xenon was supplied as the propellant gas from an annular slit along the inner surface of the ring anode so that it was ionized near the anode, and the applied electric potential was efficiently transformed to an ion kinetic energy. As an electrostatic thruster, within the examined operation conditions, the thrust, F, almost scaled with the propellant mass flow rate; the discharge current, Jd, increased with the discharge voltage, Vd. An important characteristic was that the thrust also exhibited electromagnetic acceleration performance, i.e., the so-called “swirl acceleration,” in which F≅JdBRa ∕2, where B and Ra were a magnetic field and an anode inner radius, respectively. Such a unique thruster performance combining both electrostatic and electromagnetic accelerations is expected to be useful as another option for in-space electric propulsion in its broad functional diversity.

The molten contact surfaces, which are caused by interrupting high currents in a vacuum interrupter (VI), can be induced to form Taylor cones by electrostatic force after current zero. Moreover, the deformation caused by the Taylor cones can further enhance the local electric field at their tips, which lead to breakdowns during post-arc phases, and make the interruptions fail. The objective of this paper is to simulate how the Taylor cones are brought out, considering influences of initial shapes, electric fields. An electrohydrodynamic (EHD) model was adopted, and the interface between vacuum and liquid metal region was traced by the level-set method. The results show that the presence of a protuberance indeed can transform to a local field enhancing Taylor cone if the liquid metal surface is exposed to a large enough electric field. Besides, the threshold for Taylor cone formation decreases when temperature of electrode increases. Consequently, the possibility of breakdown in a VI after high-current interruptions increases as well.

Electro hydro dynamic (EHD) and electrostatic propulsion devices were developed in the 1920s by Thomas Townsend Brown. One such device, called a “lifter”, has no moving parts and, in the air, operates on electrical energy. It is a fashionable device and has a very simple structure, basically consisting of a narrow wire electrode and a large, flat one. However, it has a low ratio of propulsion force to unit electrical input power. According to theory, the propulsion force it generates depends on the interaction between the ion density of the ionized air and the charges on the surface of the large electrode. EHD and electrostatic propulsion models using the polarization effect are proposed to improve the ratio of the propulsion force to unit electrical input power. The propulsion device generates propulsion force through the use of an asymmetrical metal structure with charges generated by the polarization effect. The propulsion force the new devices generated for the same electric energy was 5.7 times higher than that of a basic type lifter owing to additional propulsion force being generated by the maximum polarization effect in the experiments. It was found that combining other effects with this polarization effect results in the ratio of generated propulsion force to electric power being close to 100N/kW when the electric power is high. This value is as high as that of a helicopter. We also performed numerical analysis was also performed for capacitances and charges for various kinds of EHD and electrostatic propulsion devices. An optimized system was developed and is discussed in this paper.