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The subject matter of the article is the radial distribution of electrons movement parameters inside electric propulsion thrusters with closed electrons drift. The radial magnetic field in Hall effect thrusters is the limits the axial flow of electrons because of interaction with azimuth electron current. In turn, this azimuth current exists as a result of rivalry between the attempt of the magnetic field to transform electrons current completely closed one and the loss of electrons rotation moment in collisions. Similar processes take place in the ionization chamber of plasma-ion thrusters with the radial magnetic field. The attempts to estimate electrons parameters through only collisions with ions and atoms inside volume have given the value of axial electrons current much lower than really being. This phenomenon is called anomalous electrons conductivity, which was tried to be explained as a consequence of various effects including "near-the-wall-conductivity", which was explained as a result of non-mirror reflection of electrons from the Langmuir layer near the walls of the thruster channel. The disadvantage of this name is the fact that the reflection of the electron occurs before reaching the surface from the potential barrier at the plasma boundary with any environment: the wall, but also with the environment vacuum. The potential distribution in the Langmuir layer is non-stationary and inhomogeneous due to the presence of so-called plasma oscillations. The definition of "conductivity" is just as unfortunate in this name, because the collisions are always not a factor of conductivity, but on the contrary – of resistance. The goal is to solve the task of electrons rotation moment distribution in the thruster channel. The methods used are the formulation of the kinetic equation for electrons distribution function over the velocities, radius, and projections of the coordinates of the instantaneous center of cyclotron rotation; solution of this equation and finding with its use the distribution of the gas-dynamic parameters of electrons along the cross-section of the channel. Conclusions. A mathematical model of electrons rotation moment dynamics is proposed, which allows using plasma-dynamics equations to analyze its distribution along the cross-section of thruster channel and to estimate the effect of "near-the-wall-conductivity" using appropriate boundary conditions.

This Paper investigates the role of microcracks in Hall thruster wall erosion. The formation and growth of microcracks on the surface of M26 grade boron nitride composite due to repeated thermal sh...

Hall thrusters must be frequently ignited and extinguished during operation in orbit. The thruster is typically extinguished by terminating the anode voltage. However, the equilibrium state of the plasma in the discharge channel is lost when the external electric field disappears. This consequently changes the state of the plasma, causing rotating spokes to appear. However, spoke rotation ceases after nine rotations. Therefore, in this study, a high-speed camera was used to study the instantaneous spoke rotations of a Hall thruster during the extinguishing process. Additionally, the characteristics of the azimuthal rotating spoke during this process were studied. It was found that, because the anode voltage does not instantaneously decrease to 0 V during the extinguishing process, the spoke mode changes from multi-mode to single-mode. Furthermore, as the voltage dropped from 250 to 0 V, a single-mode spoke appeared, and the spoke speed linearly decreased.

The design of the magnetic field is crucial to the performance of the hall thruster. This paper investigates the effects of magnetic field on the discharge characteristics of the low-power hall thruster in the method of combining simulation and experiment. Based on a 40-mm-diameter hall thruster (LHT-40) operating with hundreds of Watts, a two-dimensional axisymmetric model is established in fluid method to investigate the mechanism, how the magnetic field influences the performance of the thruster, and the experiments are also carried out to investigate the performance effected by magnetic field. Both the numerical and the experimental results indicate that there is an optimal magnetic flux density in which the thruster could achieve a higher efficiency compared with other conditions. And in order to maintain its high efficiency, the line of magnetic-curve inflexion inside the discharge channel should coincide with the center line of the channel to decrease the interaction between the plasma and the wall. The LHT-40 could generate about 12.2 mN of thrust at an optimal efficiency of 27% when the maximum magnetic flux density along the center line of the channel is about 200 Gs.

A radial particle-in-cell model of the weakly-collisional plasma discharge in a Hall thruster, provides the non-Maxwellian velocity distribution functions (VDF) of ions and electrons. The model considers a radial magnetic field, secondary electron emission from the two walls, and phenomenological models of anomalous electron scattering. The electron VDF is used to assess the different terms in the macroscopic momentum and energy equations, identifying those differing from the standard fluid model for a near-Maxwellian VDF. The pressure tensor consists of an anisotropic gyrotropic part and a small gyroviscous part. Nonetheless, the gradient of this last one affects the cross-field electron current density, generating radial undulations that resemble those reported for near-wall conductivity. A gyroviscous energy flux is identified too. The heat flux parallel to the magnetic lines does not follow a conductive-type law but a convective-type one, already found in other weakly-collisional plasmas. The tails of the electron velocity distribution function are partially depleted due to wall collection, leading to reduced electron fluxes of particles and energy, which are characterized with parameters useful for fluid models. Differences in the plasma response for annular and planar channel geometries are highlighted. The levels of replenishment of the electron VDF and of the asymmetries in radial profiles differ for isotropic and anisotropic anomalous scattering models.

We revisit the electrostatic instability model of Singh and Malik (IEEE Trans. Plasma Sci. 39:10 (2011) 1910–1918) and correct qualitative and quantitative errors.

The analysis is represented of some works devoted to the mathematical modeling of processes in plasma-ion thrusters and Hall effect thrusters. It is shown that the common in these works is the use of approximate forms of the equations of gas dynamics, which are applicable to the description of relatively dense gases, but not to analyze the processes in the rarefied plasma of electric propulsion thrusters. As a result, the above mathematical models do not represent the processes that are significantly responsible for the values of the thruster operating parameters.Authors try to partially correct this drawback by insertion into the initial approximate forms of the equations written for a point in the plasma volume, the parameters that actually represent the boundary effects and should be written not in the equations of gas dynamics themselves, but in the boundary conditions for these equations.The most complete forms of the necessary equations are given in this paper. It is shown that it is necessary to take into account electrons thermal conductivity as well as at least one (radial-azimuth) component of viscosity tensor to describe the "wall scattering" effect.It is concluded that the most productive approach in mathematical modeling is to write the most complete forms of equations with their subsequent simplification – removing the terms responsible for the processes recognized on the basis of primary numerical estimates as such, which can be neglected.

Technically, there are two types of propulsion systems namely chemical and electric depending on the sources of the fuel. Electrostatic thrusters are used for launching small satellites in low earth orbit which are capable to provide thrust for long time intervals. These thrusters consume less fuel compared to chemical propulsion systems. Therefore for the cost reduction interests, space scientists are interested to develop thrusters based on electric propulsion technology. This chapter is intended to serve as a general overview of the technology of electric propulsion (EP) and its applications. Plasma based electric propulsion technology used for space missions with regard to the spacecraft station keeping, rephrasing and orbit topping applications. Typical thrusters have a lifespan of 10,000 h and produce thrust of 0.1–1 N. These devices have E→×B→ configurations which is used to confine electrons, increasing the electron residence time and allowing more ionization in the channel. Almost 2500 satellites have been launched into orbit till 2020. For example, the ESA SMART-1 mission (Small Mission for Advanced Research in Technology) used a Hall thruster to escape Earth orbit and reach the moon with a small satellite that weighed 367 kg. These satellites carrying small Hall thrusters for orbital corrections in space as thrust is needed to compensate for various ambient forces including atmospheric drag and radiation pressure. The chapter outlines the electric propulsion thruster systems and technologies and their shortcomings. Moreover, the current status of potential research to improve the electric propulsion systems for small satellite has been discussed.

Hall effect thrusters (HETs) are an increasingly utilized proportion of electric propulsion devices due to their high thrust-to-power ratio. To enable an accessible research thruster, our team used inexpensive materials and simplified structures to fabricate the 44-mm-diameter Western Hall Thruster (WHT44). Anode flow, discharge voltage, magnet current, and cathode flow fraction (CFF) were independently swept while keeping all other parameters constant. Simultaneously, a Faraday probe was used to test plume properties at a variety of polar coordinate distances, and an oscilloscope was used to capture discharge oscillation behavior. Plasma plume divergence angle at a fixed probe distance of 4.5 thruster diameters increased with increasing anode flow, varying from 36.7∘ to 37.4∘. Moreover, divergence angle decreased with increasing discharge voltage, magnet current, and CFF, by 0.3∘, 0.2∘, and 8∘, respectively, over the span of the swept parameters. Generally, the thruster exhibited a strong oscillation near 90 kHz, which is higher than a similarly sized HET (20–60 kHz). The WHT44 noise frequency spectra became more broadband and the amplitude increased at a CFF of less than 1.5% and greater than 26%. Only the low flow and low voltage operating conditions showed a quiescent sinusoidal discharge current, otherwise, the discharge current probability distribution was Gaussian. This work demonstrates that the WHT44 thruster, designed for simplicity of fabrication, is a viable tool for research and academic purposes.

Ion impact emission cross sections for eleven transitions from the 5p6p configuration to the 5p6s configuration of neutral xenon occurring in the spectral region between 700 nm and 1000 nm have been measured experimentally. Collisions between both singlyand doublyionized xenon and neutral xenon have been studied. These cross sections are of primary use in the development of a spectrographic diagnostic for Hall effect thruster plasma. A detailed discussion of the experimental methods and the subsequent data reduction is included. The results are presented and the importance of these data for spectrographic emission models of Hall effect thruster plasmas is discussed.

The chapter discussed the technological application of plasma physics in space science. The plasma technology is using laser-plasma fusion, inertial fusion, Terahertz wave generation and welding of metals. In this chapter, the application of plasma physics in the field of electric propulsion and types has been discussed. These devices have much higher exhaust velocities, longer life time, high thrust density than chemical propulsion devices and useful for space missions with regard to the spacecraft station keeping, rephrasing and orbit topping applications. The mathematical relation has been derived to obtain the performance parameters of the propulsion devices.

The Hall thruster is a high-efficiency spacecraft propulsion device that utilizes plasma to generate thrust. The most common variant of the Hall thruster is the stationary plasma thruster (SPT). Erosion of the SPT discharge chamber wall by plasma sputtering degrades thruster performance and ultimately ends thruster life. Many efforts over the past few decades have endeavored to understand wall erosion so that novel thrusters can be designed to operate for the thousands of hours required by many missions. However, due to the challenges presented by the plasma and material physics associated with erosion, a complete understanding has thus far eluded researchers. Sputtering rates are not well quantified, erosion features remain unexplained, and computational models are not yet predictive. This article reviews the physics of plasma-induced SPT erosion, highlights important experimental findings, provides an overview of modeling efforts, and discusses erosion mitigation strategies.

Hall thrusters are widely applied to space propulsion missions because of characteristics such as high specific impulse, great thrust-to-power ratio, high efficiency, and simple structure. Because the scope of space missions using Hall thrusters has expanded, a higher total impulse is required to extend a north–south station-keeping mission to an orbit transfer mission, requiring Hall thrusters with improved lifetimes. Existing research shows that improving the anti-sputtering capacity of the discharge channel and significantly reducing the ion energy and flux bombardment on the wall of the discharge channel are the best measures to improve the lifetime of Hall thrusters. This study summarizes the lifetime restraining factors of Hall thrusters and analyzes the magnetic shielding technology, wall-less technology, aft-magnetic technology, and cylindrical Hall thruster technology to reduce ion energy and flux bombardment on the discharge-channel wall, adopting low-sputtering yield materials to reduce the wall erosion rate, which can improve the lifetime of Hall thrusters. This paper also identifies the limitations of each technology, such as the erosion of magnetic poles (MS) and ceramic cover plates (CHT), larger plume divergence angle (wall-less), hard-to-select aft-magnetic degree (aft-magnetic technology), low compactness, and great porosity (graphite material), providing a reference for further research.

Optimization of magnetic and electric fields has been central concern for the design of a Hall current plasma accelerator since its inception decades ago. However, neutral flow dynamics in the discharge channel may have as much impact on the accelerator performance, operation stability and lifetime as the magnetic and electric fields due to its strong coupling with plasma properties. In this article, auxiliary gas injection is numerically studied for a low-power accelerator using a two-dimensional fully kinetic particle-in-cell code. Gas injection through the discharge channel sidewalls increases lifetime of the accelerator, but also degrades thrust performance suggesting that there is an optimum gas injection ratio. Although reduction in the maximum erosion rate is substantially lower than that predicted by a two-dimensional hybrid model for a high-power accelerator [14], extension of lifetime by approximately 20% appears to be possible with little impact (2%) on the thrust. The anode efficiency analysis supported by the simulated plasma properties clarifies that reduction in voltage utilization is the main cause of the observed alterations in the plasma properties and thrust performance deterioration.

Two distinct discharge modes were observed in a 50 W-class micro-Hall thruster plasma under different operating conditions. A ball-shaped plasma with a broad plume (mode A) was observed at low mass flow rates (less than 0.37 mg s−1) over the entire operational anode voltage range (160–280 V). Raising the anode voltage beyond 200 V with the mass flow rate fixed (larger than 0.37 mg s−1) produced a narrow plume and stretched jet-like structure (mode B). In mode B, the thruster showed performance improvements in terms of thrust (3.8 mN vs 3.3 mN), specific impulse (913 s vs 800 s), and anode efficiency (28% vs 22%), with only a 2 W difference in the anode power (61 W in mode B and 59 W in mode A). This suggests that operation is more advantageous in mode B than in mode A for the utilization of such low-power Hall thrusters. Unique plume properties were observed in the two modes and considerable differences were measured in the Xe II ion acceleration structure, beam angle, and ionization rate as the mode changes, which were not reported in previous studies. Mode A exhibits an axially extended ion acceleration structure outside the discharge channel, where 75% of the final ion velocity is achieved at approximately 40 mm from the thruster exit, while most of the ion acceleration occurs within 10 mm from the thruster exit in mode B. Measurements show that the full width at half maximum of the Xe II ion energy distribution function, electron temperature, and Xe II emission intensity decreased after the plasma transitioned from mode A to mode B. Based on the optical emission spectroscopy, the ionization rate in the plasma plume decreased by 30%–41% after the mode change, which is likely related to the reduction of the beam angle and electron current by 24% and 30%, respectively.

This paper presents a brief overview of the application of advanced plasma modeling techniques to several space science and engineering problems currently of significant interest. Recent advances in both kinetic and fluid modeling provide the ability to study a wide variety of problems that may be important to space plasmas including spacecraft–environment interactions, plasma–material interactions for propulsion systems such as Hall thrusters, ionospheric plasma instabilities, plasma separation from magnetic nozzles, active space experiments, and a host of additional problems. Some of the key findings are summarized here.

Plasma engines are used for space propulsion as an alternative to chemical thrusters. Due to the high exhaust velocity of the propellant, they are more efficient for long-distance interplanetary space missions than their conventional counterparts. An advanced laboratory of plasma space propulsion (PlaNS) at the Institute of Plasma Physics and Laser Microfusion (IPPLM) specializes in designing and testing various electric propulsion devices. Inside of a special vacuum chamber with three performance pumps, an environment similar to the one that prevails in space is created. An innovative Micro Pulsed Plasma Thruster (LμPPT) with liquid propellant was built at the laboratory. Now it is used to test the second prototype of Hall effect thruster (HET) operating on krypton propellant. Meantime, an improved prototype of krypton Hall thruster is constructed.

The paper presents the results of Hall thruster plasma simulation using a 2D3V axial-azimuthal Full PIC model. Two types of solution are investigated in detail for two different magnitude of the magnetic field. In the calculations, different types of azimuthal waves were observed. By their characteristics these waves can be corresponded to the electron drift instabilities, gradient-drift waves and ion plasma resonance. The ion plasma resonance was observed in calculations with the magnetic field increase.

An investigation on the influence of relevant parameters on an annular Hall effect thruster plasma discharge is performed using a radial particle-in-cell simulation code with secondary electron emission from the walls and prescribed axial electric and radial magnetic fields. A simulation with true-secondary electrons only is taken as reference. First, the near-wall conductivity effects on the magnetized secondary electrons are illustrated by doubling the , allowing a further code validation. Second, when secondary backscattered electrons are included, the enhanced secondary emission yields lower sheath potential drops and primary electron temperature. Moreover, the dominant backscattered electrons increase the average secondary electrons emission energy, greatly affecting its temperature anisotropy ratio and increasing the replenishment level of the wall collectable tails of the primary electrons velocity distribution function. Third, the effect of the true-secondary electrons emission energy on the potential profile is shown to be negligible, the latter being mainly set by the dominant magnetic mirror effect. Finally, a planar case featuring symmetric plasma profiles permits to confirm the validity of the large cylindrical asymmetries present in the reference case, induced by the combined effects of the geometric expansion, the magnetic mirror and the centrifugal force (due to the drift). A smaller deviation of the primary electron momentum equation from the Boltzmann relation along the magnetic lines is still found in the planar case, induced by the parallel temperature non-uniformity. The UC3M researchers have been supported by the PROMETEO-CM project, Grant number Y2018/NMT-4750 (Comunidad de Madrid/FEDER/FSE). Additional funding for A Domínguez-Vázquez came from Project ESP2016-75887 (Spain's National Research and Development Plan - MINECO/FEDER) F Taccogna has been supported by the italian Ministero dell'Istruzione, dell'Università e della Ricerca (MIUR) under the CLOSE project (grant ARS01_00141).

Electric propulsion systems are a more mass efficient method for providing a change in velocity, ΔV, to on-orbit spacecraft, than their chemical counterparts. In comparison, electric systems generally have a much higher specific impulse, Isp, than chemical systems. One option within the realm of electric propulsion is Hall Effect Thrusters, which have moderately high specific impulse values. From their advent in the 1960s, Hall Effect Thrusters have been used for orbit station keeping, attitude control, and orbit transfer. Although the discharge cavity is conventionally circular, pseudo-linear or racetrack shaped cavities have been developed. Even though Hall thrusters have decades of flight heritage, there are still many plasma behavioral characteristics which are still unknown. Multiple non-intrusive measurement techniques were used to investigate plasma behavior both in the plume and in the channel of a pseudo linear Hall thruster. Through the visible emissions captured by a high-speed camera, breathing mode characteristics were induced and analyzed. Spoke structures were observed in only certain parts of the thruster channel. Additionally, the plume divergence was characterized by use of a Faraday probe along two axes of the thruster, indicating significantly different “keep out” regions for potential thruster use on spacecraft. Also, an irregularity was observed in the channel of the pseudo-linear thruster, which potentially could affect the lifespan of the thruster.

We present measurements of the electron density, the effective electron temperature, the plasma potential, and the electron energy probability function (EEPF) in the plume of a 1.5 kW-class SPT-100 Hall thruster, derived from cylindrical Langmuir probe measurements. The measurements were taken on the plume axis at distances between 550 and 1550 mm from the thruster exit plane, and at different angles from the plume axis at 550 mm for three operating points of the thruster, characterized by different discharge voltages and mass flow rates. The bulk of the electron population can be approximated as a Maxwellian distribution, but the measured distributions were seen to decline faster at higher energy. The measured EEPFs were best modelled with a general EEPF with an exponent alfa between 1.2 and 1.5, and their axial and angular characteristics were studied for the different operating points of the thruster. As a result, the exponent alfa from the fitted distribution was seen to be almost constant as a function of the axial distance along the plume, as well as across the angles. However, the exponent alfa was seen to be affected by the mass flow rate, suggesting a possible relationship with the collision rate, especially close to the thruster exit. The ratio of the specific heats, the gamma factor, between the measured plasma parameters was found to be lower than the adiabatic value of 5/3 for each of the thruster settings, indicating the existence of non-trivial kinetic heat fluxes in the near collisionless plume. These results are intended to be used as input and/or testing properties for plume expansion models in further work. This work was performed in the framework of the 'Model and Experimental validation of spacecraft-thruster Interactions (erosion) for electric propulsion thrusters plumes' (MODEX) project. MODEX is a collaboration between Airbus-DS, ESA, UC3M, ONERA, CNRS-ICARE and KTH aiming to provide a better understanding of the plasma properties in the far-plume of a Hall thruster. The project aimed at providing experimental measurements to better constrain the modelling, and therefore includes both the theoretical/modelling aspect (UC3M and ONERA) and the experimental aspect (KTH, CNRS, ESA and Airbus-DS). The test campaign was conducted at ESA/ESTEC in April-May 2017, using a SPT-100 Hall thruster provided by Airbus-DS. G Giono and J T Gudmundsson were partially supported by the Swedish Government Agency for Innovation Systems (VINNOVA) contracts no. 2016-04094 and 2014-0478, respectively.

Negative ion sources are a key component of neutral beam injection systems. A novel concept for a negative ion source based on existing well tested Hall Thrusters (HT) is presented. The thruster has been designed in order to maximize the hydrogen dissociation process so that the atom flux which exits the thruster has an optimized energy to generate negative ions when impinging on a low work function surface. The novel concept can in principle offer several advantages, for instance a limited amount of co-extracted electrons, a more uniform generation of negative ions and a lower rate of destruction of negative ions. The thruster has been completed and is now in operation. In this contribution, the behavior of the thruster with different mass flow rate and voltage is presented. In particular, the first measurements of the plume plasma parameters by a Langmuir probe are shown, both with nitrogen and with hydrogen as injected gas. The spatial profile of the plasma electronic temperature and density have been obtained by scanning the region outside the exit plane of the thruster. The measurements are aimed to determine the best position of a caesiated sample in order to maximize the generation of the negative hydrogen ions.

The E × B drift instability is studied in Hall thruster using one-dimensional particle in cell (PIC) simulation method. By using the dispersion relation, it is found that unstable modes occur only in discrete bands in k space at cyclotron harmonics. The results indicate that the number of unstable modes increases by increasing the external electric field and decreases by increasing the radial magnetic field. The ion mass does not affect the instability wavelength. Furthermore, the results confirm that there is an instability with short wavelength and high frequency. Finally, it is shown that the electron and ion distribution functions deviate from the initial state and eventually the instability is saturated by ion trapping in the azimuthal direction. Also for light mass ion, the frequency and phase velocity are very high that could lead to high electron mobility in the axial direction.

Synopsis To assess and improve the design aspects during the design and development of HETs, it is essential to perform plasma dynamic simulations on these thruster designs. A 2D-3V (two dimensional-three velocity) Particle-in-cell Monte Carlo (PIC-MC) model has been used to simulate the standard SPT-100, and the design was optimized. Later, a completely new HET geometry has been simulated with an aim to optimize the thruster design with reduced plasma wall interaction.

According to experiments, a jumplike repartition of the discharge burning takes place at some parameters of Hall-thruster operating. This paper presents results of probing measurements demonstrating the difference in the distribution of floating and plasma potentials, electron temperature, and electric-field strength in a channel of the Hall thruster at two different types of discharge burning. The measurements were carried out along the discharge-channel axis using Langmuir probes.

This study analyzes the dependence of electron extraction efficiency, which is defined as the ratio of the extracted electron current to the generated electron current, on the orifice shapes and magnetic fields of a miniature microwave discharge xenon neutralizer via three-dimensional particle-in-cell simulations with Monte Carlo collisions (PIC–MCCs). The PIC–MCC simulation results show that the orifice shapes do not significantly affect the discharge characteristics or the electron extraction efficiency. However, the efficiency achieves a 1.5-times higher value in a new magnetic field configuration, referred to as MF-2, where the magnetic field lines pass through nearly the entire area of the orifices. This improvement is attributed to the reduction in the electron backflow and the electron loss toward both the downstream inside surface and the outside wall of the discharge chamber. In addition, there are relatively small plasma fluctuations in the discharge chamber for MF-2 due to its low Bohm diffusion coefficient, where no rotating spokes, which are often seen in other E ×B devices, are observed. As a result, the electron loss toward the downstream surface inside the discharge chamber is reduced, and this decrease in the electron loss also contributes to the increase in the extraction efficiency., This article may be downloaded for personal use only. Any other use requires prior permission of the author and AIP Publishing. This article appeared in Y. Sato et al., Jnl. of App. Phy. 126, 243302 (2019) and may be found at https://doi.org/10.1063/1.5127805.

Time-resolved measurements of the plasma parameters are performed in the plume of a cross-field discharge. The plasma potential is measured with a cylindrical Langmuir probe and an emissive probe. The electron temperature and density are measured with a cylindrical Langmuir probe. The cross-field discharge is maintained in a harmonic oscillation regime to guarantee reproducible conditions for all measurements (© 2013 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim)

International audience; Collective light scattering (CLS) has been recently extended to the observation of Hall effect thruster plasmas. A better understanding of the physics of Hall thrusters is key to improving thruster operation and lifetimes. In particular, the origins of thruster phenomena such as anomalous electron transport need to be determined. A high-performance CLS diagnostic has therefore been developed for identifying and measuring electron density fluctuations in the thruster, at millimetric length scales and MHz frequencies. Such modes are believed to play a role in anomalous transport, and experiments so far performed have provided information on their dispersion relations, amplitude and directivity. This work describes the technical aspects of the optical bench and the range of accessible experiments (© 2011 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim)

Since the inception of the Hall-effect thruster more than five decades ago, many theories have been proposed about the source(s) of the anomalous cross-field transport that is known to occur in these devices. Yet, none of these theories has been implemented in 2-D (r-z) simulations and successfully predicted the behavior of the discharge and the erosion of a thruster over different operating conditions and geometries. We present results from numerical experiments with a 2-D (r-z) axisymmetric hydrodynamics code that show the plasma solution is relatively insensitive to very large variations in the anomalous transport in some regions of the discharge channel and near-plume. Because the changes of the plasma properties in these regions can be too small or impossible to detect in the laboratory by conventional diagnostics, the verification of a transport model becomes very challenging. Hence, without more advanced diagnostics and improved numerical models, comparisons between plasma measurements and simulation results can lead to wide-ranging theories and scalings of the anomalous resistivity.

Radio blackout that occurs during hypersonic reentry flight is an important issue for the operation of the vehicle. Since the radio blackout problem is caused by a high plasma number density around a vehicle, it is necessary to manipulate the plasma to allow communication. We suggest a configuration of an E B layer as a reentry blackout mitigationmethod.ThesuggestedE Blayerconfigurationwithatwo-dimensionalmagnetic fieldissimulatedusing thethermalized potentialmodelandthePoisson-like modelin ordertoillustrate theeffectiveness of thisapproachas amitigationmethod.Thenumericalmodelusesamagnetohydrodynamicsapproximationandissolvedusinga finite volumemethodwithaRiemannsolver.Theresultsofthenumericalmodelareassessedusingavailableexperimental results.Astrongplasmadensityreductionisobtainedwhenthehighelectricandmagnetic fieldsareappliednearthe cathode. The manipulated plasma region provides a possibility to communicate through a plasma layer during a reentry blackout.

Simulation on the control of the ion current oscillation is performed by introducing the effect of the Xe ion transport in plasmas produced in the channel of Hall thrusters. Numerical results are in good agreement with the preceded experimental observations. It is shown for the first time that the ion current becomes stable and the thrust efficiency is improved when the ion mobility and the cross sectional area decrease. (© 2008 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim)