Study on Similarity Criteria of Hall Effect Thrusters

The scaling design of Hall effect thrusters is based on similarity criteria. Up to now, few of the similarity criteria proposed concern about the inside physical processes of the thruster except Melikov‐Morozov similarity criterion which embodies the ionization of propellant. As many other significant processes, such as electron conduction, ion acceleration and energy exchange, are out of consideration, it is far from enough to direct the thruster design appropriately at present. Therefore, in this paper, we have deduced out many new similarity criteria by analyzing the neutral continuity equation, the ion/electron momentum equations and the electron energy equation with the equation analysis method. By further comparing the magnitudes of source terms of those equations, we obtain certain similarity criteria which are primary and should be guaranteed for modeling design of thrusters. These dominant similarity criteria are finally verified in experiments (© 2009 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

Scaling Design and Experimental Study on Hall Thrusters with Curved Magnetic Field

The existing scaling theories of Hall thrusters are based on the hypothesis of a one‐dimensional straight magnetic field, which is not suitable for the design of modern thrusters with a two‐dimensional curved magnetic field. In this paper, using the equation analysis method, we derive new similarity criterions in a curved magnetic field by analyzing the momentum equations of charged particles; consequently, we propose a new modeling design method for Hall thrusters with a constant discharge voltage. This method is further validated by experiments. A designed model with a power of 1.5 kW is made based on our proposed method from a prototype model with a power of 1 kW. The experimental results demonstrate that these two thrusters have little differences in performance and physical processes as expected from the scaling. Therefore, our method is well suited for designing a Hall thruster with a curved magnetic field (© 2011 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

Particle‐in‐Cell Simulations for Ion Thrusters

The Particle‐in‐Cell (PIC) method was used to study two different ion thruster concepts: Hall Effect Thrusters (HETs) and High Efficiency Multistage Plasma Thrusters (HEMPs), in particular the plasma properties in the discharge chamber due to the different magnetic field configurations. Special attention was paid to the simulation of plasma particles fluxes on the thrusters inner surfaces. In both cases PIC proved itself as a powerful tool, delivering important insight into the basic physics of the different thruster concepts.The simulations demonstrated that the new HEMP thruster concept allows for a high thermal efficiency due to both minimal energy dissipation and high acceleration efficiency. In the HEMP thruster the plasma contact to the wall is limited only to very small areas of the magnetic field cusps, which results in much smaller ion flux to the thruster channel surface as compared to HET. (© 2009 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

Fluid Simulation of a Cusped Field Thruster

The cusped field thruster is a kind of newly developed electric propulsion device. The electric field at the channel exit and the low frequency oscillation were measured by former experiments. While the formation mechanism of them have not been fully interpreted yet. Through studying two distinguishing typical electron paths in the thrusters, a fluid model is built up along two electron paths, and then the model is completed by synthetically analyzing the effect of two electron paths. Time‐averaged electric potential distribution, anode current oscillation curve and time‐synchronized atom density and ion density distribution are obtained by simulation. (© 2015 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

Effects of Coupling Oscillations on the Electron Conductivity of Hall Thrusters

It is known that the static magnetic field of a Hall thruster remains unchanged with the product of coil current and coil turn kept constant. This is called the ampere‐turns equivalence principle, which is used in the static magnetic field design of Hall thrusters. It is found that the discharge characteristic is different with the same static magnetic field which is formed by the same ampere‐turns but different coil turns. The effects of the fluctuating magnetic field on the static magnetic field are studied experimentally and theoretically, and the effective collision frequency is calculated based on the dispersion relation of coupling oscillations and the estimation of the fluctuating magnetic field. Results indicate that the different coil turns are accompanied by different coupling intensities which lead to different magnetic field fluctuations. The different magnetic field fluctuations cause differences in the effective electron collision frequency and therefore the electron conductivity and the discharge characteristic of Hall thrusters (© 2012 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

离子推力器工作性能参数控制模型

为了研究离子推力器工作参数对输出特性的影响,通过离子推力器工作性能参数的理论计算公式,建立起离子推力器输入参数与输出参数的Simulink控制模型,根据模型分别对我国研制的30 cm口径以及20 cm口径离子推力器的工作输出参数进行了理论计算,并通过推力测量试验对理论值进行了比对和分析。比对结果表明:在推力理论计算过程中引入二价Xe离子比率和束流密度分布推力修正,以及推力均方误差修正后,推力理论值与实测值符合性较好,计算误差小于1 mN,证明了推力修正方法的合理性。     

Magnetic confinement studies for performance enhancement of a 5-cmion thruster

Several magnet assemblies relevant to ion thrusters were investigated using a numerical code to calculate the primary electron mean containment time. An analytical model was used to determine, consistently with the plasma stability condition, the ion thruster performance parameters, namely, the plasma ion energy cost, the plasma density, the plasma potential, and the doubly charged ion production rate as a function of the propellant mass flow rate. The numerical code was checked by reproducing the experimental performance parameters obtained from a 7-m checkerboard ion thruster. Using this approach, performance curves were obtained for a 5-cm ion thruster devised to produce a 1-mN thrust with low power and propellant consumption     

栅极热变形对离子推力器工作过程影响分析

栅极热变形是影响离子推力器性能和寿命的重要因素. 采用三维粒子方法对栅极系统等离子体输运过程进行模拟, 对比、分析栅极热变形前后栅极系统的电子返流限制、导流系数限制、离子通过率和发散角损失. 结果表明: 栅极热变形增大了屏栅离子通过率和推力器推力值, 并由于加速栅截止电流阈值的提高拓展了推力器工作电流区间, 但电子返流阈值的明显降低对栅极系统可靠工作造成了不利影响.     

Particle‐in‐cell simulation of an optimized high‐efficiency multistage plasma thruster

Electric propulsion attracts increasing attention in contemporary space missions as an interesting alternative to chemical propulsion because of the high efficiency it offers. The High‐Efficiency Multistage Plasma thruster, a class of cusped field thruster, is able to operate at different anode voltages and operation points and thereby generate different levels of thrust in a stable and efficient way. Since experiments of such thrusters are inherently expensive, multi‐objective design optimization (MDO) is of great interest. Several optimized thruster designs have resulted from a MDO model based on a zero‐dimensional (0D) power balance model. However, the MDO solutions do not warrant self‐consistency due to their dependency on estimation from empirical modelling based on former experimental studies. In this study, one of the optimized thruster designs is investigated by means of particle‐in‐cell (PIC) analysis to examine the predicted performance characteristics with self‐consistent simulations. The 0D power balance model is used to develop additional diagnostics for the PIC simulations to improve the physics analysis. Using input parameters for the 0D power balance model from the PIC simulations allows further improvement for the design optimization.     

Space-charge-limited current of vacuum arc thruster

To optimize thrust performance, the expression of space-charge-limited current for vacuum arc thruster is derived from Poisson's equation. The commonly used ring-type and coaxial-type vacuum arc thrusters are simplified to the equivalent current sheet in planar geometry and cylindrical capacitor, respectively, for this calculation. Both the spatial distribution and peak magnitude of space-charge-limited current are given explicitly, together with their dependences on gap distance, applied voltage, charge number, and ion mass. For typical experimental parameters of the vacuum arc thruster, it is shown that the maximum current density drops significantly when the gap distance becomes large and grows when the applied voltage increases; moreover, a cathode material of lower atomic weight yields a higher current density. The expressions of total current for these two types of vacuum arc thruster are also presented. This work, to our best knowledge, is the first application of space-charge-limited current to the vacuum arc thruster and practically very interesting for engineering design.     

Similarity scaling-application and limits for high-efficiency-multistage-plasma-thruster particle-in-cell modelling

To suit a wide variety of space mission profiles, different designs of ion thrusters were developed, such as the High-Efficiency-Multistage-Plasma thrusters (HEMP-T). In the past, the optimization of ion thrusters was a difficult and time-consuming process and evolved experimentally. Because the construction of new designs is expensive, cheaper methods for optimization were sought-after. Computer-based simulations are a cheap and useful method towards predictive modelling. The physics in HEMP-T requires a kinetic model. The Particle-in-Cell (PIC) method delivers self-consistent solutions for the plasmas of ion thrusters, but it is limited by the high amount of computing time required to study a specific system. Therefore, it is not suited to explore a wide operational and design space. An approach to decrease computing time is self-similarity scaling schemes, which can be derived from the kinetic equations. One specific self-similarity scheme is investigated quantitatively in this work for selected HEMP-Ts, using PIC simulations. The possible application of the scaling is explained and the limits of this approach are derived.     

Study on the Relation Between Discharge Voltage and Magnetic Field Topography in a Hall Thruster Discharge Channel

The relation between magnetic field topography and operating voltage is investigated in a 1kW Hall thruster discharge channel in order to focus the ion beam effectively and optimize the performance. The curvature of magnetic field line (α) is introduced to characterize the differences of topologies. The optimized magnetic field distribution under each operating voltage is obtained by experiment. Through the curvature transformation, we find that the area of (α > 1) in the channel gradually decreases with the increase of the operating voltage. In response to the results above, two dimensional plasma flows are simulated employing Particle‐in‐Cell method. The distributions of the electric potential, ion density and ion radial velocity are calculated to understand the important influence of the relation above on ion beam focusing. The numerical results indicate that magnetic field curvature and thermal electric field control the ion beam in the ionization and acceleration zone, respectively. The magnetic field topography and discharge voltage interact with each other and together form the focusing electric field. The ion radial mobility is suppressed effectively and the ion beam is focused to the channel centerline. In addition, for a given voltages, when the area of (α > 1) is larger than the optimal scope, the electric potential lines excessively bend to the anode causing ion over focus; contrarily, the electric potential lines will bend to the exit and defocus ions. All these results suggest the relation between magnetic field topography and discharge voltage is important to the ion radial flow control and performance optimization of the Hall thruster (© 2009 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

Experimental and numerical investigation of a Hall thruster with a chamfered channel wall

A discharge channel with a chamfered wall not only has application in the design of modern Hall thrusters, but also exists where the channel wall is eroded, and so is a common status for these units. In this paper, the laws and mechanisms that govern the effect of the chamfered wall on the performance of a Hall thruster are investigated. By applying both experimental measurement and particle-in-cell simulation, it is determined that there is a moderate chamfer angle that can further improve the optimal performance obtained with a straight channel. This is because the chamfering of the wall near the channel exit can enhance ion acceleration and effectively reduce ion recombination on the wall, which is favorable to the promotion of the thrust and efficiency. However, the chamfer angle should not be too large; otherwise, both the density of the propellant gas and the distribution of the plasma potential in the channel are influenced, which is undesirable for efficient propellant utilization and beam concentration. Therefore, it is suggested that the chamfer shape of the channel wall is an important factor that must be carefully considered in the design of Hall thrusters.     

On the Effect of the Surface State and Geometry of the Discharge- Chamber Edges and Sizes of Morozov’s Stationary Plasma Thruster upon its Long-Term Operation

The results of long-term tests of Morozov’s stationary plasma thrusters are presented. It is revealed how the surface state and geometry of the discharge chamber’s edges influence the thruster’s parameters. It is shown that, during the ground tests of thrusters with cylindrical geometry of the acceleration channel under initial stage of operation, material sputtered from the discharge chambers’ walls is deposited onto the nearanode segment of the walls. Films of deposited material fail during thruster operation causing fragment formation, which jut out towards the discharge volume and disturb the motion of drifting electrons in the area of their acceleration. As a result the thruster reactive force and specific impulse decrease. The way in which the forming fragments influence thruster performance and operation is examined. It is shown that it decreases under long-term operation and significant channel widening since the ion flux to the wall and the quantity of the sputtered material decrease, and since the profile of the walls changes due to their wear and cleaning effect of the discharge. As a result the thruster’s parameters are restored to a level close to the initial one. It is shown that the dynamics of thruster parameters variation in space and during ground tests is different. This means that it is necessary to simulate more properly the conditions of thruster operation in space when conducting ground development tests. Thrusters with a long lifetime should be designed with widening of the acceleration channel beyond the loop which surrounds the magnetic system so that areas of acceleration and the erosion of walls are located in the widened part of the channel.     

Doubly‐Charged Xe Ions Evidenced by Time Resolved RPA Measurement in the Far Field Plume of a Low‐Power HET

A low power Hall Effect Thruster (HET), based on a permanent magnet circuit, was investigated in the GREMI laboratory facility. The thruster operated in the working range between 50 and 300 W and the previously measured thrust is between 4 and 16 mN for an anodic efficiency respectively between 15% and 27%. The pulsed character of the thruster current is an important feature of this HET. The ion current's bursts are recorded at 30 and 70 cm from the exit plane in the thruster plume and are time‐resolved, which lead to a preliminary analysis of the time of flight (TOF) phenomena. This paper presents a detailed study of these bursts of ion current in the plume. The total ion current is shown to be a superposition of 2 distinct contributions of charged species. In complement, a controlled single current interruption in stable anodic current condition leads to exactly the same features than in oscillating mode. This crucial verification garantees the validity of the time of flight origine of the two distinct contributions. Then, the slower one is the more intense and is proportional to the ion Xe+ current whereas the faster one could be attributed either to doubly‐charged Xe⁺⁺ or to superfast Xe⁺. The work presents a way to determine unambiguously the nature of the fast contribution by recording the Retardated Potential Analyser (RPA) signals at various repelling grid potentials with respect to time. The energy distribution of the 2 wellseparated contributions are reconstructed and confirms the contribution of doubly‐charged xenon ions (Xe⁺⁺) in the plume. This way of RPA collecting data and interpretation presents the main advantage tobe an easy way for the identification of the nature of the charged species in the plume. (© 2015 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)     

Preliminary investigation of power flow and performance phenomenain a multimegawatt coaxial plasma thruster

Preliminary experimental and theoretical research that was directed toward the study of quasi-steady-state power flow in a large, unoptimized, multimegawatt coaxial plasma thruster is summarized. Large coaxial thruster operation is discussed, and the experimental results are evaluated and interpreted with a view to the development of efficient, steady-state, megawatt-class magnetoplasmadynamic (MPD) thrusters     

多模式离子推力器栅极系统三维粒子模拟仿真

栅极系统是离子推力器推力产生的主要部件,推力器的性能和寿命都与栅极系统密切相关.对于具有多种工作模态的离子推力器,基于电流电压入口的仿真可以有效评估推力器的工作状况.采用三维粒子模拟方法对两栅极系统等离子体输运过程进行了仿真,获得了不同模式下的推力器性能参数,对比NSTAR的在轨测试参数,验证了模型的正确性;分析了工作模式变化对栅极区域电场分布和束流状态的影响以及离子推力器多模式设计需求.分析结果表明:远离栅极系统的外凸型屏栅鞘层和内凹型零等势面、低鞍点电势值和平缓的下游电势分布,有利于提高栅极系统离子通过率,抑制电子返流,减小Pits-and-Grooves腐蚀,是离子推力器工作模式的设计方向;提高束流电压会导致发散角损失增大,但可扩展栅极工作电流范围,在束流强度较大的模式下,使束流具有较好的聚焦状态,有利于减小Barrel腐蚀.研究结果为多模式离子推力器工作模式设计提供了参考.     

离子推力器推力密度特性

离子推力器推力密度分布对航天器轨道维持和修正具有重要影响.采用粒子模拟-蒙特卡罗碰撞方法模拟束流等离子体输运过程,分析束流多组分粒子喷出数量和速度等微观参数,并计算得到单孔束流推力,结合放电室出口等离子体密度分布,进一步对推力密度分布特性分析,最后开展实验验证.研究结果显示:束流中单价离子、双荷离子以及交换电荷离子的推力贡献比分别为84.63%,15.35%和1.82%,可见推力主要来源于束流中的单价离子和双荷离子,交换电荷离子对推力贡献很小;推力密度分布具有较好的中心轴对称性,从推力器中心沿着径向先快速下降后趋于缓慢;与实验结果对比,经验模型相对误差约为4.1%,数值模型相对误差约为2.8%,相比经验模型,数值模型具有更好的准确性.研究结果可为离子推力器推力密度分布均匀性等优化提供参考.     

Characterization of electrospray beams for micro-spacecraft electric propulsion applications

Electric propulsion for spacecraft offers many advantages compared to other traditional counterparts such as chemical propulsion. An electrostatic colloid thruster, well suited for small spacecrafts due to its inherent small size, is investigated. The underlying phenomena governing the plume, namely collisions and focusing are parameterized. The determining properties of the electrospray-based thruster, such as specific impulse, are measured for several propellants and demonstrated to be higher than cold-gas thrusters even for singly charged droplet mode. Moreover, the interesting concept of a dual colloid system with positively and negatively charged cone-jet beams was shown to be feasible by simulations and experiments.     

Collective Electrostatic Interaction of Particles in a Complex Plasma with Ion Flow

Linearized electrostatic potential of a test charge in a complex (dusty) plasma with ion flow is found. Dust component is treated as a continuous medium. Positions of dust particles are assumed to be fixed (unperturbed by the test charge). Calculations are performed using the static dielectric response function found in the framework of the fluid model. The model includes ion loss and ion creation caused respectively by absorption on dust particles and ionization. Dust charge variations and friction force on ions (ion‐neutral and ion‐dust friction) are also present in the model. The main point of the paper is the potential distribution in the plane containing the test charge and oriented perpendicular to the ion flow. The possibility of the electrostatic attraction of two same sign charges in the plane perpendicular to the ion flow is investigated. (© 2005 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)