Influence of the position relationship between the cathode and magnetic separatrix on the discharge process of a Hall thruster

Previous studies have shown that there is an obvious coupling relationship between the installation location of the external cathode and the magnetic separatrix in the plume region of a Hall thruster.In this paper,the particle-in-cell simulation method is used to compare the thruster discharge process under the conditions of different position relationships between the cathode and the magnetic separatrix.By comparing the distribution of electron conduction,potential,plasma density and other microscopic parameters,we try to explain the formation mechanism of the discharge difference.The simulation results show that the cathode inside and outside the magnetic separatrix has a significant effect on the distribution of potential and plasma density.When the cathode is located on the outer side of the magnetic separatrix,the potential above the plume region is relatively low,and there is a strong potential gradient above the plume region.This potential gradient is more conducive to the radial diffusion of ions above the plume,which is the main reason for the strong divergence of the plume.The distribution of ion density is also consistent with the distribution of potential.When the cathode is located on the outer side of the magnetic separatrix,the radial diffusion of ions in the plume region is enhanced.Meanwhile,by comparing the results of electron conduction,it is found that the traiectories of electrons emitted from the cathode are significantly different between the inner and outer sides of the magnetic separatrix.This is mainly because the electrons are affected by the magnetic mirror effect of the magnetic tip,which makes it difficult for the electrons to move across the magnetic separatrix.This is the main reason for the difference in potential distribution.In this paper,the simulation results of macroscopic parameters under several conditions are also compared,and they are consistent with the experimental results.The cathode is located on the inner side of the magnetic separatrix,which can effectively reduce the plume divergence angle and improve the thrust.In this paper,the cathode moves from R=50 mm to R=35 mm along the radial direction,the thrust increases by 3.6 mN and the plume divergence angle decreases by 23.77%.Combined with the comparison of the ionization region and the peak ion density,it is found that the main reason for the change in thrust is the change in the radial diffusion of ions in the plume region.

Influence of acceleration stage electrode voltage on the performance of double-stage Hall effect thruster with adjustable zero magnetic point

The configuration of electrode voltage and zero magnetic point position has a significant impact on the performance of the double-stage Hall effect thruster. A 2D-3V model is established based on the two-magnetic peak type double-stage Hall thruster configuration, and a particle-in-cell simulation is carried out to investigate the influences of both acceleration electrode voltage value and zero magnetic point position on the thruster discharge characteristics and performances.The results indicate that increasing the acceleration voltage leads to a larger potential drop in the acceleration stage, allowing ions to gain higher energy, while electrons are easily absorbed by the intermediate electrode, resulting in a decrease in the anode current and ionization rate. When the acceleration voltage reaches 500 V, the thrust and efficiency are maximized, resulting in a 15%increase in efficiency. After the acceleration voltage exceeds 500 V, a potential barrier forms within the channel, leading to a decrease in thruster efficiency. Further study shows that as the second zero magnetic point moves towards the outlet of the channel, more electrons easily traverse the zero magnetic field region, participating in the ionization. The increase in the ionization rate leads to a gradual enhancement in both thrust and efficiency.

Performance investigation of a low-power Hall thruster fed on iodine propellant

The common propellants used for electric thrusters, such as xenon and krypton, are rare, expensive,and difficult to acquire. Solid iodine attracts much attention with the advantages of low cost,extensive availability, low vapor pressure, and ionization potential. The performance of a lowpower iodine-fed Hall thruster matched with a xenon-fed cathode is investigated across a broad range of operation conditions. Regulation of the iodine vapor's mass flow rates is stably achieved by using a temperature control method of the iodine reservoir. The thrust measurements are finished utilizing a thrust target during the tests. Results show that thrust and anode-specific impulse increase approximately linearly with the increasing iodine mass flow rate.At the nominal power of 200 W class, iodine mass flow rates are 0.62 and 0.93 mg/s, thrusts are7.19 and 7.58 m N, anode specific impulses are 1184 and 826 s, anode efficiencies are 20.8%and 14.5%, and thrust to power ratios are 35.9 and 37.9 m N/k W under the conditions of 250 V,0.8 A and 200 V, 1.0 A, respectively. The operating characteristics of iodine-fed Hall thruster are analyzed in different states. Further work on the measurements of plasma characteristics and experimental optimization will be carried out.

Growth mechanism and characteristics of electron drift instability in Hall thruster with different propellant types

The existence of a significant electron drift instability(EDI) in the Hall thruster is considered as one of the possible causes of the abnormal increase in axial electron mobility near the outlet of the channel. In recent years, extensive simulation research on the characteristics of EDI has been conducted, but the excitation mechanism and growth mechanism of EDI in linear stage and nonlinear stage remain unclear. In this work, a one-dimensional PIC model in the azimuthal direction of the thruster near-exit region is established to gain further insights into the mechanism of the EDI in detail, and the effects of different types of propellants on EDI characteristics are discussed. The changes in axial electron transport caused by EDI under different types of propellants and electromagnetic field strengths are also examined. The results indicate that EDI undergoes a short linear growth phase before transitioning to the nonlinear phase and finally reaching saturation through the ion Landau damping. The EDI drives a significant ion heating in the azimuthal direction through electron–ion friction before entering the quasi-steady state, which increases the axial mobility of the electrons. Using lighter atomic weight propellant can effectively suppress the oscillation amplitude of EDI, but it will increase the linear growth rate, frequency, and phase velocity of EDI. Compared with the classical mobility, the axial electron mobility under the EDI increases by three orders of magnitude, which is consistent with experimental phenomena. The change of propellant type is insufficient to significantly change the axial electron mobility. It is also found that the collisions between electrons and neutral gasescan significantly affect the axial electron mobility under the influence of EDI, and lead the strength of the electric field to increase and the strength of the magnetic field to decrease, thereby both effectively suppressing the axial transport of electrons.

Performance of pulsed plasma thruster at low discharge energy

As the size of satellites scales down, low-power and compact propulsion systems such as the pulsed plasma thruster(PPT) are needed for stabilizing these miniature satellites in orbit. Most PPT systems are operated at 2 J or more of discharge energy. In this work, the performance of a PPT with a side-fed, tongue-flared electrode configuration operated within a lower discharge energy range of 0.5-2.5 J has been investigated. Ablation and charring of the polytetrafluoroethylene propellant surface were analyzed through field-effect scanning electron microscopy imaging and energy-dispersive X-ray spectroscopy. When the discharge energy fell below 2 J, inconsistencies occurred in the specific impulse and the thrust efficiency due to the measurement of the low mass bit. At energy ≥2 J, the performance parameters are compared with other PPT systems of similar configuration and discussed in depth.

Thrust estimate method of an on-orbit Hall thruster using Hall drift current

In order to realize the thrust estimation of the Hall thruster during its flight mission,this study establishes an estimation method based on measurement of the Hall drift current.In this method,the Hall drift current is calculated from an inverse magnetostatic problem,which is formulated according to its induced magnetic flux density detected by sensors,and then the thrust is estimated by multiplying the Hall drift current with the characteristic magnetic flux density of the thruster itself.In addition,a three-wire torsion pendulum micro-thrust measurement system is utilized to verify the estimate values obtained from the proposed method.The errors were found to be less than 8%when the discharge voltage ranged from 250 V to 350 V and the anode flow rate ranged from 30 sccm to 50 sccm,indicating the possibility that the proposed thrust estimate method could be practically applied.Moreover,the measurement accuracy of the magnetic flux density is suggested to be lower than 0.015 mT and improvement on the inverse problem solution is required in the future.

Experimental investigation of the polarityswitching process with different bipolar ionic liquid thruster operating frequencies

The bipolar ionic liquid thruster employs ionic liquid as a propellant to discharge positively and negatively charged high-energy particles under an alternating current(AC)power source,effectively suppressing electrochemical reaction and ensuring charge neutrality.Determining an optimal AC supply power source frequency is critical for sustained stable thruster operation.This study focuses on the emission characteristics of the ionic liquid thruster under varied AC conditions.The AC power supply was set within the frequency range of 0.5-64 Hz,with eight specific frequency conditions selected for experimentation.The experimental results indicate that the thruster operates steadily within a voltage range of±1470 to±1920 V,with corresponding positive polarity current ranging from 0.41 to 4.91μA and negative polarity current ranging from−0.49 to−4.10μA.During voltage polarity switching,an emission delay occurs,manifested as a prominent peak signal caused by circuit capacitance characteristics and a minor peak signal resulting from liquid droplets.Extended emission test was conducted at 16 Hz,demonstrating approximately 1 h and 50 min of consistent emission before intermittent discharge.These findings underscore the favorable impact of AC conditions within the 8-16 Hz range on the self-neutralization capability of the ionic liquid thruster.

Power transfer efficiency in an air-breathing radio frequency ion thruster

Due to a series of challenges such as low-orbit maintenance of satellites, the air-breathing electric propulsion has got widespread attention. Commonly, the radio frequency ion thruster is favored by low-orbit missions due to its high specific impulse and efficiency. In this paper, the power transfer efficiency of the radio frequency ion thruster with different gas compositions is studied experimentally, which is obtained by measuring the radio frequency power and current of the antenna coil with and without discharge operation. The results show that increasing the turns of antenna coils can effectively improve the radio frequency power transfer efficiency, which is due to the improvement of Q factor. In pure N_2 discharge,with the increase of radio frequency power, the radio frequency power transfer efficiency first rises rapidly and then exhibits a less steep increasing trend. The radio frequency power transfer efficiency increases with the increase of gas pressure at relatively high power, while declines rapidly at relatively low power. In N_(2)/O_(2) discharge, increasing the N_(2) content at high power can improve the radio frequency power transfer efficiency, but the opposite was observed at low power. In order to give a better understanding of these trends, an analytic solution in limit cases is utilized, and a Langmuir probe was employed to measure the electron density. It is found that the evolution of radio frequency power transfer efficiency can be well explained by the variation of plasma resistance, which is related to the electron density and the effective electron collision frequency.

Study on Characteristics of a High-Precision Cold Gas Micro Thruster

In order to improve the reliability of the spacecraft micro cold gas propulsion system and realize the precise control of the spacecraft attitude and orbit, a micro-thrust, high-precision cold gas thruster is carried out, at the same time due to the design requirements of the spacecraft, this micro-thrust should be continuous working more than 60 minutes, the traditional solenoid valve used for the thrusts can’t complete the mission, so a long-life micro latching valve is developed as the control valve for this micro thruster, because the micro latching valve can keep its position when it cuts off the outage. Firstly, the authors introduced the design scheme and idea of the thruster. Secondly, the performance of the latching valve and the flow characteristics of the nozzle were simulated. Finally, from the experimental results and compared with the numerical study, it shows that the long-life micro cold gas thruster developed in this paper meets the mission requirements.

离子液体电喷雾推力器发射极制造技术现状及展望

离子液体电喷雾推力器是目前热点研究的一种静电式微电推进装置,影响该类推力器性能的核心部件之一是发射极,其制造技术已成为离子液体电喷雾推力器研制的关键技术之一。结合离子液体电喷雾推力器工作原理和发展历程,分析了毛细管型、外部浸润型和多孔材料型三类发射极的推进剂输运特点和制造需求;回顾和梳理了三类发射极的典型制造材料与相关制造技术,总结评述了离子刻蚀等不同制造技术的优缺点;针对较为成功的基于多孔材料的超快激光制造技术,从发射极设计、新型材料制备、超快激光与材论文拓展料作用机理等角度提出了发展建议。

TM_(011) MPT等离子体和电场特征分析与调谐实验

微波等离子体推力器(microwave plasma thruster,MPT)属于电热型推力器,其圆柱腔内等离子体过程、微波场分布与TM011模谐振状态是影响性能的重要因素.针对前人研究的可连续调节谐振状态的千瓦级MPT,需要开展结构固定的MPT研究,使其调谐过程简单、谐振状态良好,为深入研究奠定基础.本文对结构固定的千瓦级MPT圆柱腔内等离子体过程进行分析,探寻最佳放电条件.计算腔体内TM011模谐振状态下的微波电场和功率密度分布,分析其影响因素.对圆柱腔进行精细调谐实验,研究圆柱腔尺度和微波耦合探针尺度对谐振状态的影响.理论分析和数值计算结果发现489 Pa压强条件下氦气放电消耗功率最低,长径比大于1的圆柱腔内电场分布规律有利于气体放电.调谐实验结果发现长度和半径适中的球形探针使最短圆柱腔TM011谐振状态最佳、谐振频率最接近工作频率2.45 GHz.实验证明该腔体及匹配的探针使微波功率利用率高、氦气易放电,其结构方案具有正确性和可靠性.

离子电推进发展历史回顾与启示(上)

离子电推进是空间电推进中最具代表性的技术之一,追溯离子电推进从概念到工程应用的历史过程,能够为洞见离子电推进技术的未来发展提供有力依据。从早期孕育、初期发展、中期发展、当代发展和创新发展等五个方面,对世界各国离子电推进技术的发展历史进行了包括技术发展、产品研发、工程应用等内容的全面回顾。在此基础上,将离子电推进从初始概念到工程应用经历的漫长过程分为五个方面系统总结了离子电推进发展的成功经验和失败教训,从正确认识离子电推进等七个方面梳理了对离子电推进未来发展具有借鉴意义的启示。

霍尔推力器中电子碰撞及等离子体密度和磁场梯度激发的不稳定性

在霍尔推力器中,电子漂移、电子碰撞,以及等离子体密度、温度、磁场梯度所蕴含的自由能会驱动各种频率和波长的不稳定性.不稳定性的存在会破坏等离子体的稳定放电,削弱推力器与电源处理单元的匹配度,降低推力器的性能.基于此,本文利用基于流体模型推导的色散关系研究了霍尔推力器中由电子碰撞、等离子体密度和磁场梯度驱动的不稳定性.结果表明:1)在考虑电子惯性、电子与中性原子的碰撞,以及电子E×B漂移时能够在推力器近阳极区到羽流区内的任一轴向位置处激发不稳定性.随着角向波数ky的增大(k=2π/λ,λ为波长),模式将从由碰撞激发的低杂波不稳定性转变为离子声波不稳定性.当ky=10 m−1时,最大增长率γ_(max)对应频率ωr随着碰撞频率νen的增大而轻微减小;当ky=300 m−1时,γ_(max)对应的频率ωr以及最大频率ω_(rmax)几乎不随碰撞频率变化.不依赖于ky的大小,对于碰撞激发的不稳定性,模式的增长率随着碰撞频率的增大而增大.2)同时考虑电子惯性、电子碰撞效应,以及密度梯度时,密度梯度对驱动不稳定性占主导作用.模式的动力学行为不会随ky的增大而变化,但模式的本征值随ky的增大而增大.在密度梯度κN=0的两侧,由于密度梯度引起的抗磁性漂移频率ωs的符号发生了变化,模式的本征值在κN=0两侧有相反的变化趋势:当ω∗与ωr符号相反时,密度梯度对不稳定性的激发有削弱作用(κN>0);当ω∗与ωr符号相同时,密度梯度对不稳定性的激发有增强作用(κN<0).3)在模型中同时考虑等离子体密度梯度、磁场梯度,以及电子惯性和碰撞效应时,模式本征值的变化依赖于电子的漂移频率,以及密度和磁场梯度引起的抗磁性漂移频率的相对大小.当仅包含密度梯度和磁场梯度时,推力器放电通道内将出现稳定窗,即增长率为0的区间;包含电子惯性和碰撞效应后,稳定窗消失.

Catalytic bed slenderness ratio and ADN/methanol ratio for decomposition and combustion characteristics within ammonium dinitramide(ADN)-based green aerospace thruster

The decomposition and combustion characteristics of ammonium dinitramide (ADN) based non-toxic aerospace propellant are analytically studied to determine the effects of catalytic bed structure (slenderness ratio) and operation parameters (mass fraction ratio of ADN/CH3OH) on the general performance within the ADN-based thruster. In the present research, the non-equilibrium temperature model is utilized to describe the heat transfer characteristics between the fluid phase and solid phase in the fixed bed. We determined the fluid resistance characteristics in the catalytic bed by experiments involving the method of pressure-mass. We have done the simulation study based on the available results in the literature and found the complex physical and chemical processes within the ADN thruster. Furthermore, an optimized catalytic bed slenderness ratio was observed w让h a value of 1.75 and the mass fraction ratio of 5.73 significantly influenced the propellant performance. These results could serve as a reference to explore the combustion characteristics within the thruster and the preparation of future propellants.

Particle Simulation of Three-Grid ECR Ion Thruster Optics and Erosion Prediction

Ion optics is a critical component of ion thrusters. A two-dimensional axisymmetric model is developed to study the characteristics of three-grid electron cyclotron resonance ion thruster optics. The code is based on a particle-in-cell combined with the Monte Carlo collision method to simulate ion dynamics and charge-exchange processes in the grid region. The simulation results show that the mode can give a reasonable estimate of the physics characteristics of the ion optics. The design of the ion optics satisfies the requirement of preventing electron backstreaming. Charge-exchange ions can cause damage to the grids, especially to the accelerator grid. ＇Barrel＇ erosion can increase the accelerator grid aperture radius at a rate of 1.91~ 10-11 m/s, while the decelerator grid plays an important role in reducing ＇pits-and-grooves＇ erosion.

Study of the key factors affecting the triple grid lifetime of the LIPS-300 ion thruster

In order to ascertain the key factors affecting the lifetime of the triple grids in the LIPS-300 ion thruster,the thermal deformation,upstream ion density and component lifetime of the grids are simulated with finite element analysis,fluid simulation and charged-particle tracing simulation methods on the basis of a 1500 h short lifetime test.The key factor affecting the lifetime of the triple grids in the LIPS-300 ion thruster is obtained and analyzed through the test results.The results show that ion sputtering erosion of the grids in 5 kW operation mode is greater than in the case of 3 kW.In 5 kW mode,the decelerator grid shows the most serious corrosion,the accelerator grid shows moderate corrosion,and the screen grid shows the least amount of corrosion.With the serious corrosion of the grids in 5 kW operation mode,the intercept current of the acceleration and deceleration grids increases substantially.Meanwhile,the cold gap between the accelerator grid and the screen grid decreases from 1 mm to 0.7 mm,while the cold gap between the accelerator grid and the decelerator grid increases from 1 mm to 1.25 mm after 1500 h of thruster operation.At equilibrium temperature with 5 k W power,the finite element method（FEM）simulation results show that the hot gap between the screen grid and the accelerator grid reduces to 0.2 mm.Accordingly,the hot gap between the accelerator grid and the decelerator grid increases to 1.5 mm.According to the fluid method,the plasma density simulated in most regions of the discharge chamber is 1?×?10;-8?×?10;m;.The upstream plasma density of the screen grid is in the range 6?×?10;-6?×?10;m;and displays a parabolic characteristic.The charged particle tracing simulation method results show that the ion beam current without the thermal deformation of triple grids has optimal perveance status.The ion sputtering rates of the accelerator grid hole and the decelerator hole are 5.5?×?10;kg s;and 4.28?×?10;kg s;,respectively,while after the thermal deformation of the triple grids,the ion beam current has over-perveance status.The ion sputtering rates of the accelerator grid hole and the decelerator hole are 1.41?×?10;kg s;and 4.1?×?10;kg s;,respectively.The anode current is a key factor for the triple grid lifetime in situations where the structural strength of the grids does not change with temperature variation.The average sputtering rates of the accelerator grid and the decelerator grid,which were measured during the 1500 h lifetime test in5 k W operating conditions,are 2.2?×?10;kg s;and 7.3?×?10;kg s;,respectively.These results are in accordance with the simulation,and the error comes mainly from the calculation distribution of the upstream plasma density of the grids.

Experimental Study on the Effects of Magnetic Field Configuration near the Channel Exit on the Plume Divergence of Hall Thrusters

Experimental investigations into the effects of the magnetic field configuration near the channel exit on the plume of Hall thrusters were conducted. The magnetic field configuration near the channel exit is characterized by the inclination angle between the magnetic field lines and the thruster radial direction. Different inclination angles were obtained by varying the current ratio in the coils. The plume divergence angles were measured by a dual-directed probe. The results showed that the plume divergence angle increased obviously with the increase in the magnitude of the inclination angle near the channel exit. Therefore, in order to optimize the magnetic field for reducing plume divergence, the magnitude of the inclination angle should be reduced as much as possible. It suggests that the magnetic field configuration near the channel exit is another important factor that affects plume divergence.

Influence of channel length on discharge performance of anode layer Hall thruster studied by particle-in-cell simulation

Hall thruster has the advantages of simple structure, high specific impulse, high efficiency, and long service life, and so on. It is suitable for spacecraft attitude control, North and South position keeping, and other track tasks. The anode layer Hall thruster is a kind of Hall thruster. The thruster has a longer anode area and a relatively short discharge channel. In this paper, the effect of the channel length on the performance of the anode layer Hall thruster is simulated by the PIC simulation method. The simulation results show that the change of the channel length has significant effect on the plasma parameters, such as potential and plasma density and so on. The ionization region mainly concentrates at the hollow anode outlet position, and can gradually move toward the channel outlet as the channel length decreases. The collision between the ions and the wall increases with the channel length increasing. So the proper shortening of the channel length can increase the life of the thruster. Besides, the results show that there is a best choice of the channel length for obtaining the best performance. In this paper, thruster has the best performance under a channel length of 5 mm.

PIC simulation of plasma properties in the discharge channel of a pulsed plasma thruster with flared electrodes

Plasma in the discharge channel of a pulsed plasma thruster(PPT) with flared electrodes is simulated by a self-developed two-dimensional code. The fully particle-in-cell method with Monte Carlo collision is employed to model the particle movement and collisions and investigate the plasma properties and acceleration process. Temporal and spatial variations of the electron density distribution and the ion velocity between electrodes are calculated and analyzed in detail.The computational results of the electron number density, which is in the order of 1023 m-3,show good agreements with experimental results of a PPT named ADD SIMP-LEX. The ion velocity distributions along the center line of the channel lead to a comprehensive understanding of ions accelerated by electromagnetic field. The electron distributions of PPT with discharge voltages varying from 1300 to 2000 V are compared. The diffusion of electrons presents strong dependency on discharge voltage and implies higher degree of ionization for higher voltage.

Weak thruster fault detection for AUV based on stochastic resonance and wavelet reconstruction

When the bi-stable stochastic resonance method was applied to enhance weak thruster fault for autonomous underwater vehicle(AUV), the enhancement performance could not satisfy the detection requirement of weak thruster fault. As for this problem, a fault feature enhancement method based on mono-stable stochastic resonance was proposed. In the method, in order to improve the enhancement performance of weak thruster fault feature, the conventional bi-stable potential function was changed to mono-stable potential function which was more suitable for aperiodic signals. Furthermore, when particle swarm optimization was adopted to adjust the parameters of mono-stable stochastic resonance system, the global convergent time would be long. An improved particle swarm optimization method was developed by changing the linear inertial weighted function as nonlinear function with cosine function, so as to reduce the global convergent time. In addition, when the conventional wavelet reconstruction method was adopted to detect the weak thruster fault, undetected fault or false alarm may occur. In order to successfully detect the weak thruster fault, a weak thruster detection method was proposed based on the integration of stochastic resonance and wavelet reconstruction. In the method, the optimal reconstruction scale was determined by comparing wavelet entropies corresponding to each decomposition scale. Finally, pool-experiments were performed on AUV with thruster fault. The effectiveness of the proposed mono-stable stochastic resonance method in enhancing fault feature and reducing the global convergent time was demonstrated in comparison with particle swarm optimization based bi-stochastic resonance method. Furthermore, the effectiveness of the proposed fault detection method was illustrated in comparison with the conventional wavelet reconstruction.