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.

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.

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.

A study of the influence of different grid structures on plasma characteristics in the discharge chamber of an ion thruster

The grid structure has significant effects on the discharge characteristics of an ion thruster.The discharge performances of a 30 cm diameter ion thruster with flat,convex and concave grids are studied.The analysis results show that the discharge chamber with a convex grid has a larger’magnetic-field free area’than the others,and the parallelism of the magnetic-field isopotential lines and anode is generally the same in the three models.Plasma densities of the three structures at the grid outlet are in the range of 3.1×1016-6.9×1017m-3.Along the thruster axis direction,the electron temperature in the chamber with the convex and concave grids is in the range of 3.3-3.5 eV,while that with a flat grid is lower,in the range of 3.1-3.5 eV.In addition,the convex and the concave grids have better uniform distribution of electron temperature.Moreover,the collision frequency ratios show that the axial degree of ionization of the three models is the highest,and the flat grid has the highest discharge efficiency,followed by the convex grid and the concave grid is the least efficient.The test and simulation results of the 30 cm diameter ion thruster with the convex grid show that the measurement and calculation results are 3.67 A and 3.44 A,respectively,and the error above mainly comes from the ignorance of the doubly charged ions and parameter settings in the model.The comparison error between the simulation and measurement of beam current density is mainly caused by the actual thermal deformation of the grids during the discharge process,which leads to the change in electric potential distribution and variation of the focusing characteristics of the grids.Upon consideration of discharge performance and the thermal grid gap variation,it can be concluded that the flat and concave grids are more suitable for small-diameter ion thrusters,while the convex grid is a more reasonable choice for the higher-power and larger-diameter thrusters.

Influence of neutral gas supply position on wall erosion of Hall thruster studied by particle-in-cell simulation

In this work,we have carried out a simulation study on the discharge process of Hall thrusters under the conditions of different neutral gas radial supply positions based on the particle-in-cell(PIC)and Monte Carlo collision(MCC)methods.This paper compares the two-dimensional(2D)distributions of neutral gas,plasma and wall erosion-related parameters under different neutral gas supply positions.The comparison results show that the change of the neutral gas supply position affects the radial distribution uniformity of the neutral gas and plasma in the channel.From the comparison of the density peaks,it can be found that the neutral gas density and the plasma density peak under the upper gas supply condition are relatively low,and the plasma density peak is 22.49%lower than the density peak under the middle gas supply condition.Meanwhile,as the radial position of the gas supply moves from the lower gas supply to the upper gas supply,the position of the ionization zone also gradually moves toward the anode.The results of erosion-related parameter distribution comparison show that the change of gas supply location has an obvious influence on erosion rate and erosion range.In terms of erosion rate,the wall erosion rate is relatively low under the upper gas supply condition,and the peak erosion rates of the inner and outer walls are 33.3%and 29.9%lower than those under the other two conditions.In terms of erosion range,as the gas supply position moves from the lower gas supply position to the upper gas supply position,the erosion range gradually increases from5 to 7.5 mm.

Experimental and numerical simulation study of the effect for the anode positions on the discharge characteristics of 300W class low power Hall thrusters

Low-power Hall thruster(LHT) generally has poor discharge efficiency characteristics due to the large surface-to-volume ratio.Aiming to further refine and improve the performance of 300 W class LHT in terms of thrust and efficiency,and to obtain the most optimal operating point,the experimental study of the discharge characteristics for three different anode positions was conducted under the operation of various discharge voltages(100-400 V) and anode mass flow rates(0.65 mg·s-1and 0.95 mg·s-1).The experimental results indicated that the thruster has the most excellent performance in terms of thrust and efficiency etc at a channel length of 27 mm for identical operating conditions.In addition,particle in cell simulations,employed to reveal the underlying physical mechanisms,show that the ionization and acceleration zone is pushed downwards towards the channel exit as the anode moves towards the exit.At the identical operating point,when the channel length is reduced from 32 to 27 mm,the ionization and acceleration zone moves towards the exit,and the parameters such as thrust and efficiency increase due to the high ionization rate,ion number density,and axial electric field.When the channel length is further moved to 24 mm,the parameters in terms of thrust(F) and efficiency(ηa) are reduced as a result of the decreasing ionization efficiency(ηm) and the larger plume divergence angle(α).In this paper,the results indicated that an optimum anode position(ΔL=27 mm) exists for the optimum performance.

Experimental study of a neutralizer-free gridded ion thruster using radio-frequency self-bias effect

An experimental study on the quasi-neutral beam extracted by a neutralizer-free gridded ion thruster prototype was presented.The prototype was designed using an inductively coupled plasma source terminated by a double-grid accelerator.The beam characteristics were compared when the accelerator was radio-frequency(RF)biased and direct-current(DC)biased.An RF power supply was applied to the screen grid via a blocking capacitor for the RF acceleration,and a DC power supply was directly connected to the screen grid for the DC acceleration.Argon was used as the propellant gas.Furthermore,the characteristics of the plasma beam,such as the floating potential,the spatial distribution of ion flux,and the ion energy distribution function(IEDF)were measured by a four-grid retarding field energy analyzer.The floating potential results showed that the beam space charge is compensated in the case of RF acceleration without a neutralizer,which is similar to the case of classical DC acceleration with a neutralizer.The ion flux of RF acceleration is 1.17 times higher than that of DC acceleration under the same DC component voltage between the double-grid.Moreover,there are significant differences in the beam IEDFs for RF and DC acceleration.The IEDF of RF acceleration has a widened and multipeaked profile,and the main peak moves toward the high-energy region with increasing the DC self-bias voltage.In addition,by comparing the IEDFs with RF acceleration frequencies of3.9 and 7.8 MHz,it is found that the IEDF has a more centered main peak and a narrower energy spread at a higher frequency.

Endurance-test and theoretical prediction of a rare earth nanocathode for the applied field magnetoplasmadynamic thruster

The erosion loss of cathode is essential for the lifetime of magnetoplasmadynamic thruster(MPDT).In this work,an endurance test system for MPDT cathodes was designed and developed,and the erosion characteristics,erosion rate and erosion mechanism of the cathode were studied using the system under vacuum condition.The WCe20 hollow cathode was selected to carry out the long-term erosion of 540 h with the argon propellant supply flow rate of40 ml min^(-1),the input current of 25 A,and the central magnetic field intensity of 96 Gs.In order to predict the theoretical service life of cathode,a steady state erosion numerical model was established.The calculation results show that the total erosion rate of sputtering and evaporation is 11.58 mg h^(-1),which is slightly smaller than the test data of the average cathode corrosion rate of 12.70 mg h^(-1) in the experiment,because the experimental value includes start-up erosion rate.

A study of grid failure mode drivers and methods for accelerated life testing of a 30cm diameter ion thruster

In view of the high cost caused by the 1:1 lifetime verification test of ion thrusters,the lifetime acceleration test should be considered.This work uses the PIC-MCC(Particle-in-Cell MonteCarlo Collision)method to analyze the five failure factors that lead to the failure of the accelerator grid of a 30 cm diameter ion thruster under the working mode of 5 k W.Meanwhile,the acceleration stress levels corresponding to different failure factors are obtained.The results show that background pressure has the highest stress level on the grid's erosion.The accelerator grid aperture's mass sputtering rate under the rated vacuum degree(1×10^(-4)Pa)of 5 k W work mode is 8.78 times that of the baseline vacuum degree(1×10^(-6)Pa),and the mass sputtering rate under worse vacuum degree(5×10^(-3)Pa)is 5.08 times that of 1×10^(-4)Pa.Under the influence of the other four failure factors,namely,the voltage of the accelerator grid,upstream plasma density,the screen grid voltage and mass utilization efficiency,the mass sputtering rates of the accelerator grid hole are 2.32,2.67,1.98 and 2.51 times those of the accelerator grid hole under baseline condition,respectively.The ion sputtering results of two 30 cm diameter ion thrusters(both installed with new grids assembly)after working for 1000 h show that the mass sputtering rate of the accelerator grid hole under vacuum conditions of 5×10^(-3)Pa is 4.54 times that under the condition of 1×10^(-4)Pa,and the comparison error between simulation results and test results of acceleration stress is about 10%.In the subsequent ion thruster lifetime verification,the working vacuum degree can be adjusted according to the acceleration stress level of background pressure,so as to shorten the test time and reduce the test cost.

Numerical study of the radio-frequency biased accelerating system in ion thrusters

A 2D-3V implicit immersed-finite-element particle-in-cell(IFE-PIC)model is introduced to investigate the radio-frequency(RF)self-bias accelerating system applied in the RF ion thruster.A set of holes in a two-grid system with slit apertures is simulated in Cartesian coordinates.The characteristics of the plasma plume,such as the ion density,the neutralization rate and the ion and electron current density were investigated for different RF voltage amplitudes(600-1200V)and frequencies(6-30 MHz).Furthermore,the performance of the thruster was also carefully studied.The simulation results show that a well-focused plasma beam can be formed when the voltage amplitude is larger than 900 V and the frequency exceeds the reciprocal of ion transit time(≥12 MHz)in our simulation cases.The performance of the system can be evidently improved by increasing the voltage amplitude and the frequency,and the losses of the particle and thrust are reduced correspondingly.The bulk region of the plasma beam downstream shows good quasi-neutrality,and the ions are dominant in the peripheral region when a well-focused state is achieved.The high ion density beamlet in the periphery of the ion beam is closer to the axis when the voltage amplitude is increasing,while it is expanded radially when increasing the frequency.Backstream electrons have been observed upstream,and this mainly occurs in the phase in which the electron cannot escape.

Cathode position effects on microwave discharge cusped field thruster

The microwave discharge cusped field thruster is a novel concept of electric micropropulsion device,which operatesμN level thrust in low mass flow rate conditions,making use of a coaxial transmission line resonator.With its advantages of low thrust noise and high thrust resolution over a wide range of thrust,the thruster has emerged as a candidate thruster for the space-borne gravitational wave detection mission.The cathode effects commonly exist in many kinds of electric propulsion,and they are typically significant in micropropulsions.In order to find out the cathode position effects on a microwave discharge cusped field thruster,a thermionic cathode is mounted on a cross-slider for coupling.Under different cathode positions,the plume is analyzed by a Faraday probe and a retarding potential analyzer to analyze the performance and discharge characteristics.The results show that the magnetic mirror effect leads to significant degradation of anode current and an increase in low-energy ion ratio as the cathode moves away from the thruster exit.The electron conduction route also significantly impacts anode current efficiency,related to the cathode-exit distance and the thruster magnetic topology.

Numerical study of viscosity and heat flux role in heavy species dynamics in Hall thruster discharge

A two-and three-dimensional velocity space axisymmetric hybrid-PIC model of Hall thruster discharge called Hybrid2D has been developed.The particle-in-cell(PIC) method was used for neutrals and ions(heavy species),and fluid dynamics on a magnetic field-aligned(MFA) mesh was used for electrons.A time-saving method for heavy species moment interpolation on a MFA mesh was developed.The method comprises using regular rectangle and irregular triangle meshes,connected to each other on a pre-processing stage.The electron fluid model takes into account neither inertia terms nor viscous terms and includes an electron temperature equation with a heat flux term.The developed model was used to calculate all heavy species moments up to the third one in a stationary case.The analysis of the viscosity and the heat flux impact on the force and energy balance has shown that for the calculated geometry of the Hall thruster,the viscosity and the heat flux terms have the same magnitude as the other terms and could not be omitted.Also,it was shown that the heat flux is not proportional to the temperature gradient and,consequently,the highest moments should be calculated to close the neutral fluid equation system.At the same time,ions can only be modeled as a cold non-viscous fluid when the sole aim of modeling is the calculation of the operating parameters or distribution of the local parameters along the centerline of the discharge channel.This is because the magnitude of the viscosity and the temperature gradient terms are negligible at the centerline.However,when a simulation’s focus is either on the radial divergence of the plume or on magnetic pole erosion,three components of the ion temperature should be taken into consideration.The non-diagonal terms of ion pressure tensor have a lower impact than the diagonal terms.According to the study,a zero heat flux condition could be used to close the ion equation system in calculated geometry.

Study of beam divergence and thrust vector eccentricity characteristics of the Hall thruster based on dual Faraday probe array planes and its applications

The accurate knowledge of the thrust vector eccentricity and beam divergence characteristics of Hall thrusters are of significant engineering value for the beneficial integration and successful application of Hall thrusters on spacecraft.For the characteristics of the plume bipolar diffusion due to the annular discharge channel of the Hall thruster,a Gaussian-fitted method for thrust vector deviation angle and beam divergence of Hall thrusters based on dual Faraday probe array planes was proposed in respect of the Hall thruster beam characteristics.The results show that the ratios of the deviation between the maximum and minimum values of the beam divergence angle and the thrust vector eccentricity angle using a Gaussian fit to the optimized Faraday probe dual plane to the mean value are 1.4%and 11.5%,respectively.The optimized thrust vector eccentricity angle obtained has been substantially improved,by approximately 20%.The beam divergence angle calculated using a Gaussian fitting to the optimized Faraday probe dual plane is approximately identical to the non-optimized one.The beam divergence and thrust vector eccentricity angles for different anode mass flow rates were obtained by averaging the beam divergence and thrust vector eccentricity angles calculated by the dual-plane,Gaussian-fitted ion current density method for different cross-sections.The study not only allows for an immediate and effective tool for determining the design of thrust vector adjustment mechanisms of spacecraft with different power Hall thrusters but also for characterizing the 3D spatial distribution of the Hall thruster plume.

叶片数对无轴泵喷推进器性能的影响

为丰富无轴泵喷推进器噪声性能优化理论,研究了叶片数对无轴泵喷推进器性能的影响,重点探究了不同叶片数下无轴泵喷推进器内的压力脉动变化规律。基于计算流体力学中的有限体积法离散RANS方程,应用RNG k-ε湍流模型封闭计算模型,采用SIMPLEC进行压力-速度耦合,对不同叶片数下的无轴泵喷推进器内部流场进行了非定常数值模拟计算,并对计算结果进行了详细的对比分析。研究结果表明:随叶片数的增加,无轴泵喷推进器模型的扬程、效率和推力均逐渐增大,叶片表面压力分布更加均匀,叶轮轮缘和中部的压力脉动均明显降低,叶轮进出口主频处的压力脉动幅值从轮缘向中心逐渐减小。

矢量泵喷推进潜航器回转性能研究

泵喷推进器已经逐渐成为现代潜航器的首选,矢量推进器也在航空航天领域得到了广泛的应用。为了解决传统依赖船舵控制在低航速下控制效率低的问题以及提高潜航器回转性能,矢量推进器在潜航器上的应用已逐步成为国内外关注的热点。本文考虑大舵角(喷管偏角)下潜航器各参数非线性的影响,建立潜航器水平面操纵运动的非线性模型。通过仿真研究分析船-舵-推进器、船-矢量推进器和船-舵-矢量推进器三种不同的控制方法下潜航器在低航速、控制泵喷推进器转速、控制潜航器轴向航速三种不同条件下作水平面回转运动时的回转性能,结果表明矢量泵喷推进器可以有效改进潜航器的回转性能。

离子推力器溅射沉积物防剥离放电室的设计及激光选区熔化增材制造研究

为了抑制离子推力器寿命末期由溅射沉积导致的放电室膜层剥落物脱落,首先采用理论研究和试验的方法开展了剥落物抑制放电室的设计,基于零件结构和增材制造特点开展了零件工艺性分析并提出解决方案,实现了放电室增材制造,开展了离子推力器放电室溅射物防脱落性能测试。结果表明:放电室内表面设计纹理化结构排布形式为正六边形,凸起部分外接圆直径0.5 mm,间距0.25 mm,高度0.5 mm时可降低溅射沉积物脱落概率及碎屑尺寸,通过分层旋转扫描,选择激光功率425 W,扫描速度1200 mm/s工艺参数可实现大尺寸薄壁、具有复杂微小特征钛合金回转体制造,通过结构拓扑,镂空点阵结构、工艺参数和空间摆放位置优化,成型零件壁厚一致性偏差小于0.10 mm。放电室优化前后推力器性能稳定,各工作点关键参数相对误差小于4%。本文的研究为离子推力器放电室溅射沉积物脱落抑制方法和复杂结构的激光选区熔化增材制造成型提供了工程应用方案。

低功率平面型霍尔推力器的工作特性研究

为了解决低功率霍尔推力器存在的寿命问题,本文设计了一种平面型霍尔推力器。该推力器取消了内部放电通道,前壁和阳极在同一平面,电离和加速过程完全发生在推力器外部的开放空间中。推力器在0.6~1.0 mg/s的氙气流量下,实现稳定放电。本文实验研究了不同运行参数下推力器的阳极电流特性、推力特性和羽流束流特性。结果表明,阳极电流的振荡频谱主要集中在20~60 kHz,与常规霍尔推力器的呼吸振荡频率范围相似。在62~245 W的功率下,推力为3.2~10.5 mN,阳极效率为8%~22%。推力水平与同功率等级的传统霍尔推力器相当,而阳极效率偏低。利用法拉第探针获得离子的束流特性,结果表明:较低的羽流发散效率是造成目前推力器阳极效率偏低的主要因素。结合实验结果,分析了平面型霍尔推力器与传统霍尔推力器在离子加速方面的差异。本文设计的平面型霍尔推力器为解决霍尔推力器的壁面侵蚀问题提出了一种可能的方案。

一种配置矢量调节机构的全电推卫星动力学模型

为建立全电推进卫星的精确模型、设计全电推卫星的姿轨控方法及提高卫星的控制精度,针对一种配置可展开矢量调节机构的全电推卫星开展动力学模型研究.建立四关节可展开矢量调节机构的模型,推导基于可展开式矢量调节机构构型的全电推卫星的电推力及力矩模型,在此基础上建立全电推卫星的姿态轨道动力学方程.利用此模型描述全电推进卫星的姿态运动状态及轨道变化,能够准确反映出卫星的运动性能,该模型已应用于我国亚太6E卫星研制中,应用该模型对卫星进行的动力学仿真结果符合测试及分析结果,且该卫星在轨控制效果与预期一致,验证了模型的有效性.

基于并联机构的AUV矢量推进器研究

本文提出一种PUS-PRS-S并联机构形式的螺旋桨矢量推进器。对矢量推进器进行具体的结构设计,使其能够很容易地作为一个整体安装在AUV尾部。建立矢量推进器的的运动学和动力学模型,通过数值计算与仿真,结果表明,设计的矢量推进器具有良好的运动性能,能够流畅地完成俯仰、偏摆运动。通过对推力矢量的分析,可知矢量推进器在AUV前进方向上始终能提供不低于75%的推力,在偏转方向上最大能提供50%的推力,且推力大小与AUV姿态和航速均无关。对比与传统的以鳍舵作为控制面的AUV,矢量推进AUV在慢速航行状态下具有更好的操纵性。