Diagnostics of Air-Breathing Electric Thrusters

Air-Breathing Electric Thrusters (ABET) diagnostics is addressed, with on-ground prototypes in mind. It is based on detailed volume averaged Global Models and focuses on emission spectroscopy. Notably, the obtained optical emission spectroscopy diagnostics tools give important information about the thruster propellant constitution and also the ionization degree of each constituent. This allows for trade-off between various prototypes on the basis of their characterization and optimization.

Electric Thrusters Redundancy Configuration Strategy Study for All Electric Propulsion Platform Station-Keeping

Electric propulsion is used for all electric propulsion satellites to perform the orbit transfer,attitude control and station-keeping tasks. Generally electric propulsion subsystem contains 4 thrusters. But if one thruster fails in the beginning of satellite lifetime,other thrusters will undertake all the firing tasks. The firing time will be 2 to 3 times of thrusters without failure. Thus it may go beyond the allow ed lifetime of thruster. This paper puts forward two thruster redundancy configuration solutions with 6 thrusters to solve this problem. Two layout configurations and their corresponding station-keeping strategies are simulated and compared. The results show that the maximum firing time of both layout configurations can meet the lifetime limitation. This solution is a good reference for all electric propulsion satellites design.

Study on High Power Hall Electric Propulsion Technology

High power Hall electric propulsion technology is a very competitive electric propulsion technology for future large space missions such as large GEO satellites,manned space programs,deep space explorations,cargo ships,space tugs.Based on the experience of more than 20 years in research and development of Hall electric propulsion,the Shanghai Institute of Space Propulsion(SISP)has developed 3 high power Hall thrusters,i.e.,the 10 k W class HET-500,20 k W class HET-1000,and 50 k W class HET-3000.This paper presents the development status of the high power(≥10 k W)Hall electric propulsion at SISP,including tests of 3 high power Hall thrusters in the power range from 10 k W to 50 k W,the qualification of a single string of a 10 k W Hall electric propulsion system,and the study of a cluster of two 1.35 k W HET-80 Hall thrusters to understand the technical issues related to multi-thruster high power electric propulsion systems.

Machine learning-based method to adjust electron anomalous conductivity profile to experimentally measured operating parameters of Hall thruster

The problem of determining the electron anomalous conductivity profile in a Hall thruster,when its operating parameters are known from the experiment,is considered.To solve the problem,we propose varying the parametrically set anomalous conductivity profile until the calculated operating parameters match the experimentally measured ones in the best way.The axial 1D3V hybrid model was used to calculate the operating parameters with parametrically set conductivity.Variation of the conductivity profile was performed using Bayesian optimization with a Gaussian process(machine learning method),which can resolve all local minima,even for noisy functions.The calculated solution corresponding to the measured operating parameters of a Hall thruster in the best way proved to be unique for the studied operating modes of KM-88.The local plasma parameters were calculated and compared to the measured ones for four different operating modes.The results show the qualitative agreement.An agreement between calculated and measured local parameters can be improved with a more accurate model of plasma-wall interaction.

On the heating mechanism of electron cyclotron resonance thruster immerged in a non-uniform magnetic field

To study the heating mechanism of electron cyclotron resonance thruster(ECRT)immersed in a non-uniform magnetic field,experiments and simulations are performed based on an electron cyclotron resonance plasma source at ASIPP.It is found that the first harmonic of electron cyclotron resonance is essential for plasma ignition at high magnetic field(0.0875 T),while the plasma can sustain without the first and second harmonics of electron cyclotron resonance at low magnetic field(till 0.0170 T).Evidence of radial hollow density profile indicates that upper hybrid resonance,which has strong edge heating effect,is the heating mechanism of low-field ECRT.The heating mode transition from electron cyclotron resonance to upper hybrid resonance is also revealed.Interestingly,the evolutions of electron temperature and electron density with input power experience a‘delayed’jump,which may be correlated with the different power levels required for cyclotron and ionization.Moreover,when the field strength decreased,the variation of electron density behaves in an opposite trend with that of electron temperature,implying a possible competition of power deposition between them.The present work is of great interest for understanding the plasma discharge in ECRT especially immersed in a non-uniform magnetic field,and designing efficient ECRT using low magnetic field for economic space applications.

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.

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.

Three-dimensional particle simulation of ion thruster plume flows with EX-PWS

Ion thruster plumes from a multi-thruster array of different working configurations are simulated by a hybrid fluid-particle software. The particle in cell method is employed to model the transports of ions. The direct simulation Monte Carlo method is used to model momentum and charge exchange （CEX） collisions. The software is based on unstructured grids which make it easy to handle with complex geometry. The results of chamber simulation are compared with experimental data in ion current density and number density, which show good agreements. The maximum difference of current density along the thruster centerline is less than 9.30%. The interaction effects of plumes when multiple thrusters are operating in vacuum are predicted. Distributions of single charged xenon ions are significantly different in the near-field plume flow, however, merge into one in the far downstream region. Moreover, the interaction effect on the spatial distribution of CEX xenon ions is displayed as well.

Cathode erosion site distributions in an applied-field magnetoplasmadynamic thruster

Erosion can influence cathode life,and is thus considered to be one of the main factors limiting the application of applied-field magnetoplasmadynamic thrusters.In this paper,erosion sites on graphite cathodes are studied so as to identify the influence of applied magnetic field and the ratio of propellant mass flow rate supplied from cathode and anode.The experiment results show that the application of applied magnetic field can significantly reduce the erosion rate of the cathode compared to that without magnetic field.The erosion sites on the cathode vary with the relative position of the convergent-divergent magnetic field,and are mainly distributed in the divergent part of the field.The erosion sites on the cathodes are found to be related to the propellant supply.The decreasing anode mass flow rate enlarges the range of erosion.These results are much helpful for the analysis of cathode erosion site location since they provide evidences of erosion mechanisms and point out the directions for further research.

GEO Satellite Thruster Configuration and Optimization

A thruster configuration and optimization method for GEO satellite is proposed in this paper.All thrusters are installed on the back panel of satellite,which is not only compatible with the design of satellite subdivision structure,but also able to provide a large installation room for the payloads and components carried by the satellite.Chemical thruster and electric thruster were selected because they have complementary advantages and can effectively reduce propellant consumption.The working mode and tasks of the thruster were analyzed in detail for thruster configuration,which was optimized in order to minimize propellant consumption.Lastly,a GEO satellite thruster was configured and optimized using this method,with results showing that the method is feasible and effective,and consequently has value for engineering applications.

Numerical Simulation of Ion Extraction Through Ion Thruster Optics

Based on the particle-in-cell （PIC） method, a two-dimensional numerical scheme was developed to investigate the ion beam extraction phenomena through the ion thruster optics. According to the calculated results, the plasma sheath upstream of the screen grid, the electric field in the calculation domain, and the ion and electron spatial distributions are obtained for different accelerator grid voltages. The results indicate that the accelerator grid voltage affects the plasma sheath upstream of the screen grid significantly. It is found that a moderate accelerator grid voltage results in an ion optical performance better than either a higher or lower voltage, from a point of ion extraction from the discharge chamber and erosion mitigation of the accelerator grid due to the direct ion impingement.

Study on the Discharge Characteristics of a Coaxial Pulsed Plasma Thruster

A new coaxial pulsed plasma thruster （PPT） laboratory model is designed and employed in this study. A Teflon sleeve is connected with the anode, which is shaped as a nozzle, and a cathode is mounted in the cavity of the Teflon sleeve and kept in close contact with it. A thread is then designed in the internal surface of the Teflon sleeve, and because of the strong field strength of the cathode triple junction （CTJ）, vacuum flashover occurs and a plasma jet is acquired behind the anode. The electric field distribution of the designed coaxial PPT laboratory model is simulated by MAXWELL 3D simulation software, and the plasma density and thrust are measured by a Langnmir probe and a piezoelectric thin-film sensor, respectively. Through a series of comparative experiments, we discuss the impact of optimal designs, such as the thread and the nozzle-shaped anode, on the discharge characteristics of the coaxial PPT. The experimental and simulation results indicate that the designed coaxial PPT laboratory model presents better discharge characteristics in view of its higher plasma density and greater thrust.

The plume diagnostics of 30 cm ion thruster with an advanced plasma diagnostics system

In order to achieve a better understanding of plume characteristics of LIPS-300 ion thruster,the beam current density,ion energy and electron number density of LIPS-300 ion thruster plume are studied with an Advanced Plasma Diagnostics System(APDS)which allows for simultaneous in situ measurements of various properties characterizing ion thruster,such as plasma density,plasma potential,plasma temperature and ion beam current densities,ion energy distribution and so on.The results show that the beam current density distribution has a double‘wing'shape.The high energy ions were found in small scan angle,while low energy ions were found in greater scan angle.Electron number density has a similar shape with the beam current density distribution.

Development of HEP-XXMF Series Hall Thrusters in BICE

Hall thruster is a type of electric propulsion thruster which is often regarded as a moderate specific im- pulse space propulsion technology and is used primarily for station keeping and orbit maintenance tasks. Magnetic field is the most important aspect in Hall thruster＇s design. Each time the performance improvement of Hall thrusters is accompanied with the regeneration of magnetic field design. Now, all the maior space nations treat the magnetic field design as the key technology for Hall thrusters, and many Hall thrusters have emerged in recent years, for example, BPT-4000 and NASA- 173M. In China, based on magnetic focus technology, the Beijing Institute of Control Engineer- ing （BICE） has been developing Hall thrusters ranged from hundreds of watts to 5 kilowatts, including the IS00 W HEP- 100MF and the 5 kW HEP- 140blF Hall thrusters. This paper briefly reviews the development of the HEP-XXMF series Hall thruster. The principle of the Hall thruster, the design methods and the current status of HEP-100MF and HEP- 140MF are discussed in the paper.

Production of a High-Mach-Number Plasma Flow for an Advanced Plasma Space Thruster

A higher specific impulse and a larger thrust are required for a mannedinterplanetary space thruster. Prior to a realization of a fusion-plasma thruster, amagneto-plasma-dynamic arcjet (MPDA) powered by a fission reactor is one of the promising candidatesfor a manned Mars space thruster. The MPDA plasma is accelerated axially by a self-induced j x Bforce. Thrust performance of the MPDA is expected to increase by applying a magnetic nozzle insteadof a solid nozzle. In order to get a much higher thruster performance, two methods have beeninvestigated in the HITOP device, Tohoku University. One is to use a magnetic Laval nozzle in thevicinity of the MPDA muzzle for converting the high ion thermal energy to the axial flow energy. Theother is to heat ions by use of an ICRF antenna in the divergent magnetic nozzle. It is found thatby use of a small-sized Laval-type magnetic nozzle, the subsonic flow near the muzzle is convertedto be supersonic through the magnetic Laval nozzle. A fast-flowing plasma is successfully heated byuse of an ICRF antenna in the magnetic beach configuration.

Experimental investigation on the evolution of plasma properties in the discharge channel of a pulsed plasma thruster

Pulsed plasma thrusters(PPTs)are an attractive form of micro-thrusters due to advantages such as their compactness and lightweight design compared to other electric propulsion systems.Experimental investigations on their plasma properties are beneficial in clarifying the complex process of plasma evolution during the micro-second pulse discharge of a PPT.In this work,the multi-dimensional evolutions of the light intensity of the PPT plasma with wavelength,time,and position were identified.The plasma pressure was obtained using an iterative process with composition calculations.The results show that significant ion recombination occurred in the discharge channel since the line intensities of CII,CIII,CIV,and FII decreased and those of CI and FI increased as the plasma moved downstream.At the center of the discharge channel,the electron temperature and electron density were in the order of 10000 K and 1017 cm-3,respectively.These had maximum values of 13750 K and 2.3?×?1017 cm-3 and the maximum temperature occurred during the first half-cycle while the maximum number density was measured during the second half-cycle.The estimated plasma pressure was in the order of 105 Pa and exhibited a maximum value of 2.69?×?105 Pa.

The State of Arts of Hall Thruster

Electric propulsion is broadly defined as the acceleration of a working fluid for propulsion by electrical heating and/or by electric and magnetic body forces. Compared with chemical propulsion, electric propulsion has the characteristic of higher specific impulse, lower thrust, lighter weight and longer lifetime. So electric propulsion is generally suitable for satellite attitude control, the orbit transfer and raising, orbit correction, resistance compensate, position keeping, reposi- tion, space exploration and interplanetary flight.

国外离子推力器栅极几何形貌测试评价技术的应用与发展

栅极组件作为离子推力器的核心部件,对电推力器的性能、可靠性及寿命起着非常重要的作用。简述了离子推力器栅极几何形貌测试评价技术的主要研究内容,重点论述了国外航天技术先进国家在寿命试验中应用栅极几何形貌测试评价新技术的重要进展,总结了几何形貌测试评价技术的发展趋势。

光谱卫星LPPT-25微电推进系统飞行试验工作性能评价

基于光谱卫星任务需求,兰州空间技术物理研究所开展了脉冲等离子体微电推进系统方案设计,以此完成25W级脉冲等离子体电推进系统飞行样机研制,并开展了地面各项测试,均满足需求。为了进一步验证该系统的空间环境适应性、与航天器的相互兼容性、空间工作特性及空间飞行性能与地面数据的差异性,LPPT-25微电推进系统搭载长光卫星公司光谱星一号(GP-1)卫星开展了在轨飞行试验,对脉冲等离子体电推进系统在轨飞行试验结果进行了评价。结果表明:在整个飞行试验期间,脉冲等离子体电推进系统各项工作性能参数符合设计指标要求,电推进分系统工作正常,推力输出稳定,各遥测温度满足推进要求的控温范围;电推进推力标定为306.3μN,相比地面测试推力300μN,偏差在5%以内,体现了良好的天地一致性。

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.