Thrust measurement of dimethyl ether arcjet thruster

The present paper describes thrust measurement results for an arcjet thruster using Dimethyl ether (DME) as the propellant. DME is an ether compound and can be stored as a liquid due to its relatively low freezing point and preferable vapor pressure. The thruster successfully produced high-voltage mode at DME mass flow rates above 30 mg/s, whereas it yielded low-voltage mode below 30 mg/s. Thrust measurements yielded a thrust of 0.15 N and a specific impulse of 270 s at a mass flow rate of 60 mg/s with a discharge power of 1300 W. The DME arcjet thruster was comparable to a conventional one for thrust and discharge power.     

稳态等离子体电推进技术研究现状及其关键技术

研究表明,稳态等离子体推力器(SPT)具有比冲大、效率高、寿命长的优点,是具有较高性能的先进空间推进系统之一,已广泛应用于小卫星的姿态控制和轨道保持。概述了SPT的系统组成和工作原理,重点介绍了SPT的研究进展、主要性能参数、关键技术及其在潜在应用领域中的性能优势,指出其关键技术有空心阴极的热设计和结构设计以及弯曲磁场位形的设计。对我国SPT研究内容提出了建议。     

圆柱形Hall推力器磁场的设计与数值模拟

霍尔(Hall)推力器的磁场设计是提高其性能的关键技术之一。针对传统环形Hall推力器小型化带来的缺点,提出了圆柱形的Hall推力器。简要对比了环形和圆柱形两种推力器,对圆柱形推力器的工作原理、磁场设计的要求以及利用ANSYS有限元软件进行的电磁场数值模拟的结果进行了详细分析。结果表明,圆柱形推力器磁路设计是合理的,另外也证明了用ANSYS有限元软件进行电磁场模拟的可靠性。     

一种用于临近空间飞行器的吸气式电推进技术

针对低速临近空间飞行器提出了一种新型吸气式电推进方案,该方案采用单介质阻挡放电（SDBD）作为等离子体源,因此能在较大气压范围（数Pa~1atm）内电离大气产生等离子体并产生推力。为探究该吸气式电推进方案的推力性能,测量了实验样机在多个气压和电压条件下产生的推力。推力测量结果显示在10~90kPa气压范围内,实验样机产生的推力在10 2~10 3μN量级;气压一定时,产生的推力与驱动电压呈幂次相关;而电压一定时,随着气压自1atm逐渐降低,产生的推力先增大后减小,且达到最大推力的气压与所加驱动电压相关。     

Very high delta-V missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept

A new and innovative type of gridded ion thruster, the “Dual-Stage 4-Grid” or DS4G concept, has been proposed and its predicted high performance validated under an ESA research, development and test programme. The DS4G concept is able to operate at very high specific impulse and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. This makes it a possible candidate for ambitious missions requiring very high delta-V capability and high power. Such missions include 100 kW-level multi-ton probes based on nuclear and solar electric propulsion (SEP) to distant Kuiper Belt Object and inner Oort cloud objects, and to the Local Interstellar medium. In this paper, the DS4G concept is introduced and its application to this mission class is investigated. Benefits of using the DS4G over conventional thrusters include reduced transfer time and increased payload mass, if suitably advanced lightweight power system technologies are developed.A mission-level optimisation is performed (launch, spacecraft system design and low-thrust trajectory combined) in order to find design solutions with minimum transfer time, maximum scientific payload mass, and to explore the influence of power system specific mass. It is found that the DS4G enables an 8-ton spacecraft with a payload mass of 400 kg, equipped with a 65 kW nuclear reactor with specific mass 25 kg/kW (e.g. Topaz-type with Brayton cycle conversion) to reach 200 AU in 23 years after an Earth escape launch by Ariane 5. In this scenario, the optimum specific impulse for the mission is over 10,000 s, which is well within the capabilities of a single 65 kW DS4G thruster. It is also found that an interstellar probe mission to 200 AU could be accomplished in 25 years using a “medium-term” SEP system with a lightweight 155 kW solar array (2 kg/kW specific mass) and thruster PPU (3.7 kg/kW) and an Earth escape launch on Ariane 5. In this case, the optimum specific impulse is lower at 3500 s which is well within conventional gridded ion thruster capability.     

A ballistic-pendulum test stand to characterize small cold-gas thruster nozzles

This paper deals with the design, development and experimentation of a new test stand for the accurate and precise characterization of small cold-gas nozzles having thrust of the order of 0.1 N and specific impulse of the order of 10 s. As part of the presented research, a new cold-gas supersonic nozzle was designed and developed based on the quasi one-dimensional theory. The test stand is based on the ballistic-pendulum principle: in particular, it consists of a suspended gondola hosting the propulsion system and the sample nozzle. The propulsion system consists of an air tank, pressure regulator, solenoid valve, battery and digital timer to command the valve. The gondola is equipped with a fin, immersed in water, to provide torsional and lateral oscillation damping. A laser sensor measures the displacement of the gondola. The developed test stand was calibrated by using a mathematical model based on the inelastic collision theory. The obtained accuracy was of ～1%. Sample experimental results are reported regarding the comparison of the new supersonic nozzle with a commercially available subsonic nozzle. The obtained measurements of thrust, mass flow rate and specific impulse are precise to a level of ～3%. The broad goal of the presented research was to contribute to an upgraded design of a spacecraft simulator used for laboratory validation of guidance, navigation and control algorithms for autonomous docking manoeuvres.     

A performance comparison of ultrasonically aided electric propulsion extractor configurations

As a novel propulsion technology, ultrasonically aided electric propulsion (UAEP) offers a high specific impulse and a high thrust density. In this paper, the effects of extractor grid configuration on performance of a UAEP thruster have been investigated by both experimental studies and numerical simulation. Relationships between spray current and operation parameters, including applied voltage, propellant flow rate, and vibration power and frequency, are explored for different extractor mesh sizes and shapes. Numerical simulation is also carried out for a better understanding of the formation of capillary standing waves as well as the electric field distribution in the acceleration zone. Experimental results show that compared with a circular shaped extractor, a reticular shaped extractor is able to produce a higher spray current. The current density increases with a denser mesh, which agrees well with the numerical simulation results. This phenomenon indicates that optimizing extractors with appropriate shapes and sizes can be an effective way to improve the performance of a UAEP system. A performance evaluation based on hydrodynamic and electrostatic calculations indicates that the present UAEP system can produce a thrust competitive to that of the colloid thruster with an emitter array.     

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 impulse 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 major 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-173 M. In China, based on magnetic focus technology, the Beijing Institute of Control Engineering(BICE) has been developing Hall thrusters ranged from hundreds of watts to 5 kilowatts, including the 1500 W HEP-100 MF and the 5 k W HEP-140 MF 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-100 MF and HEP-140 MF are discussed in the paper.     

Endurance testing of a pulsed plasma thruster for nanosatellites

The mission complexity of Nanosatellites has increased tremendously in recent years, but their mission range is limited due to the lack of an active orbit control or ∆v capability. Pulsed Plasma Thrusters (PPT), featuring structural simplicity and very low power consumption are a prime candidate for such applications. However, the required miniaturization of standard PPTs and the adaption to the low power consumption is not straightforward. Most investigated systems have failed to show the required lifetime. The present coaxial design has shown a lifetime of up to 1 million discharges at discharge energies of 1.8 J in previous studies. The present paper focuses on performance characterizations of this design. For this purpose direct thrust measurements with a µN thrust balance were conducted. Thrust measurements in conjunction with mass bit determination allowed a comprehensive assessment. Based on those measurements the present µPPT has a total impulses capability of approximately I≈1.7 Ns, an average mass bit of 0.37 µg s⁻¹ and an average specific impulse of I_sp≈904 s. All tests have shown very good EM compatibility of the PPT with the electronics of the flight-like printed circuit board. Consequently, a complete µPPT unit can provide a ∆v change of 5.1 m/s or 2.6 m/s to a standard 1-unit or 2-unit CubeSat respectively.     

脉冲等离子体推力器内部流场的数值分析

应用所建立的广义拉格朗日乘子形式的磁流体力学模型对平行电极型脉冲等离子体推力器的内部流场进行了数值研究.对等离子体流动过程的分析表明,放电初期等离子体在很强的电磁力作用下高速喷出推力器,而到放电后期等离子体受到的气动力比电磁力大得多,等离子体的运动非常缓慢.电磁冲量占推力器元冲量的大部分,增强电磁加速作用可以进一步提高推力器的性能.     

离子液体推力器性能影响因素仿真研究

文章主要研究离子液体推力器发射阵列几何结构参数对推力器性能的影响。建立单个发射极粒子网格法仿真模型,模拟获得了单个发射极束流分布及推力、比冲等性能参数曲线;通过改变电极间距及引出极孔径获得了单个发射极平均推力和角向效率的变化曲线,分析了电极间距及引出极孔径对推力器性能的影响。研究表明:在离子被引出极拦截前,推力器角向效率与电极间距成正比,与引出极孔径成反比;在不考虑发射离子数量和比例改变情况下,推力器推力随电极间距的增加先增大后减小,一定范围内随引出极孔径的减小而增大,但过小的引出极孔径会造成严重的推力损失与栅极侵蚀。以上研究结果可为离子液体推力器的优化设计提供参考。     

双组元离心式喷注器10 N发动机偏工况试验

根据国内外同类发动机研制经验,双组元10 N发动机在入口压力为0.8~2.2 MPa范围内,入口压力偏差会使发动机真空比冲、燃气温度等性能产生较大变化。为了获得双组元离心式喷注器10 N发动机在落压推进系统要求的入口压力范围内性能,通过采用小流量喷雾试验台和42 km高模试验台,对偏工况条件下的冷态性能及热试性能进行试验研究。试验结果表明:该发动机额定入口压力1.58 MPa时真空比冲为2881 N·s/kg;当入口压力在0.6~2.5 MPa变化时,对应真空推力从4.7 N增加到14 N,落压比为3;入口压力0.6 MPa时真空比冲为2600 N·s/kg,入口压力2.5 MPa时真空比冲为2956 N·s/kg;入口压力在0.6~2.5 MPa试验范围内,发动机燃烧室壁温均低于材料许用温度,表明发动机热设计优良,可满足双组元落压推进系统对姿控发动机的性能需求。     

氨电弧喷射推力器实验和数值研究

为了证实氨电弧喷射推力器良好的工作性能，对其进行了系统的实验和数值研究。在实验研究中建立了高真空实验系统、先进的小推力测量系统和恒定电流控制系统，在数值研究中采用二阶精度无波动、无自由参数的耗散差分格式(NND格式)数值求解耦合电磁源项的扩展N-S方程组。通过实验获得了不同结构尺寸的推力器不同工作条件下的工作性能，通过数值模拟给出了推力器的流场结构，预计了其推力、比冲和推进效率。研究结果表明，氨电弧喷射推力器工作性能良好，能够担负一定的空间应用任务。     

Generation of experimental plasma parameter maps around the baffle aperture of a Kaufman (UK-25) ion thruster

The plasma produced by a Kaufman (UK-25) ion thruster possesses four distinct regions. The interaction of two of these regions, the coupling and discharge plasmas, is important for thruster design and efficiency, since it controls the production of the primary electrons responsible for most of the ionisation in the thruster. This paper details the measurement and analysis techniques used to obtain two-dimensional experimental plasma parameter maps across a plasma double layer in the baffle aperture region of this type of ion thruster. This plasma-interaction region was mapped experimentally using Langmuir probes.The Langmuir probe data were collected and analysed to produce maps of plasma potential, Maxwellian and primary electron energy and number density. Detailed static magnetic field measurements using Hall-effect probes were also obtained. A number of derived plasma parameter maps were then made possible, such as electron pressure. The application of the analysis techniques used here allowed ion thruster plasma property maps to be constructed of a spatial extent and resolution previously unseen. Two-dimensional maps of the spatial location of primary electrons are presented, as effected by applied magnetic field changes, along with the detection of regions showing depletion of primary electron energies.     

一种新颖的微空心阴极放电等离子体推力器

微小卫星的发展和成功应用迫切需要新型微推力器的研制。微放电技术是等离子体放电中重要的一类,近几十年来成为各国的研究热点。其中,微空心阴极放电(MHCD)是一种新颖的非平衡高气压辉光放电,其优点是可以在高气压下稳定放电,并且只需要非常低的电压(几百伏特)或者输入功率(百毫瓦数量级)。MHCD建立在2个几百微米厚度的金属平面电极上,材料可以是钼、铝等,由电介质(云母或氧化铝)隔开。"三明治"的布局结构上从一个电极到另一个电极钻有直径为几十微米到几百微米的孔,气体压强可以很高,甚至超过大气压。微空心阴极放电较小的尺寸结构与强烈并可控的气体加热相结合,可以开发应用在新型的电热式微等离子体推进上。由于微空心阴极放电等离子体推力器在微放电等离子体中加热了工质气体,随后通过微喷管喷出产生推力,因此与传统的冷气微推力器相比,可大大提高推力器的比冲和推力。     

Five millinewton field emission thruster testing

A compact cluster of two 80-mm-wide emitters has been designed, built and tested. The field emission thruster using this cluster is able to produce thrust in excess of 5 mN at high mass efficiency.     

A thrust formula for an MPD thruster with applied-magnetic field

In this paper a thrust formula for applied field MPD will be presented. The swirling velocity will be derived from the magnetic stress tensor and its conversion into axial energy into the magnetic nozzle will be analytically treated. The theoretical prediction of both swirling velocity and thrust will be compared to the measurements showing a reasonable agreement.     

航天器推力器测试台高速实时数据采集系统的设计与实现

为了满足航天器卫星推力器真空热试车工程背景的需要,在VC++环境下利用多线程技术设计航天器推力器试验台高速数据采集系统,本文详细的研究了设计中的关键性技术及解决方案,试验验证该方案是可行的,并有一定的通用性。     

A thrust formula for an MPD thruster with applied-magnetic field

In this paper a thrust formula for applied field MPD will be presented. The swirling velocity will be derived from the magnetic stress tensor and its conversion into axial energy into the magnetic nozzle will be analytically treated. The theoretical prediction of both swirling velocity and thrust will be compared to the measurements showing a reasonable agreement.     

月球探测器悬停控制技术的研究

月球探测器近月面定点悬停是月球探测器实现月面软着陆的必要保证。提出一种近月面定点悬停的控制方案,选用常推力液体发动机和脉冲式发动机实现悬停任务。对于常推力液体发动机,以燃耗最省为指标采用最优控制,使月球探测器悬停在特定高度上;对于脉冲式发动机采用PID控制,实现月球探测器悬停状态的保持。给出具体设计方法,仿真结果表明方案可行。