Experimental investigation of a low-power stationary plasma thruster

Integral characteristics of a laboratory model of the new-generation α-40 stationary plasma thruster (SPT) designed to operate at a total consumed power of 150 W have been studied. Integral parameters of the SPT prototype, including thrust, discharge current, specific impulse, and efficiency, determined on the test stands of MIREA and Fakel Special Design Bureau proved to be virtually identical. At a deposited power of ～150W, discharge voltage of 210 V, and consumption rate of propellant (xenon) of 0.7 mg/s, the thrust reached 9.4 mN, the anodic specific impulse was ～1370 s, and the anode thrust efficiency (with allowance for the power consumed by magnetization coils) was about 41%. In a different regime with a greater thrust (9.7 mN) but lower specific impulse (1240 s) at the same deposited power (～150 W), which was observed at a propellant-consumption rate of 0.8 mg/s and a discharge voltage of 190 V, the efficiency was ～40%. The jet semidivergence angle was ±25°. The predicted SPT service life is 2500 h. The obtained data show that the efficiency of α-40 SPT is higher than that of the known analogs.     

Directional acoustic underwater thruster

This study describes a tested prototype for a controllable directional underwater thruster with no moving parts. During operation, a high-intensity acoustic wave creates directional water jets and the device moves itself in the opposite direction. When the underwater thruster moves along a non-vertical angle, it can produce straight backward thrust of 2.3 mN and lateral thrust of 0.6 mN in parallel with the device surface, with a total thrust-to-weight ratio of 2:1. To enhance the acoustic streaming effect, a self-focusing acoustic transducer (SFAT) with air reflectors is used to focus the acoustic wave.     

Development and testing of a new thrust stand for micro-thrust measurement in vacuum conditions

A miniature thrust stand in man-made vacuum environment, based on the principle of leaf spring bending deformation, capable of supporting testing of thrusters having a total mass of up to 300 g and producing thrust levels between 1 μN and 760 μN has been developed and tested. A case-study measurement was conducted for a Colloid Micro-Newton Thruster (CMNT) to investigate the feasibility of the proposed system under vacuum conditions. Displacement of the leaf spring deforming was measured using a linear capacitive displacement sensor (LCDS). An in situ calibration rig allows for steady-state calibration before, during and after thruster operation. To validate this method, extensive thrust measurements were carried out on a colloid thruster that produces μN magnitude level thrust under vacuum conditions. Results show that the measurements were very repeatable, producing results that compare favorably with measured data with an uncertainty of approximately 10%.     

离子推力器推力矢量偏角测试

研制了离子推力器推力矢量偏角测试系统,包括测试装置的设计、制造、测试处理软件的设计和调试。应用研制的测试系统对20cm离子推力器的推力矢量偏角进行了测试。     

离子推力器束流下降原因分析

离子推力器的推力与其引出的束流成正比,束流的大小直接确定了推力.离子推力器在点火启动后,在工作条件不变的条件下,其引出束流随工作时间而下降.为找出束流下降的原因,以离子推力器为研究对象,通过分析引起离子推力器束流下降的各种因素,并对这些因素进行分析与验证.经过对因素的分析定位,找出引起束流下降的主要原因.分析与验证表明：影响离子推力器束流下降的决定因素为栅极组件固有特性、磁场固有特性和阴极固有特性,其中阴极固有特性是导致束流下降的主要原因.     

Experimental and numerical study of an electrothermal pulsed plasma thruster for small satellites

A pulsed plasma thruster (PPT) with a propellant feeding mechanism was designed using two poly tetrafluoroethylene (PTFE) bars as propellants. The PPT was mounted on a thrust stand with a 1-m-long perpendicular pendulum which was developed for precise measurements of impulse bits. Initial thrust performance showed thrust-to-power ratio of 43–48 μN/W, specific impulse of 470–500 s and thrust efficiency of 10–12% with energy of 4.5–14.6 J. Ten thousand shots achieved a total impulse of approximately 3.6 Ns, and the PTFE bars were consumed approximately 2 mm in length. However, uneven receding of the PTFE surface was observed. In order to investigate physical phenomena in a whole system, an unsteady numerical simulation of initial discharge, generation of plasma, heat transfer to the PTFE, heat conduction inside the PTFE, ablation from the PTFE surface and acceleration of plasma was performed. The calculated results were used to explain the physical phenomena in the cavity, especially ablated mass of the PTFE.     

Experimental investigation of thrust-vector deviation in a plasma thruster

We have studied deviation of the thrust vector of a plasma thruster with the aid of a magnetic system arranged behind the thruster edge. The magnetic field generated by the system acted upon the directional motion of ions in the jet. The experiments were performed on a laboratory model of the α-100 stationary plasma thruster (SPT) with an output channel diameter of 100 mm and overall dimensions L = 70 mm and D = 200 mm. The results of measurements showed that, in the range of parameters studied, the angle of rotation of the output plasma jet exhibited a nearly linear dependence on the current in coils of the deviating magnetic system.     

基于超导差分加速度测量原理的微推力测量方案

随着微推进技术的快速发展,微推力器推力的测试与标定也日益困难。为实现微推力的精密测量和动态特性测试,通过分析超导重力梯度仪的差分加速度测量原理、仪器的构成和原理电路等,提出了一种基于高温超导差分加速度测量原理和SQUID检测技术微推力测量方案,其推力测量范围为10^-7-10^-2N、测量频带从直流至4Hz。根据检验质量的动力学方程和超导回路磁通量守恒方程,导出了推力与输出差分电流的传递函数,分析了影响测量的主要噪声。方案可应用于微推力的测量,并有进一步提高推力测量精度的潜力。     

多档位扭摆推力测量装置的设计与标定实验

针对卫星用中高功率电推进器产生的推力范围跨度大，现有的推力测量装置存在测量范围不全、抗干扰能力差导致测不准等问题，开展了多档位三丝扭摆推力测量装置研究。首先，建立了推力测量装置物理模型，研究了推力与偏转位移之间的数学关系，并实现了多档位三丝扭摆推力测量装置的设计。接着，采用标准砝码与定制砝码对测量装置进行标定，并通过标定实验确定各档位下的测量误差。然后，综合考虑了装置的不确定度影响因素，设计相关试验开展不确定评估。实验结果表明：在实验环境不变的情况下，多档位三丝扭摆推力测量装置设计的小档位推力为98 mN的不确定度为0.030 mN（k=2）；中档位推力为490 mN的不确定度为0.068 mN（k=2），大档位推力为980 mN的不确定度为0.092 mN（k=2）。多档位扭摆推力测量装置采用换档位的测量方法实现了9.8~1029 mN范围的推力测量，测量精度高、抗干扰能力强，解决了宽范围推力测量过程中全量程精度难以保证的问题，为中高功率电推进器推力测量提供技术支撑。     

Effect of solute mixing in the liquid propellant of a pulsed plasma thruster

Sodium chloride or ammonia was dissolved in the water propellant of pulsed plasma thrusters to improve the performance. Pulsed plasma thrusters using liquid propellant utilize water as attractive alternative instead of Teflon. Water propellant enables in controlling propellant mass flow and leads to high specific impulse. However, liquid propellant pulsed plasma thrusters have larger plasma resistance and lower thrust power ratio than the common Teflon propellant thruster. Here, sodium chloride and ammonia solution of water were examined to decrease that plasma resistance. As a result, emission lines attributed from the solute were observed using sodium chloride aqueous solution propellant, and a 5% reduction of the plasma resistance was shown, and the thrust to power was increased. However, ammonia aqueous solution decreased the thruster performance.     

Performance characteristics of very low power cylindrical Hall thrusters for the nano-satellite “PROITERES-3”

Development of Hall thrusters for nano, small and low power satellites below 100 W is expected. In lowering Hall thruster power, the cylindrical type Hall thruster is more advantage than conventional coaxial-type Hall thrusters. In this study, a very low power cylindrical Hall thruster for nano-satellite “PROITERES-3” under development in Osaka Institute of Technology was designed, and the thruster performance was measured. As a result, a stable operation was achieved even with 10 W. The anode specific impulse and the anode thrust efficiency were 1570 s and 18.1%, respectively, with 66 W. Also, the discharge current oscillation was lower compared with that of SPT-type Hall thruster.     

LIPS-200离子推力器束流模型及其应用

建立了基于测量数据基础的LIPS-200离子推力器束流模型,给出了模型中积分发散问题的解决方法。应用模型计算了LIPS-200推力器中远场束流分布,给出了束流发散角、推力修正因子、南北位保推力器安装角等应用实例。     

Total impulse improvement of coaxial pulsed plasma thruster for small satellite

Research and development on pulsed plasma thruster (PPT) for small satellites has been conducted in Tokyo Metropolitan University (TMU). A PPT with coaxial electrode and PTFE cavity has been considered to gain larger impulse bit than that of parallel plate PPT. In this paper, the objectives are development of the PPTs with large impulse bit and large total impulse, to apply to a propulsion system for a small satellite. The electrodes and the propellant configuration were considered. As a result, divergent nozzle and larger cathode diameter improved thrust performances and larger impulse bit from 300 to 800 μNs was achieved. And the current reversal was appeared remarkable as its operation accumulated. Then, propellant-feeding mechanism was considered to improve its total impulse. With exchange of five propellants in each 20,000 shots, total impulse of 40 Ns was obtained. These results indicate that the PPT developed in TMU would provide various missions for a small satellite.     

Investigation of thrust characteristics of a dummy hollow-anode hall thruster

To measure the thrust of a dummy hollow-anode Hall Effect thruster, an IT-20 thrust meter was used whose mode of functioning is based on automatic compensation for the engine’s thrust by the force generated by a magnetoelectric compensator in the negative feedback loop. The dependences of the thrust on the discharge voltage and the magnitude of the magnetic field were obtained. The existence of oscillations in the discharge current was shown. The frequency of these oscillations matches the transit-time oscillation frequency of neutral particles, which may have an effect on the divergence between the calculated thrust values with those measured by the thrust meter.     

Fracture behavior of silicon nitride ceramics subjected to hypervelocity impact

A new advanced ceramic thruster made of monolithic silicon nitride ceramics was developed for the planetary exploration spacecraft AKATSUKI (PLANET-C) at Institute of Space and Astronautical Science (ISAS), Japan Aerospace Exploration Agency (JAXA). To ensure its operation onboard the spacecraft, the reliability of the ceramic thruster against micrometeoroid hypervelocity impact has been investigated. Silicon nitride plates were impacted by spheres of stainless-steel and other materials with 0.2-0.8-mm diameters in the velocity range up to 8.0 km/s using a two-stage light-gas gun. Using crater depth data under various impact conditions, the penetration equation of silicon nitride was determined. The impacted samples showed fracture patterns of three types: cratering, cratering with spallation, and perforation. These fracture patterns were well categorized by the multiple forms of the penetration equation.     

The criterion of optimal configuration of magnetic field in a thruster with closed electron drift

An integral criterion is suggested, which characterizes the configuration of magnetic field in a thruster with closed electron drift from the standpoint of efficiency of confinement of electrons in the ionization region. For the purpose of investigating the validity of the criterion, a thruster is developed with a new design of magnet system, which enables one to vary the structure of magnetic field in a wide range. The experimental investigation of the effect of magnetic field structure on the efficiency and specific impulse of plasma thruster revealed correlation between the results of application of the criterion and the parameters of efficiency of the thruster.     

双单摆式微推力测量系统的研究

为了实现空间微推力器的高精度推力测量,提出了一种电容桥式差分读出的双单摆推力测量方案,其测量频带从10-3Hz～0.016 Hz,可实现分辨率为0.1μN,量程1 mN的微推力测量。与意大利帕多瓦大学设计的双单摆微推力差分测量系统相比,方案具有更强的共模抑制能力,操作简单易行。     

Two-mode operation of plasma thruster

Much effort is presently devoted to developing stationary plasma thrusters (SPTs) capable of operating at elevated discharge voltages with large specific impulses. Classical SPTs in nominal regimes are characterized by a specific impulse of 1500 s at a thrust of 85 mN. We present experimental data on the characteristics of an SPT of the α-100 type capable of operating in two-mode regimes at the same supplied electric power ranging within 1.9–2.5 kW. The first (low-voltage) mode provides a thrust of about 120 mN, while the second (high-voltage) mode ensures a specific impulse of up to ∼3000 s.     

Determining the efficiency of a plasma thruster with closed electron drift

We propose a new method for the analysis of the efficiency of ion generation and acceleration, which can be used to determine the ion current in a plasma accelerator (thruster) with closed electron drift. The method is based on the measurement of the energy and angular distributions of ions in the plasma jet, the discharge power, and the jet thrust. The laws of variation of the plasma thruster efficiency as a function of the discharge voltage in the interval of 300–1000 V have been studied.     

Micro-multi-plasmajet array thruster for space propulsion applications

Thrust performance tests were conducted for an arrayed multi-plasmajet structure consisting of rectangular nozzle elements each with an exit height of 0.5 mm and length of 0.5 mm. To evaluate the thrust characteristics of the arrayed micro-multi-plasmajet, the thrust was measured using a calibrated cantilever-type thrust stand in vacuum. Using 3 × 3 nozzle elements, the micro-plasmajet array showed stable DC operation. Compared to conditions with cold-gas flow, the application of a DC discharge showed a significant improvement in thrust performance of at least 20% for thrust and specific impulse. Typical values achieved for thrust, specific impulse and thrust efficiency at an input power of 6.3 W were 8.5 mN, 77 s and 0.21, respectively.