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.

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.

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.

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.

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.

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.

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.

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.

Low-frequency oscillations in Hall thrusters

In this paper, we summarize the research development of low-frequency oscillations in the last few decades. The findings of physical mechanism, characteristics and stabilizing methods of low-frequency oscillations are discussed. It shows that it is unreasonable and incomplete to model an ionization region separately to analyze the physical mechanism of low-frequency oscillations. Electro-dynamics as well as the formation conditions of ionization distribution play an important role in characteristics and stabilizing of low-frequency oscillations. Understanding the physical mechanism and characteristics of low- frequency oscillations thoroughly and developing a feasible method stabilizing this instability are still important research subjects.

Effect of low-frequency oscillation on performance of Hall thrusters

In this paper, a direct connection between the discharge current amplitude and the thruster performance is established by varying solely the capacitance of the filter unit of the Hall thrusters. To be precise, the variation characteristics of ion current, propellant utilization efficiency, and divergence angle of plume at different low-frequency oscillation amplitudes are measured. The findings demonstrate that in the case of the propellant in the discharge channel just meets or falls below the full ionization condition, the increase of low-frequency oscillation amplitude can significantly enhance the ionization degree of the neutral gas in the channel and increase the thrust and anode efficiency of thruster. On the contrary, the increase in the amplitude of low-frequency oscillation will lead to increase the loss of plume divergence, therefore the thrust and anode efficiency of thruster decrease.

Effect of Magnetic Mirror on the Asymmetry of the Radial Profile of Near-Wall Conductivity in Hall Thrusters

Considering the actual magnetic field configuration in a Hall thruster, the effect of magnetic mirror on the radial profile of near-wall conductivity （NWC） is studied in this paper. The plasma electron dynamic process is described by the test particle method. The Monte Carlo scheme is used to solve this model. The radial profile of electron mobility is obtained and the role of magnetic mirror in NWC is analysed both theoretically and numerically. The numerical results show that the electron mobility peak due to NWC is inversely proportional to the magnetic mirror ratio and the asymmetry of electron mobility along the radial direction gets greater when the magnetic mirror is considered. This effect indicates that apart from the disparity in the magnetic field strength, the difference in the magnetic mirror ratio near the inner and outer walls would actually augment the asymmetry of the radial profile of NWC in Hall thrusters.

Design and fabrication of a full elastic sub-micron-Newton scale thrust measurement system for plasma micro thrusters

In this work,a force measurement system is proposed to measure the thrust of plasma microthruster with thrust magnitude ranging from sub-micro-Newtons to hundreds micro-Newtons.The thrust measurement system uses an elastic torsional pendulum structure with a capacitance sensor to measure the displacement,which can reflect the position change caused by the applied force perpendicular to the pendulum axis.In the open-loop mode,the steady-state thrust or the impulse of the plasma micro-thruster can be obtained from the swing of the pendulum,and in the closed-loop mode the steady-state thrust can be obtained from the feedback force that keeps the pendulum at a specific position.The thrust respond of the system was calibrated using an electrostatic weak force generation device.Experimental results show that the system can measure a thrust range from 0 to 200μN in both open-loop mode and closed-loop mode with a thrust resolution of 0.1μN,and the system can response to a pulse bit at the magnitude of 0.1 m N s generated by a micro cathode arc thruster.The background noise of the closed-loop mode is lower than that of the open-loop mode,both less than 0.1 m N/Hz in the range of 10 mHz to 5 Hz.

Characterization and thrust measurements from electrolytic decomposition of ammonium dinitramide(ADN)based liquid monopropellant FLP-103 in MEMS thrusters

Although ammonium dinitramide(ADN)has been targeted as a potential green monopropellant in future space vehicles,its application potential in Micro-electrical–Mechanical System(MEMS)thrusters or microthrusters has been seldom reported in open literature.In this paper,electrolytic decomposition of Ammonium Dinitramide(ADN)-based liquid monopropellant FLP-103 was carried out in an open chamber and MEMS thrusters were fabricated from poly-dimethylsiloxane(PDMS)to characterize the power consumption.Two thrust measurement methods were employed to investigate the electrolytic decomposition of FLP-103 in MEMS microthrusters.The results show that the monopropellant can be successfully ignited at room temperature through 80 V,0.1 A(8 W)using copper wire as electrodes.In the current thruster design,low thrust was obtained at FLP-103 flowrate of 40μl·min^(-1)but it generated the highest specific impulse,Isp,among all the flowrates tested.The experiments successfully demonstrated the potential application of electrolytic decomposition of FLP-103 in MEMS thrusters.

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.

Realization of proportional navigation and lateral acceleration autopilot for endoatmospheric kinetic interceptor with pulse thrusters

The guidance and control strategy for endoatmospheric kinetic interceptor controlled by lateral pulse thrusters was detailed.The pulse thruster control system was firstly proposed.The sample-and-hold approach was introduced to design a novel lateral acceleration autopilot on the basis of traditional two-loop topology.Combined with proportional navigation guidance law and the novel autopilot,the overall ballistic trajectory was presented and examined.Simulation results show that the pulse thruster control strategy can greatly improve the control system response speed and the maximal acceleration capability for realizing kinetic kill interception.

Correlation Method of Average Plasma Velocity Measurement in Hall-Effect Thrusters

A new method to measure the average plasma velocity in a Hall-effect thruster is presented. The method is brought forward in virtue of the characteristics of low frequency oscillation induced by the propellant ionization in the channel and the oriented movement feature of the plasma density out of the channel. The method, equivalent to the correlation method generally used in the signal processing field, provides a solution to the problem of specific impulse measurement on a timescale of hundreds of microseconds and makes the time evolution of average plasma velocity clear. The comparison between the measured value and the calibrated value shows that the relative error is about 3%.