Experimental and numerical simulation study of the effect for the anode positions on the discharge characteristics of 300 W 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^-1 and 0...     

The effect of anode axial position on the performance of a miniaturized cylindrical Hall thruster with a cusp-type magnetic field

A 200 W cylindrical Hall thruster with a cusp-type magnetic field was proposed, manifesting convergent plume and high specific impulse. In this paper, a series of ring-shaped anodes are designed and the influence of anode axial position on the performance of CHT with a cusp-type magnetic field is studied. The experimental results indicate that the thruster keeps stable operation at the condition of 140–270 W discharge power. When the anode moves axially towards the upstream cusp field, the thrust enhances from 6.5 mN to 7.6 mN and specific impulse enhances from 1658 s to 1939 s significantly. These improvements of thruster performance should be attributed to the enhancement of current utilization, propellant utilization and acceleration efficiency. According to the analyses on the discharge characteristics, it is revealed that as the anode moves upstream, the electron transport path could be extended, the magnetic field in this extended path could impede electron cross-field transport and facilitate the ionization intensity, yielding to the enhancement of current utilization and propellant utilization efficiency. Moreover, along with this enhancement of upstream ionization at the given anode flow rate, the main ionization region is thought to move upstream and then separate more apparently from the acceleration region, which has been demonstrated by the narrowing of ion energy distribution function shape. This change in acceleration region could decrease the ion energy loss and enhance acceleration efficiency. This work is beneficial for optimizing the electrode structure of thruster and recognize the ionization and acceleration process under the cusp magnetic field.     

Study on the influence of the discharge voltage on the ignition process of Hall thrusters

A high-speed charge-coupled device camera was used to capture images of the plume and acceleration channel of a Hall effect thruster during ignition at different discharge voltages. To better understand the influence of changes in the discharge voltage on the plasma parameters during thruster ignition, a particle-in-cell numerical model was used to calculate the distribution characteristics of the ion density and electric potential at different ignition moments under different discharge voltages. The results show that when the discharge voltage is high, the ion densities in the plume and acceleration channel are significantly higher at the initial phase of thruster ignition; with the gradual strengthening of the ignition process, the propellant avalanche ionization during thruster ignition occurs earlier and the pulse current peak increases. The main reason for these phenomena is that the change in the discharge voltage results in different energy acquisitions of the emitted electrons entering the thruster channel.     

Investigation of short-channel design on performance optimization effect of Hall thruster with large height–radius ratio

In this study, the neutral gas distribution and steady-state discharge under different discharge channel lengths were studied via numerical simulations. The results show that the channel with a length of 22 mm has the advantage of comprehensive discharge performance. At this time, the magnetic field intensity at the anode surface is 10% of the peak magnetic field intensity. Further analysis shows that the high-gas-density zone moves outward due to the shortening of the channel length, which optimizes the matching between the gas flow field and the magnetic field, and thus increases the ionization rate. The outward movement of the main ionization zone also reduces the ion loss on the wall surface. Thus, the propellant utilization efficiency can reach a maximum of 96.8%. Moreover, the plasma potential in the main ionization zone will decrease with the shortening of the channel. The excessively short-channel will greatly reduce the voltage utilization efficiency. The thrust is reduced to a minimum of 46.1 mN. Meanwhile, because the anode surface is excessively close to the main ionization zone, the discharge reliability is also difficult to guarantee. It was proved that the performance of Hall thrusters can be optimized by shortening the discharge channel appropriately, and the specific design scheme of short-channel of HEP-1350PM was defined, which serves as a reference for the optimization design of Hall thruster with large height–radius ratio. The short-channel design also helps to reduce the thruster axial dimension, further consolidating the advantages of lightweight and large thrust-to-weight ratio of the Hall thruster with large height–radius ratio.     

Simulation of the effect of a magnetically insulated anode on a low-power cylindrical Hall thruster

The intersection point of the characteristic magnetic field line (CMFL) crossing the anode boundary with the discharge channel wall,and its influence on thruster performance and the energy and flux of ions bombarding the channel wall,have been studied numerically.The simulation results demonstrate that with the increase in distance from the crossover point of the CMFL with the channel wall to the bottom of the thruster channel,the ionization rate in the discharge channel gradually increases;meanwhile,the ion energy and ion current density bombarding the channel wall decreases.When the point of the CMFL with the channel wall is at the channel outlet,the thrust,specific impulse,and efficiency are at a maximum,while the ion energy and ion current density bombarding the channel wall are at a minimum.Therefore,to improve the performance and lifetime of the thruster,it is important to control the point of intersection of the CMFL with the channel wall.     

Particle-in-cell simulation for effect of anode temperature on discharge characteristics of a Hall effect thruster

Propellant gas flow has an important impact on the ionization and acceleration process of Hall effect thrusters (HETs). In this paper, a particle-in-cell numerical method is used to study the effect of the anode temperature, i.e., the flow speed of the propellant gas, on the discharge characteristics of a HET. The simulation results show that, no matter the magnitude of the discharge voltage, the calculated variation trends of performance parameters with the anode temperature are in good agreement with the experimental ones presented in the literature. Further mechanism analysis indicates that the magnitude of the electron temperature is responsible for the two opposing variation laws found under different discharge voltages. When the discharge voltage is low, the electron temperature is low, and so is the intensity of the propellant ionization; the variation of the thruster performance with the anode temperature is thereby determined by the variation of the neutral density that affects the propellant utilization efficiency. When the discharge voltage is high, the electron temperature is large enough to guarantee a high degree of the propellant utilization no matter the magnitude of the anode temperature. The change of the thruster performance with the anode temperature is thus dominated by the change of the electron temperature and consequently the electron-neutral collisions as well as the electron cross-field mobility that affect the current utilization efficiency.     

Study on the influences of ionization region material arrangement on Hall thruster channel discharge characteristics

There exists strong interaction between the plasma and channel wall in the Hall thruster,which greatly affects the discharge performance of the thruster.In this paper,a two-dimensional physical model is established based on the actual size of an Aton P70 Hall thruster discharge channel.The particle-in-cell simulation method is applied to study the influences of segmented low emissive graphite electrode biased with anode voltage on the discharge characteristics of the Hall thruster channel.The influences of segmented electrode placed at the ionization region on electric potential,ion number density,electron temperature,ionization rate,discharge current and specific impulse are discussed.The results show that,when segmented electrode is placed at the ionization region,the axial length of the acceleration region is shortened,the equipotential lines tend to be vertical with wall at the acceleration region,thus radial velocity of ions is reduced along with the wall corrosion.The axial position of the maximal electron temperature moves towards the exit with the expansion of ionization region.Furthermore,the electron-wall collision frequency and ionization rate also increase,the discharge current decreases and the specific impulse of the Hall thruster is slightly enhanced.     

Numerical study of low-frequency discharge oscillations in a 5kW Hall thruster

A two-dimensional particle-in-cell plasma model is built in the R–Z plane to investigate the lowfrequency plasma oscillations in the discharge channel of a 5 kW LHT-140 Hall thruster. In addition to the elastic, excitation, and ionization collisions between neutral atoms and electrons,the Coulomb collisions between electrons and electrons and between electrons and ions are analyzed. The sheath characteristic distortion is also corrected. Simulation results indicate the capability of the built model to reproduce the low-frequency oscillation with high accuracy. The oscillations of the discharge current and ion density produced by the model are consistent with the existing conclusions. The model predicts a frequency that is consistent with that calculated by the zero-dimensional theoretical model.     

Investigation on plasma characteristics in a laser ablation pulsed plasma thruster by optical emission spectroscopy

In order to further improve the propulsion performance of pulsed plasma thrusters for space micro propulsion, a novel laser ablation pulsed plasma thruster is proposed, which separated the laser ablation and electromagnetic acceleration. Optical emission spectroscopy is utilized to investigate the plasma characteristics in the thruster. The spectral lines at different times,positions and discharge intensities are experimentally recorded, and the plasma characteristics in the discharge channel are concluded through analyzing the variation of spectral lines. With the discharge energy of 24 J, laser energy of 0.6 J and the use of aluminum propellant, the specific impulse and thrust efficiency reach 6808 s and 70.6%, respectively.     

Investigation of the self-induced magnetic field characteristics in a pulsed plasma thruster with flared electrodes

The self-induced magnetic field in a pulsed plasma thruster(PPT)with flared electrodes is investigated for a better understanding of the working process and the structural design of the thruster.A two-dimensional model of the magnetic field is built and is validated by comparing the simulated results with the experimental results in literature.The magnetic flux density in the discharge channel during the working process is presented and analyzed regarding the electrode structures.The calculated magnetic field flux density decreases from 0.8 T at the upstream to 0.1 T and below at the downstream in the discharge channel(68 J).The peak of the magnetic flux density over time lags behind the current peak,which provides evidence for the existence of a moving plasma sheet in the discharge process.The magnetic field induced by the current in the extra bending part of the anode enhances the Lorentz force,which acts on the charged particles near the propellant.Finally,the geometric study indicates that the electromagnetic impulse bit does not monotonically increase with the flared angle of the electrodes.Instead,it reaches a maximum at a certain flared angle,which could provide significant suggestions for structural optimization.     

Effect of vacuum backpressure on discharge characteristics of hollow cathode

A hollow cathode is the electronic source and neutralizer of the Hall thruster and an ion thruster.When the orbit of an all-electric propulsion satellite changes from 100 km to 36 000 km, the backpressure changes by two to three orders of magnitude. In this paper, the influence of the backpressure on the discharge characteristics of the hollow cathode has been studied experimentally in the so-called diode configuration. With the increase in the backpressure, the anode voltage decreases gradually, and the amplitude of the current oscillation decreases significantly. Additionally, the plasma is relatively stable, the most probable ion energy and the width of the ion energy distribution reduces, and the electron distribution function inclines toward the Maxwell distribution under high backpressure. The analysis results show that the backpressure affects the gas ionization and the ionic acoustic turbulence, which also affects the discharge characteristics of the hollow cathode.     

Effect of the Discharge Voltage on the Performance of the Hall Thruster

In this paper,a two-dimensional physical model is established according to the discharging process in the Hall thruster discharge channel using the particle-in-cell method.The influences of discharge voltage on the distributions of potential,ion radial flow,and discharge current are investigated in a fixed magnetic field configuration.It is found that,with the increase of discharge voltage,especially during 250-650 V,the ion radial flow and the collision frequency between ions and the wall are decreased,but the discharge current is increased.The electron temperature saturation is observed between 400-450 V and the maximal value decreases during this region.When the discharge voltage reaches 700 V,the potential distribution in the axis direction expands to the anode significantly,the ionization region becomes close to the anode,and the acceleration region grows longer.Besides,ion radial flow and the collision frequency between ions and the wall are also increased when the discharge voltage exceeds 650 V.     

Discharge instability in a plasma contactor

Like the hollow cathode, discharge instability also occurs during the operation of a plasma contactor.Voltage and current probes were employed to test the change of keeper voltage, keeper current,anode voltage, and anode current parameters with time under different working conditions. The anode current range corresponding to the discharge instability phenomenon is about 0.4 A to 1.2 A,and the emission characteristic curve in this area appears to bulge wherein the four parameters all produce different degrees of oscillation, the anode current oscillation being the greatest. Its waveform is considered to consist of a small-amplitude, high-frequency triangular wave and a large-amplitude,low-frequency sawtooth wave, and we have explained the shape of the wave. Each parameter shows hundreds of Hz in oscillation frequency and the phases of the four parameters appear to be regular.After fast Fourier transform processing, the frequency and amplitude of the main peak of the anode current oscillation tend to change with changes of the anode current, and there are differences in the trends under different keeper currents and xenon flows.     

Plume diagnostics of BUSTLab microwave electrothermal thruster using Langmuir and Faraday probes

This study presents the Langmuir and Faraday probe measurements conducted to determine the plume characteristics of the BUSTLab microwave electrothermal thruster (MET). The thruster, designed to operate at 2.45 GHz frequency, is run with helium, argon and nitrogen gases as the propellant. For the measurements, the propellant volume flow rate and the delivered microwave power levels are varied. Experiments with nitrogen gas revealed certain operation regimes where a very luminous plume is observed. With the use of in-house-built Langmuir probes and a Faraday probe with guard ring, thruster plume electron temperature, plasma density and ion current density values are measured, and the results are presented. The measurements show that MET thruster plume effects on spacecraft will likely be similar to those of the arcjet plume. It is observed that the measured plume ion flux levels are very low for the high volume flow rates used for the operation of this thruster.     

Effect of Segmented Electrode Length on the Performances of an Aton-Type Hall Thruster

The influences of the low-emissive graphite segmented electrode placed near the channel exit on the discharge characteristics of a Hall thruster are studied using the particlein-cell method.A two-dimensional physical model is established according to the Hall thruster discharge channel configuration.The effects of electrode length on the potential,ion density,electron temperature,ionization rate and discharge current are investigated.It is found that,with the increasing of the segmented electrode length,the equipotential lines bend towards the channel exit,and approximately parallel to the wall at the channel surface,the radial velocity and radial flow of ions are increased,and the electron temperature is also enhanced.Due to the conductive characteristic of electrodes,the radial electric field and the axial electron conductivity near the wall are enhanced,and the probability of the electron-atom ionization is reduced,which leads to the degradation of the ionization rate in the discharge channel.However,the interaction between electrons and the wall enhances the near wall conductivity,therefore the discharge current grows along with the segmented electrode length,and the performance of the thruster is also affected.     

Life test research of a high specific impulse Hall thruster HEP-140MF

In this study, a high specific impulse Hall thruster, HEP-140MF, having a high discharge voltage, was used to accelerate ions. We aimed to obtain a high specific impulse and an acceleration zone moving downstream toward the channel exit to reduce wall sputtering erosion of the walls of the discharge channel, hence ensuring an enhanced lifetime. To study the lifetime characteristics of the high specific impulse Hall thruster, a life test was performed on the HEP- 140MF thruster for the first time, and performance parameters, such as thrust, specific impulse, and efficiency, were measured. Changes in the performance parameters and evolutions in the surface profiles of the discharge channel wall were summarized. The reasons contributing to these changes during the life test were analyzed. Moreover, the accelerated life test method was validated on the HEP-140MF.     

The far-field plasma characterization in a 600W Hall thruster plume by laser-induced fluorescence

Non-intrusive characterization of the singly ionized xenon velocity in Hall thruster plume using laser induced fluorescence(LIF) is critical for constructing a complete picture of plume plasma,deeply understanding the ion dynamics in the plume, and providing validation data for numerical simulation. This work presents LIF measurements of singly ionized xenon axial velocity on a grid ranging from 100 to 300 mm in axial direction and from 0 to 50 mm in radial direction for a600 W Hall thruster operating at the nominal condition of discharge voltage 300 V and discharge current 2 A, the influence of discharge voltage is investigated as well. The ion velocity distribution function(IVDF) results in the far-field plume demonstrate a profile of bimodal IVDFs, especially prominent at radial distances greater than channel inner radius of 22 mm at axial position of 100 mm, which is quite different from that of the near-field plume where bimodal IVDFs occur in the central core region for the same power Hall thruster when compared to previous LIF measurements of BHT-600 by Hargus(2010 J. Propulsion Power 26 135).Beyond 100 mm, only single-peak IVDFs are measured. The two-dimensional ion velocity vector field indicates the bimodal axial IVDF is merely a geometry effect for the annular discharge channel in the far-field plume. Results about the IVDF, the most probable velocity and the accelerating potential profile along the centerline all indicate that ions are still accelerating at axial distances greater than 100 mm, and the maximum most probable velocity measured at300 mm downstream of the exit plane is about 19 km s-1. In addition, the most probable velocity of ions along radial direction changes a little except the lower velocity ion populations in the bimodal IVDF cases. The ion temperature at axial distances of 10 and 300 mm oscillates along the radial direction, while the ion temperature first increases, and then decreases for the 200 mm case. Finally, the axial position for the ion peak axial velocity on the thruster centerline is shifted upstream for higher discharge voltages, and the velocity curve is becoming steeper with the discharge voltage before reaching the maximum. This observation can be used as a criterion to optimize the thruster performance.     

Performance of a 4 cm iodine-fueled radio frequency ion thruster

The performance of an iodine radio ion thruster with a 4 cm diameter(IRIT4) was studied experimentally in this paper. Regulation of the mass flow rates of the iodine propellant is achieved by using a temperature control method of the iodine reservoir. Performance of the thruster using iodine as propellants is obtained at different total thruster powers of 40.6–128.8 W,different grid voltages of 1000–1800 V and the iodine flow rate of 100 μgs~(-1). Results show that thrust and specific impulse increase approximately linearly with the increasing total thruster power and the screen grid voltage. The thrust of 2.32 mN and the specific impulse of 2361 s are obtained at the nominal total thruster power of 95.8 W and the screen grid voltage of 1800 V. It is also indicated that performance of the iodine propellant is comparable to that of the xenon propellant; and a difference between them is that the iodine thrust is slightly higher than xenon when the total thruster power is more than 62 W. At the nominal 95.8 W total thruster power, the thrust values of them are 2.32 m N and 2.15 mN respectively, and the thrust-to-power ratios of them are 24.2 mN kW~(-1) and 23.5 mN kW~(-1), respectively.     

The ablation characteristics of laser-assisted pulsed plasma thruster with metal propellant

In this study,a laser-assisted pulsed plasma thruster (LA-PPT) with a novel configuration is proposed as an electric propulsion thruster which separates laser ablation and electromagnetic acceleration.Owing to the unique structure of the thruster,metals can also be used as propellants,and a higher specific impulse is expected.The ablation quality,morphology,and plume distribution of various metals (aluminium alloy,red copper,and carbon steel) with different laser energies were studied experimentally.The ablation morphology and plume distribution of red copper were more uniform,as compared to those of other metals,and the ablation quality was higher,indicating its greater suitability for LA-PPT.The plume generated by nanosecond laser ablation of aluminium alloy expanded faster,which indicated that the response time of the thruster with aluminium alloy as the propellant was shorter.In addition,when the background pressure was 0.005 Pa,an obvious plume splitting phenomenon was observed in the ablation plume of the pulsed laser irradiating aluminium alloy,which may significantly reduce the utilisation rate of the propellant.     

Impact of exterior electron emission on the self-sustaining margin of hollow cathode discharge

Hollow cathode researches used to focus on the inner cavity or downstream plume, however,rarely on the gap between the throttling orifice plate and the keeper plate(T-K gap), which was found to impact the self-sustaining margin of hollow cathode discharge in this paper. Near the lower margin, the main power deposition and electron emission and ionization regions would migrate from inner cavity and downstream plume to the T-K gap, in which case, the source and destination of each m A current therein matter for the self-sustaining capability. Changing the metal surfaces in the T-K gap with emissive materials proved effective in lowering the lower margin by supplementing auxiliary thermionic emission, compensating electron loss on cold absorbing walls and suppressing discharge oscillations. By doing so, the lower margin of a 4 A hollow cathode was lowered from 1 to 0.1-0.2 A, enabling it to couple with low power Hall thruster without extra keeper current.