Low-power Hall Thruster Research Articles

Development of a low-power Hall thruster with permanent magnets and a dual trigger electrode hollow cathode for the Qilu satellite constellation

A low-power Hall effect thruster (LpHet-100) has been installed on the Qilu-3 satellite and has been launched by the Long March 2D rocket in January 2023. The specific features of LpHet-100 are presented in this article. An innovative dual trigger electrode poison-resistant hollow cathode and an effective magnetic shielding configuration with permanent magnets, additive manufacturing gas distributor are designed and experimentally verified. The prominent characteristics of a shorter starting time, lower ignition shock, and self-healing have been shown in tests of the dual trigger electrode hollow cathode. Thrust performance tests proved that LpHet-100 can work stably in the power range of 50–230 W and generate a thrust of 2.9–10.8 mN, with the maximum specific impulse of 1594 s and maximum thrust efficiency of 38 %. Rigorous environmental tests were conducted to verify the thruster design scheme. In addition, Torque measurement of Hall thrusters is incorporated into the ground tests for the first time. A plasma torque of 5.65–9.88 μN·m is obtained when LpHet-100 operated in 132–228 W. It provides a critical guideline for the attitude and orbit control schemes of the Qilu-3 satellite, which is an essential part of space mission planning. Generally, this study proposes new design of the thruster body and hollow cathode, providing an exhaustive reference for the development process and application of low-power Hall thrusters.

On-orbit mission overview of the low power hall thruster propulsion system aboard venμs satellite

Venμs is a satellite launched in 2017, for super-spectral Earth imaging and Electric Propulsion System (EPS) demonstration. In this paper we overview EPS design and operation throughout all five mission phases, from open/closed-loop orbit control, through orbit descent (720 → 410 km), orbit maintenance under high drag environment, to orbit raising (410 → 560 km). The EPS consisted of two throttleable Hall thrusters, PPU, Propellant Management Assembly (PMA), and a 9 L propellant tank carrying 16 kg of Xenon. Both thrusters operated in the 300–550 W power range, generated a combined total impulse of 158.1 kN-sec and consumed all propellant. Two methods are described to compute the remaining propellant mass – “Bookkeeping” and “PTV” methods. The advantages and disadvantages of each method are discussed in light of the Venµs mission. Thruster performance was measured on-orbit using the “Orbit Determination” method and compared to laboratory experiments conducted on the ground with identical thrusters. The measured performance on-orbit was found to agree within the error bars with the performance measured on the ground. Lastly, we present several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. We present the methods used to construct a solution and implement it.

Performance evaluation of a low power Hall thruster with carbon dioxide propellant

Carbon dioxide (CO2) is an attractive candidate for condensable propellants because its availability and storability. Although several studies have been reported on adopting CO2 as a propellant for Hall thrusters, there have been few studies on CO2 standalone operation, and none of them have focused on physical phenomena in plasmas. This study experimentally evaluated, thrust performance and plume characteristics of a 100 W-class Hall thruster under CO2 operation. CO2 operation was compared to xenon operation under the same discharge power conditions to determine its unique properties. As typical thrust performance under this experimental condition, a specific impulse of 727 s, thrust-power-ratio of 19.8 mN/kW, and anode efficiency of 0.0711 were achieved at a discharge voltage of 200 V and a discharge power of 352 W under a background pressure of 3.4 × 10−2 Pa. Plume measurements showed that mass utilization efficiency was 0.228 under CO2 operation, which was particularly poor compared to under xenon operation. The first reason for this is that the ionization mean free path of CO2 is 5.18 mm, about eight times longer than that of xenon. The second reason is that the energy is used for other processes, such as dissociative ionization and dissociation etc., considering the reaction process of the particles. Therefore, the results suggest narrowing the channel width to increase the channel density, and increasing the magnetic field strength to suppress the discharge current as ways to improve thrust efficiency.

Ultra-low frequency discharge instability of a barium tungsten hollow cathode at lower self-sustaining margin

Development of low-orbiting constellations stimulated the demand for low-power Hall thrusters, as well as their accompanying low-current hollow cathodes. In this paper, such cathodes were tested on quite low currents to search their self-sustaining limit, and were found to exhibit an ultra-low frequency instability at lower current limit. The instability would grow fast when the cathode current was below 0.3A, and oscillate on 0.1Hz to a few Hz with a visibly blinking plume. At the same time, the emitter temperature was fluctuating with a 15 °C peak-to-peak amplitude. The instability was confirmed to originate from the cathode interior, as a result of depleted plasma heating. Then the thermionic cathode had to rely mainly on neutral gas to heat the emitter. Because thermionic emission was in exponential relation with emitter temperature, the temperature fluctuation was then amplified to cause larger fluctuation in total current as well as in total discharge power, which led to deeper fluctuation in temperature. Regarding the heating of both neutral gas and emitter, an instability model showed that multiple unnoticed factors could determine the onset of instability, including the thermal inertia of neutral gas and emitter, the heat transfer from discharging plasma to neutral gas, the heat loss of cathode and the response time of power supply. As a result of this instability, the plume ion energy was elevated from <50eV to 30∼120eV due to the fluctuating density on cathode exit and fluctuating discharge voltage. It was estimated that the increased ion energy could accelerate the keeper erosion by 4∼22 times.

Corrigendum: Coupling plasma physics and chemistry in the PIC model of electric propulsion: application to an air-breathing, low-power Hall thruster

Corrigendum: Coupling plasma physics and chemistry in the PIC model of electric propulsion: application to an air-breathing, low-power Hall thruster

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 low-power 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 are 7.19 and 7.58 mN, 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 mN/kW 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.

Study on a hybrid-produced magnetic-field Hall thruster

The increasing interest in all-electric, low-Earth orbit satellite constellations has stimulated the rapid development of electric propulsion technology, particularly low-power Hall thrusters. In this study, the difficulties in magnetic field design were analyzed in low-power Hall thrusters, and a new and efficient hybrid production technique is proposed. Hybrid production combines the advantages of coil and permanent magnet production; the main magnetic field is generated by permanent magnets, and the regulating magnetic field is generated by coils in parallel. Accordingly, a prototype of a 600-W hybrid-produced magnetic field Hall thruster, HPHT-600, was designed and tested. The HPHT-600 has a small size (channel outer diameter 54 mm), light weight (1.2 kg), low area–volume ratio (0.148), and low magnetic power. In addition, it exhibits a power envelope of 300–1000 W and a total efficiency of 53.5 % at the discharge voltage of 300 V, anode volume flow rate of 35 sccm, and total specific impulse of 1900 s at 500 V and 20 sccm.

Effect of magnetic field strength on the performance characterization of a low-power wall-less Hall thruster

The wall-less Hall thruster is expected to overcome the issue of wall erosion in miniaturized Hall thrusters. A unique structure of negative gradient magnetic field is used to constrain the plasma behavior. In this paper, permanent magnet rings with different thicknesses were used to adjust the magnetic field strength, and its effect on discharge characteristics and thrust performance was experimentally investigated. The experiment results show that increasing the magnetic field strength can reduce the current oscillation and increase the maximum operating voltage. Within a limited range, increasing the magnetic field strength can significantly improve thrust performance. However, excessive magnetic field strength leads to a decrease in anode efficiency. The optimal magnetic field strength appears at 121 mT, with a thrust of 5.5 mN, a specific impulse of 935 s, a thrust power ratio of 51 mN/kW, and an anode efficiency of 23.1 %. Based on the analysis of different efficiency parameters and plume images, under excessively high magnetic field strength, the shortening of the ionization region may lead to intensified leakage of neutral atoms, which may be the primary reason for the reduction in propellant utilization and anode efficiency.

Magnetic focusing of low-power Hall thruster with a center-mounted hollow cathode

Magnetic focusing with xenon propellant has been verified in orbit from medium-to high-power Hall thrusters but has yet to be achieved in low-power Hall thrusters. A magnetically focused Hall thruster was designed to solve the common issue of large plume divergence angle in low-power Hall thrusters, which realizes the plasma lens magnetic field topography with permanent magnet excitation on the wide channel and adopts the center-mounted cathode layout. Experimental results indicate that the HEP-300CHC thruster exhibits an optimal magnetic focusing mode when operated at 288W with plume divergence half-angle approximately 11.7°. The magnetic focusing range is relatively narrow, concentrated within discharge voltage between 210V and 230V, and anode flowrate between 18sccm and 20sccm.The anode efficiency is higher than 40% at the nominal power, and the maximum anode efficiency reach 52% at the discharge power of 300W. This study provides a reference basis for ion beam control of low-power Hall thrusters.

96 kN-sec endurance test of the R-200 low power hall thruster

The R-200 low power Hall thruster unit completed a 96 kNs endurance test, during which the anode mass flow rate was adjusted to maintain a constant power of 250 W throughout the campaign. The test was performed at Rafael, while the initial performance validation was conducted at ESTEC. At the end of the endurance test the thruster exhibited nearly unchanged performance and did not reach its end-of-life. During the endurance test, the thruster performance was characterized 5 times in the 100–250 W discharge power range, the ceramic channel profile was measured 6 times to assess the erosion rate, thrust was continuously tracked multiple times a day, an ignition test was conducted after 60 kNs to assess continuous operation impact on ignition capability, and sensitivity to magnetic field changes was checked twice (at 12 kNs and 65 kNs). Two notable events occurred during the test – a cathode keeper breach and a ceramic channel breach. The keeper electrode was replaced, and cathode position adjusted. following the ceramic channel breach the thruster continued to operate regularly, and with no performance degradation, even after the inner pole was exposed to the plasma discharge. The thruster generated a constant thrust of 13 mN throughout the entire test while the Isp decreased from 1,250 s at the beginning of life to 1,160 s at the end of test. Performance degradation of less than 10% was measured at all operating points in the 100–250 W discharge power range at the end of the experiment. The ignition test validated 300 successful thruster ignition cycles. Using channel profile geometry data, a simple semi-empirical erosion model was calibrated to predict future breaching events and allow for an improved channel design. The information gathered during this endurance test campaign was used to redesign the cathode keeper and inner ceramic channel profile. A subsequent test of the improved thruster unit design showed an expected lifetime of over 200 kNs.

Effects of the magnetic field intensity on pole erosion of low-power Hall thrusters

Magnetic pole erosion is a key life-limiting factor of low-power Hall thrusters. In this study, the effects of the magnetic field intensity near the inner magnetic pole on the pole erosion of a 100 W permanent-magnet Hall thruster at rated operating point were investigated using magnetic field intensity adjustment technology. The experimental results indicate that the erosion rate of the inner magnetic pole decreased considerably with a decrease in the magnetic field intensity near the inner magnetic pole, and the erosion rate can be decreased by over 37.5% when the magnetic field intensity decreased by 20%. Further analysis indicates that the decrease of magnetic field intensity reduces the number of electrons moving along the magnetic field lines to the magnetic pole, and increases the surface potential of the magnetic pole. The reduction of the potential difference between the channel outlet and the magnetic pole leads to a decrease in the current density and energy of the ions sputtered to the magnetic pole, which contributes greatly to the decrease in pole erosion rate. This study provides a reference for the protection of magnetic pole erosion.

Experimental and numerical studies of low power Hall thruster performance

The low power Hall thruster has the potential to be applied in multiple apace missions. In order to study the performance of the low power Hall thruster, a 200W-class Hall thruster is designed and manufactured. The plasma properties in the discharge channel are investigated via numerical simulation. The ionic composition and ion current density are respectively measured by a Wien filter and a Faraday probe. The generated thrust is measured by a thrust target. The simulation results show that the maximal ion number density is more than 1×1018 m-3. The ionization and acceleration regions are located in the midstream and downstream of the discharge channel. The experimental results reveal that thrust, anode specific impulse and anode efficiency are 12.13 mN, 1380 s, and 41% at 200 W. The ion current density decreases along the radial direction. According to the results of simulation and experiment, the thruster is reasonable and feasible. Our research provides a reference for the design of low power Hall thruster. Structural optimization and performance improvement will be carried out further.

Demonstration and experimental characteristics of a water-vapor Hall thruster

Water is an attractive candidate for condensable propellants owing to its availability, handleability, and sustainability. This study proposes the use of water vapor as a propellant for a low-power Hall thruster, and experimentally demonstrates the feasibility of this proposal. Based on the performance estimation from the plume diagnostics, a thrust-to-power ratio of 19 mN/kW, specific impulse of 550–860 s, and anode efficiency of 5–8 % were obtained at an anode power of 233–358 W. From further efficiency analysis, the mass utilization efficiency of water was found to be the most deteriorated among the internal efficiencies compared to the conventional xenon propellant, which was consistent with the expectations from a small discharge current oscillation, large beam divergence, and increase in low-energy ions. Moreover, additional power loss via reactions unique to polyatomic molecules was indicated by evaluation of the ionization cost. In this experiment, the mass utilization efficiency was improved with an increase in the anode voltage from 200 to 240 V without degradation of the power utilization. This suggests that operating at a higher voltage is more suitable for a water-vapor Hall thruster.

Numerical simulation optimization of neutral flow dynamics in low-power Hall thruster

This paper utilizes a low-power Hall thruster to investigate its flow field’s neutral flow dynamics. Specifically, under a free molecular flow with a high Knudsen number (Kn>10), we employ the flow conductance theory to investigate the interplay between the internal structural parameters of the thruster flow field and its conductance, which is directly reflected in the neutral gas distribution.Hence, first, we increase the number and diameter of the orifices in the buffer chamber so that the baffle effectively increases the flow field conductance, and thus the number density (nn) of the neutral gas in the discharge channel increases. This phenomenon reduces the ionization mean free path (λi) and thus increases the utilization rate (ηi) of the neutral gas. Then, we alter the baffle orifices distribution to an unequally spaced distribution and stagger the outlet orifices, resulting in a highly homogeneity neutral gas distribution in the circumferential direction, ultimately increasing the effective outlet area and reducing the backflow probability of the neutral gas molecules. This operation reduces the low-frequency oscillation amplitude and thus enhances the thruster’s operation stability. Finally, we explore the influence of the entire operating conditions, especially temperature and mass flow rate variations, on the universality of the simulation results. Overall, this work provides a new concept for optimizing the neutral flow field dynamics of a Hall thruster.

Plasma diagnosis inside the discharge channel of a low-power Hall thruster working on Xe/Kr mixtures

In this paper, optical diagnostics for the internal plasma of a low-power Hall thruster working on Xe/Kr mixtures is carried out. The emission spectroscopy results show that the atomic and ionic spectral lines reach the maximum at 5 mm and 3 mm upstream of the channel exit, respectively, indicating a spatial separation between atom excitation and ionization. The krypton atomic spectral intensity obtains the maximum closer to the channel exit than xenon. Presence of xenon is demonstrated to substantially increase the density of krypton atoms in metastable state and promote the ionization of krypton atoms. Both the enhanced mixture flowrate and Kr fraction push the ionization zone downstream. Increases in discharge voltage and discharge channel length shift the ionization zone upstream. For the overall performance, a longer discharge channel improves the thrust efficiency. It is mainly attributed to a more focused plume, an increase in the current efficiency from 80% to 84%, and a reduction in the electron current of about 28.57%. Therefore, under the condition of further optimized channel erosion issue, an extended discharge channel is recommended in the practical tests of Xe/Kr mixtures to compensate for the inefficiencies caused by krypton atoms.

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 0.95 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.

Coupling plasma physics and chemistry in the PIC model of electric propulsion: Application to an air-breathing, low-power Hall thruster

This work represents a first attempt to include the complex variety of electron-molecule processes in a full kinetic particle-in-cell/test particle Monte Carlo model for the plasma and neutral gas phase in a Hall thruster. Particular emphasis has been placed on Earth’s atmosphere species for the air-breathing concept. The coupling between the plasma and the gas phase is self-consistently captured by assuming the cold gas approximation and considering gas-wall and gas recycling from the walls due to ion neutralization. The results showed that, with air molecular propellants, all the most relevant thruster performance figures degraded relative to the nominal case using Xe propellant. The main reasons can be ascribed to a reduced ionization cross-section, a larger gas ionization mean free path due to lighter mass air species, and additional electron collisional power losses. While vibrational excitations power losses are negligible, dissociation and electronic excitations compete with the ionization channel. In addition, for molecular oxygen, the large dissociation leads to even faster atoms, further reducing their transit time inside the discharge channel. Future studies are needed to investigate the role of non-equilibrium vibrational kinetics and metastable states for stepwise ionization.

Experimental investigation of an external discharge very low anode power (< 20 W) hall thruster

A sub 20 W wall-less Hall Effect Thruster (HET) was developed at the Asher Space Research Institute (ASRI), Technion. In this work, an initial study of the thruster performance and underlying physics was conducted. It was found that the anode efficiency of the thruster was low (~ 1%), mainly due to the low mass utilization efficiency. Typical performance figures are 90 mu N of thrust, specific impulse of 90 s and anode efficiency of ~ 1% at 3–4 W anode power. The thruster far-field plume was analyzed using a retarding potential analyzer. It was found that the beam divergence was relatively low at 57.7^circ (for 95% of the beam current) compared to other wall-less HETs. The voltage utilization efficiency was 38% for a discharge voltage of 1 kV and a mass flow rate of 1 sccm xenon. We speculate that the leading driver to the low mass utilization efficiency is the small ionization fraction associated with these very low power wall-less devices. It was found that the beam efficiency can be over 90% at discharge power levels < 3 W, and decreases with power down to less than 50%.

Demonstration of One Hundred Kilogram Xenon Throughput by a Low-Power Hall Thruster

The Hall thruster component of the Ascendant Sub-Kilowatt Transcelestial Electric Propulsion System (known ASTRAEUS) successfully demonstrated a propellant throughput of over 100 kg of xenon and demonstrated a total of of impulse in a long-duration wear test. The thruster was operated for a total of 7205 h, with the test broken into five segments each with a different discharge power between 200 and 1350 W: 3291 h at 300 V/1000 W, 1739 h at 500 V/1000 W, 416 h at 400 V/600 W, 388 h at 200 V/200 W, and 1371 h at 300 V/1350 W. Telemetry and temperature measurements periodically recorded throughout the test revealed constant performance across each test segment and over the full test duration. Each segment was sufficiently long to obtain erosion rate data on the thruster by an optical profilometer. Subnominal magnetic field strengths were inadvertently applied to the thruster during the first 5 khr of the test due to an electrical ground-support equipment issue, which caused a downstream shift of the discharge plasma and higher erosion of the thruster’s pole covers than originally expected. This issue was corrected for the latter three test segments; however, the thruster performance and operational behavior at all reference operating conditions remained unaffected. The long-duration wear test was voluntarily terminated after the thruster processed 100.3 kg of xenon (Xe). This marks the world’s first long-duration wear test of a low-power magnetically shielded Hall thruster with a center-mounted hollow cathode wherein over 100 kg Xe throughput was demonstrated.

Magnetic field deflection in a 100 W Hall thruster with permanent magnets

The compact structure restrains the application of magnetic shielding in low-power Hall thrusters (LpHTs), leading to an asymmetric magnetic field or partial magnetic shielding of the channel wall. This study employs a trim coil to implement an asymmetric magnetic configuration in a 100 W laboratory Hall thruster. The locations of the maximum curvature of magnetic lines are deflected toward the inner and outer channel wall corresponding to the inward and outward deflected magnetic field configurations. Effects of the magnetic field deflection on the position of the ionization zone, efficiency of the thruster, discharge oscillations, and wall erosion are studied. Optical imaging and electrostatic probes are employed to monitor and scan the plasma beam. Experimental results show that the outward deflection induces a change in the magnetic mirror effect and alters the location of the ionization zone. The radial movement of the ionization zone away from the inner channel wall decreases the near-wall conductivity, reducing the electron current by 13.5% and promoting the current efficiency. Discharge oscillations are suppressed, and the propellant utilization efficiency is improved by 8.2%. Erosion of the channel wall shows an improvement of 40%. Generally, an outward deflected magnetic configuration can significantly improve the performance of LpHTs.