Low-power Hall Thruster Research Articles

Numerical study of the effect of radial magnetic field on performance of Hall thruster

The Hall-effect thruster has wide applications for commercial aerospace because of the high thrust density and simple structure. In order to further improve the performance of low-power Hall thruster and to solve the problem that the performance of low-power Hall thruster for low-orbit satellites is limited by the input power and maximum magnetic field intensity, the influence of radial magnetic field distribution in the discharge channel on the performance of the thruster is studied by numerical simulation and theoretical analysis in this work through changing the radial magnetic gradient on condition that the axial magnetic profile and the magnetic strength remain unchanged. The results show that the potential of the acceleration zone decreases with the increase of radial distance when the discharge parameters, propellant flow rate and axial magnetic field are unchanged. Therefore, the greater the radial magnetic field gradient near the inner wall of the thruster discharge channel, the greater the kinetic energy of the ions drifting along the axial direction to the thruster outlet, , and the greater the thrust of thruster. The research results of this work provide theoretical support for the magnetic field design and performance optimization of hall thrusters.

Time-synchronized laser-induced fluorescence in the near-field of a 600 Watt Hall thruster

We report on the results of an experimental campaign to measure time-varying velocity distributions in the near-field of a low power Hall thruster. We employ a sample-hold technique, enhanced by parallelizing the measurement hardware into several signal processing channels that vastly increases the data acquisition rate. The measurements are applied to study flow field dynamics in a commercial BHT-600 Hall thruster undergoing unforced breathing mode oscillations in the 44–49 kHz range. A very detailed experimental picture of the near-field emerges from these studies. The results indicate that velocity fluctuations lessen further downstream of the exit plane. Along the thruster axis where there is a general appearance of a central jet, there is evidence of a low velocity ion population in between the periodic bursts of high velocity ions, indicative of local ionization of neutrals outside of the thruster. One possible source of this residual ionization may be background chamber gas, which is not unexpected with the limited pumping capacity of ground test facilities.

HEATER OF THE HOLLOW CATHODE WITH LaB6 FOR OPERATION WITH HALL THRUSTERS

The article presents the results of design, development and laboratory tests at Space Electric Thruster Systems (Dnipro, Ukraine) of a hollow filament cathode based on lanthanum hexaboride (LaB6). The cathode is designed for joint operation with Hall thrusters of various input powers. The emitter material LaB6 is very promising, because has a high emissivity and low ionization potential. However, to realize all the advantages of considerate material, it is necessary to provide a sufficiently high operating temperature (1400 ° C). Therefore, when developing the filament cathode, special attention was focused on the development and research of one of the key elements of the filament cathode - the heater. Thermal calculations of the hollow cathode were carried out when it was heated with a heater, as well as when the cathode was operating in auto mode. The results of thermal calculations made it possible to identify the shortcomings of the cathode structure and parameters and determine the ways to eliminate them. Laboratory tests of a cathode with a different number of heater thermal shields have been carried out. The results obtained made it possible to provide the required operating temperature of the cathode emitter with a minimum number of heat shields. Laboratory tests of the cathode were carried out both in diode mode (to the external anode) and in conjunction with the ST-25 Hall thruster developed by the company. In the course of laboratory tests of the cathode, the power supplies for the discharge, the keeper, and the heater, which are components of the power processing unit of the Hall thruster, were used. In the frame of experimental studies, a cyclogram of the cathode starting process was obtained, which is necessary to form the Hall thruster starting sequence. The tests carried out confirmed the correctness of the selected technical solutions for the choice of the cathode and its heater. The developed cathode can be used both in low-power (up to 200 W) Hall thrusters and in thrusters with increased input power (up to 5 kW).

Spatial evolution characteristics of ion and electron flow for 300 W class low-power Hall thruster

The Faraday probe and cylindrical Langmuir probe were used to characterize the downstream ion and electron spatial evolution of a 300 W class low-power Hall thruster. The time-averaged ion current density, electron energy probability function, plasma potential, electron temperature, and electron density measurements were conducted at discharge voltages of 200–400 V and anode mass flow rates of 0.65 and 0.95 mg s−1 in the range of 100–500 mm axially and −100 to 100 mm radially downstream of the thruster. The results show that the ion and electron flows exhibit a bipolar diffusion characteristic along the radial direction. Meanwhile, the radial diffusion rate of ions in the plume is greater than the axial diffusion rate. The plasma potential decreases from 18 V at 100 mm axially from the thruster exit to 4 V at 500 mm axially and 100 mm radially. Correspondingly, the electron temperature decreases from 4.2 to 1.0 eV. The electron number density decreases from 2.6 × 1016 to 4 × 1014 m−3. A variable exponential relationship between electron temperature and electron density was observed from the measurements of electron energy probability distribution functions, with an adiabatic factor γ ranging between 1.3 and 1.4 (below the adiabatic value of 5/3). The adiabatic factor γ is considered to correlate with the anode mass flux and the spatial location of plasma, which suggests a possible dependence on the collision rate. These data are of great importance for plume model validation, improvement, plume effect evaluation, and protection mechanisms.

Experimental study on the influence of magnetic inclination in plume region on performance of low-power Hall Thruster

The magnetic field distribution of Hall thrusters has a crucial impact on its performance. Studies have mostly focused on the magnetic field distribution within the channel and near the channel exit, but less attention has been paid to the magnetic field distribution in the near-field plume region. Two different magnetic field distributions were designed. The influence of the magnetic inclination in the plume region on the performance of a low-power Hall thruster was investigated using a particle-in-cell (PIC) numerical simulation and experimental study. The results indicate that the magnetic field with inclined inward can effectively improve the performance of Hall thrusters, relatively increasing the thrust by 3.6%–9.6%, anode specific impulse by 3.8%–9.8%, and anode efficiency by 2.5%–5.5% at a fixed flow rate. It can also reduce the plume divergence angle, and the maximum decrease in the plume divergence half-angle is approximately 7°. The main reason is that the radial electric field in the plume region under an inward-inclined magnetic field is lower than that with an outward-inclined magnetic field, thus making the radial velocity of ions in the former magnetic field lower than that in the latter magnetic field, ultimately driving more ions to the thruster center.

The far-field plasma characteristics of LHT40 low-power Hall thruster for commercial aerospace applications

To achieve a better insight into the far-field plasma spatial distribution and evolution characteristics of the 300 W class low-power Hall thruster (LHT) for commercial aerospace applications, a dedicated and integrated plasma diagnostic system composed of seventeen Faraday cups (FC) and two triple Langmuir probes (TLP) is established to investigate the time-averaged in situ spatial distribution characteristics of far-field ions and electrons. The ion current density (ICD), plasma potential, plasma density, and electron temperature at 1000 mm downstream of 300 W class LHT for commercial aerospace applications in the azimuthal angle range of −90° to 90° were investigated under the conditions of different anode mass flow rates and discharge voltages. The results demonstrated that ICD, beam divergence angle, and mass utilization efficiency increased with increasing anode mass rate. The double-wings phenomenon was observed in the spatial distribution of ICD at large angles from the thruster axis, which is attributed to charge exchange collisions at increasing vacuum backpressure. The plasma electron temperature, electron density, and plasma potential parameters derived from the TLP decreased rapidly in the angle range from 0° to 30° and did not exhibit significant variations above 30°, which was also in good agreement with the results of the measured divergence angle of the FC. The discrepancy of average ion speed was calculated. The maximum error is better than 31.5% which checks the consistency between the TLP’s results and that of FC to some extent.

A low-current LaB6 open-end knife-edge emitter hollow cathode for low-power Hall thrusters

Heated thermionic cathodes based on LaB6 emitters are known to consume a relatively large amount of power during the heating part of the ignition phase. Miniaturized electric propulsion systems, and in particular low-power Hall thrusters, rely on the operation of self-sustained cathodes at low-current emission, typically below 1 A. At PSAC/SPC, an open-end knife-edge LaB6 emitter hollow cathode was developed and tested, denoted herein as the PSAC-KE cathode. Here, a detailed description of the new model’s design and comparison to the older version, the PSAC cathode, is included. Electrostatic simulations suggested that the open-end emitter geometry allowed for an enhanced electric field within the emitter region during cathode startup and steady-state operation as compared to the orificed emitter PSAC cathode. The PSAC-KE cathode’s thermal management was improved based on iterative thermal simulations driven by different cathode geometries and various combinations of materials. The new cathode was tested with xenon in diode mode with an external anode and triode mode with the external anode and the keeper electrode. Furthermore, results from the coupling tests with a low-power Hall thruster are presented. Cathode performance was monitored over a range of anode current from 0.1 to 1 A, mass flow rates from 0.057 to 0.3 mg s−1 and keeper current of 0.05, 0.1, and 0.15 A. The cathode’s thermal design was assessed with thermocouple measurements and compared with thermal simulations and results obtained for the PSAC cathode. The preliminary tests showed that the PSAC-KE cathode achieved ignition at low heating power below 35 W, self-sustained operation at 1 A when in standalone mode, and <1 A when against a low-power Hall thruster.

Experimental characterization of the C12A7 hollow cathode and its joint operation with a low-power Hall thruster

As essential components of an electrical propulsion system, traditional hollow cathodes, such as BaO-W and LaB6, have stringent requirements on the ignition procedure and propellant purity. The C12A7 hollow cathode with electride inserts, with a reported low work function and resistance to poisoning effects, has shown great potential in replacing traditional cathodes. To avoid the primary problems of frequent melting and degradation of the C12A7 emitter, which might be caused by the heaterless ignition process, a designed C12A7 hollow cathode with a heater was tested and successfully ignited with Xe propellant in this study. The emitting temperature of the C12A7 hollow cathode is estimated to be approximately 1000–1100 °C. The operational characteristics of the C12A7 hollow cathode in diode and triode configurations are presented, including the operating parameters when the discharge mode transition occurred. The C12A7 hollow cathode was subsequently jointly operated with a classic SPT-type Hall effect thruster (HET) for a nominal power of 400W. The HET was specially designed for Iodine propellant and has been tested with a LaB6 hollow cathode. It is successfully ignited by the C12A7 hollow cathode and works stably at 200 W for approximately 3 h. After more than 10 h of operation in a month with multiple exposures to the atmosphere, no obvious degradation of the C12A7 emitter is detected. The experimental results of the C12A7 hollow cathode show comparable performance and fast ignition and thus offer great potential in replacing traditional hollow cathodes, especially in low-power missions.

Optimization of a Faraday Cup Collimator for Electric Propulsion Device Beam Study: Case of a Hall Thruster

A Faraday cup (FC) is an instrument dedicated to current measurement in beams, jets and plasmas. It consists of a set of polarized electrodes mounted in such a way plasma sheath effect can be neglected, yielding accurate and reliable results. A FC is composed of three main parts, namely a collector or cup, which collects the current, a collimator, which defines the collection area and can contribute to limit electrons from entering the cup and a housing which protects the instrument from perturbation caused by the surrounding medium. In this paper, we provide experimental results of the effect of the collimator upon the measured ion current within the beam of a low-power Hall thruster. Different collimator materials, aperture diameters and polarization voltages are studied to determine the optimum design. Minimum dimension as well as appropriate materials are given as a conclusion in the case of low-power Hall thruster beam investigation.

Distinct discharge modes in micro Hall thruster plasmas

Two distinct discharge modes were observed in a 50 W-class micro-Hall thruster plasma under different operating conditions. A ball-shaped plasma with a broad plume (mode A) was observed at low mass flow rates (less than 0.37 mg s−1) over the entire operational anode voltage range (160–280 V). Raising the anode voltage beyond 200 V with the mass flow rate fixed (larger than 0.37 mg s−1) produced a narrow plume and stretched jet-like structure (mode B). In mode B, the thruster showed performance improvements in terms of thrust (3.8 mN vs 3.3 mN), specific impulse (913 s vs 800 s), and anode efficiency (28% vs 22%), with only a 2 W difference in the anode power (61 W in mode B and 59 W in mode A). This suggests that operation is more advantageous in mode B than in mode A for the utilization of such low-power Hall thrusters. Unique plume properties were observed in the two modes and considerable differences were measured in the Xe II ion acceleration structure, beam angle, and ionization rate as the mode changes, which were not reported in previous studies. Mode A exhibits an axially extended ion acceleration structure outside the discharge channel, where 75% of the final ion velocity is achieved at approximately 40 mm from the thruster exit, while most of the ion acceleration occurs within 10 mm from the thruster exit in mode B. Measurements show that the full width at half maximum of the Xe II ion energy distribution function, electron temperature, and Xe II emission intensity decreased after the plasma transitioned from mode A to mode B. Based on the optical emission spectroscopy, the ionization rate in the plasma plume decreased by 30%–41% after the mode change, which is likely related to the reduction of the beam angle and electron current by 24% and 30%, respectively.

Research on anode power supply of low-power Hall thruster

Based on the small power, small size, and high-efficiency micro-thrust requirements of micro-satellites, this paper mainly studies the anode power supply topology suitable for micro-satellites. The flyback converter and the secondary volta ge doubler structure are used to achieve high gain. At the same time, in order to improve the efficiency, the resonant soft switching technology is used to achieve valley opening through the resonance of the transformer leakage inductance and the capacitor to realize the soft switching of the switch tube and the diode. This article developed a 250W experimental prototype, and conducted a series of no-load and full-load tests under closed-loop steady-state operation, which can realize the zero-voltage turn-on of the switch tube and the zero-current turn-off of the diode. The efficiency is up to 93.7% under full load conditions, while the output ripple is small and there is no oscillation, and the device stress is small, which proves to be suitable for use in the anode power supply of micro satellites.

Effects of Magnetic Shielding Configuration on Discharge Characteristics and Performance of a 60 mm-Diameter Low-Power Hall Thruster

Lifetime is a main factor restraining the application of low-power Hall thruster. Magnetic shielding configuration is regarded as a promising method to prolong the lifespan of Hall thruster. Aiming to demonstrate the feasibility and effectiveness of magnetic shielding configuration applying on low-power Hall thruster, a 60-mm diameter Hall thruster in partial magnetic shielding configuration was designated. Both the numerical and experimental methods were used to investigate the discharge characteristics of the Hall thruster and help understand the mechanism behind. The maximum anode efficiency was achieved as high as 29.7% with 1.7 mg·s–1 anode mass flow and 320 V discharge voltage. To evaluate the effectiveness of the magnetic shielding used for low-power Hall thruster, a 2000 h lifetime test has been carried out and the results indicate that the erosion rate has been decreased below 0.2 μm·h–1.

Influences of Magnetic Flux Density on Discharge Characteristics of Low-Power Hall Thruster

ABSTRACT The design of the magnetic field is crucial to the performance of the hall thruster. This paper investigates the effects of magnetic field on the discharge characteristics of the low-power hall thruster in the method of combining simulation and experiment. Based on a 40-mm-diameter hall thruster (LHT-40) operating with hundreds of Watts, a two-dimensional axisymmetric model is established in fluid method to investigate the mechanism, how the magnetic field influences the performance of the thruster, and the experiments are also carried out to investigate the performance effected by magnetic field. Both the numerical and the experimental results indicate that there is an optimal magnetic flux density in which the thruster could achieve a higher efficiency compared with other conditions. And in order to maintain its high efficiency, the line of magnetic-curve inflexion inside the discharge channel should coincide with the center line of the channel to decrease the interaction between the plasma and the wall. The LHT-40 could generate about 12.2 mN of thrust at an optimal efficiency of 27% when the maximum magnetic flux density along the center line of the channel is about 200 Gs.

SITAEL HC1 Low-Current Hollow Cathode

SITAEL is active in the field of electric propulsion and is involved in the development of different thruster technologies—mainly Hall thrusters (HTs)—of power levels ranging from 100 W up to 20 kW. Low-power HTs are the most effective choice to perform orbit transfer, drag compensation, and de-orbiting maneuvers for small satellites. This paper is dedicated to the activities regarding HC1, the hollow cathode conceived for the 100-W-class Hall thruster under development at SITAEL. Successful test campaigns were performed and are described, with emphasis on the improvements in the cathode design after an extensive research and development phase. The results are presented and discussed, along with future developments of the ongoing activities.

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 mA 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.

Incoherent Thomson scattering measurements of electron properties in a conventional and magnetically-shielded Hall thruster

Reliable measurements of electron properties are key for the validation of simulations of Hall thrusters and similar plasma sources. In this paper, a recently-developed incoherent Thomson scattering diagnostic is successfully applied to electron property measurement in a low-power Hall thruster discharge. Investigations of electron properties along radial and azimuthal directions in both conventional and magnetically-shielded architectures are performed. The high diagnostic sensitivity gives access to electron property measurements in a plasma environment with densities as low as . Electron temperatures reaching several tens of eV and drift velocities on the order of are measured in the exit plane region. The impact of the magnetic field (intensity and direction) and discharge voltage are investigated.

Discharge mode transition in a Krypton-fed 1 A-class LaB6 cathode for low-power Hall thrusters for small satellites

Thermionic cathodes are essential for the operation of various electrostatic propulsion devices. They strongly influence the performance and lifetime of the propulsion system. In this study, a 1 A-class LaB6 laboratory model hollow cathode has been tested with krypton in diode and triode configurations in order to assess the cathode discharge mode transition behavior. Measurements have been performed over a range of krypton mass flow rates (0.1, 0.15, and 0.21 mg/s, or 1.6, 2.4, and 3.4 sccm), keeper (0.1, 0.15, and 0.2 A), and anode currents (0.1–1 A) at a fixed cathode-to-anode distance. Seven criteria were used to distinguish between the spot and plume mode operations. The results show that the mode transition in low-current cathodes may be a nonlinear phenomenon, and only some of the existing mode transition criteria can be used to accurately predict the spot/plume discharge regions at low emission currents.

Effects of gas supply direction on the discharge characteristics of a low-power Hall thruster

Micro-nano satellites show considerable potential in the application of low-orbit satellite networks, and low-power Hall effect thrusters (HETs) that provide the impetus for them have also attracted considerable attention worldwide. However, the relatively large surface-to-volume ratio of low-power HETs leads to significant loss of ions on the wall surface. After the channel width is increased or the ionization area is moved outward, ensuring sufficient ionization of the propellant is very difficult, and thus its integrated performance is not high. Thus, in this study, a 100-W permanent magnet HET with an aft-magnetic field is considered as the carrier to investigate the effect of gas supply direction on the thruster discharge characteristics through a simulation and experiments. Results indicate that under the same working conditions, adopting the circumferential direction gas supply mode effectively increases the density of the propellant in the ionization zone (up to 7%), improves the propellant ionization rate (up to 12.7%), and improves integrated discharge performance (the anode efficiency can be increased by up to 5.3%). In addition, the performance optimization effect of this mode is more evident under the working condition of low flow. When combined with an aft-magnetic field, this method is an effective means to ensure high-efficiency discharge of a low-power HET, which can be used as a reference for the miniaturization development of HETs.

Performance and plume characteristics of an 85 W class Hall thruster

An 85 W class Hall thruster with inner and outer electromagnetic coils was developed, and its thrust performance and plume characteristics were experimentally evaluated. The 85-W Hall thruster required a maximum magnetic flux density along the channel centerline above 24.3 mT to achieve stable and efficient operation above a discharge voltage of 225 V. A specific impulse of 1050 s, thrust-to-power ratio of 59.5 mN/kW, and anode efficiency of 0.306 were achieved at a discharge voltage of 225 V and a discharge power of 88.9 W under a background pressure of 7.9 × 10−4 Pa. Moreover, the thruster achieved throttling at a discharge power of 47.6–118.5 W with stable operation. The low anode efficiency of the 85-W Hall thruster was due to low propellant utilization and a large beam divergence. The short effective length for the propellant ionization owing to the narrow channel is deduced to be the cause for the low anode efficiency in low-power Hall thrusters. The mass utilization efficiency of the 85-W Hall thruster improved as the anode mass flow density increased. Therefore, an increase in the anode mass flow density is required to improve the performance of the 85-W Hall thruster.

Operation of a low-power Hall thruster: comparison between magnetically unshielded and shielded configuration

The operation of a 250 W low-power Hall thruster called ISCT-200 has been studied using a two-dimensional hybrid model. Two different magnetic field topologies have been tested. One topology is called the unshielded configuration and corresponds to a standard magnetic configuration with a quasi-radial magnetic field, and the other is called the magnetic shielding (MS) configuration where the zone of maximum of magnetic field is shifted in the near field plume. In the MS configuration, close to the channel walls, magnetic lines are forced to be parallel to walls. In the shielded configuration, the ionization takes place very close to the exhaust region and the acceleration occurs downstream the exit plane in the near field plume. The MS configuration reduces the erosion very effectively since the cooling of the electron temperature inside the channel strongly diminishes the sheath potential drop and consequently the kinetic energy of ions impacting on channel walls. The shift of ionization and acceleration regions towards the near field plume also contributes to the reduction of erosion. Calculations show very similar performances for both magnetic field configurations, with a larger than measured thrust in the shielded version of the thruster. Also, thanks to a larger electron temperature a larger fraction of doubly charged ions is found in the shielded magnetic configuration of the ICST-200.