Anode Mass Flow Rates Research Articles

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.

Study of beam divergence and thrust vector eccentricity characteristics of the Hall thruster based on dual Faraday probe array planes and its applications

The accurate knowledge of the thrust vector eccentricity and beam divergence characteristics of Hall thrusters are of significant engineering value for the beneficial integration and successful application of Hall thrusters on spacecraft. For the characteristics of the plume bipolar diffusion due to the annular discharge channel of the Hall thruster, a Gaussian-fitted method for thrust vector deviation angle and beam divergence of Hall thrusters based on dual Faraday probe array planes was proposed in respect of the Hall thruster beam characteristics. The results show that the ratios of the deviation between the maximum and minimum values of the beam divergence angle and the thrust vector eccentricity angle using a Gaussian fit to the optimized Faraday probe dual plane to the mean value are 1.4% and 11.5%, respectively. The optimized thrust vector eccentricity angle obtained has been substantially improved, by approximately 20%. The beam divergence angle calculated using a Gaussian fitting to the optimized Faraday probe dual plane is approximately identical to the non-optimized one. The beam divergence and thrust vector eccentricity angles for different anode mass flow rates were obtained by averaging the beam divergence and thrust vector eccentricity angles calculated by the dual-plane, Gaussian-fitted ion current density method for different cross-sections. The study not only allows for an immediate and effective tool for determining the design of thrust vector adjustment mechanisms of spacecraft with different power Hall thrusters but also for characterizing the 3D spatial distribution of the Hall thruster plume.

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.

Expanding the design freedom of the chamfered wall shape of a magnetically shielded Hall thruster

The channel wall shape of a magnetically shielded Hall thruster is typically designed to be tangential to the near-wall magnetic field line, which significantly constrains wall design freedom. In this study, the wall chamfer was expanded based upon tangent matching, and the effectiveness of magnetic shielding was determined based on the wall color after the experiments. The experiments indicated that the thruster would maintain the same high performance and no visible erosion at constant power levels and constant anode mass flow rate levels in the expanded chamfer wall, even if the wall was not perfectly tangent to the grazing line. The results provide new insights into the design of the topology of magnetically shielded Hall thrusters.

Multi-Objective Design Optimization of Cusped Field Thruster via Surrogate-Assisted Evolutionary Algorithms

Cusped field thrusters are an efficient cost-effective electric propulsion technology that offer promise for small spacecraft platforms. They offer various advantages, including simple design and low wall erosion, and hence long-life operation owing to high plasma confinement using permanent magnets over other electrostatic alternatives. Accurate physical modeling and performance characterization are crucial to achieve successful downscaling of cusped field thrusters while maintaining high performance. Multiobjective design optimization is performed by incorporating magnetic simulation coupled with an improved power balance model into evolutionary algorithms assisted by surrogate modeling, aiming to simultaneously maximize thrust, total efficiency, and specific impulse. Physical insights have been gained into key design factors to maximize thruster performance by probing into the plasma behavior inside the channel and the plume region. It has been found that altering the magnetic field distribution using an additional magnet can effectively enhance thruster performance by suppressing the plume divergence loss. Global sensitivity analysis has identified the anode current as the most influential design parameter on the performance parameters due to the active role it plays in the ionization process. Uncertainty analysis has found the considerable influence of the variations of anode potential and mass flow rate on those of thrust and specific impulse.

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.

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.

Background Pressure Effects on the Performance of a 20 kW Magnetically Shielded Hall Thruster Operating in Various Configurations

The paper reports the characterization results of a 20 kW-class magnetically shielded Hall thruster in three different configurations and operating with a centrally mounted cathode. The characterization was carried out at two different pumping speeds in SITAEL’s IV10 vacuum chamber, resulting in two different background pressure levels for each tested operating point. A linear behavior of discharge current and thrust values versus the anode mass flow rate was noticed for both pumping speeds levels and for all the three configurations. In addition, the thrust and discharge current values were always found to be lower at lower background pressure levels. From the performance levels, a preliminary estimate of the ingested mass flow rates was performed, and the values were then compared to a recently developed background flow model. The results suggested that, for this thruster and in the tested operating regimes, the change in performance due to background pressure could be ascribed not only to the ingestion of external mass flow coming from the chamber but also to other physical processes caused by the flux of residual background neutrals.

Cathode erosion site distributions in an applied-field magnetoplasmadynamic thruster

Erosion can influence cathode life, and is thus considered to be one of the main factors limiting the application of applied-field magnetoplasmadynamic thrusters. In this paper, erosion sites on graphite cathodes are studied so as to identify the influence of applied magnetic field and the ratio of propellant mass flow rate supplied from cathode and anode. The experiment results show that the application of applied magnetic field can significantly reduce the erosion rate of the cathode compared to that without magnetic field. The erosion sites on the cathode vary with the relative position of the convergent-divergent magnetic field, and are mainly distributed in the divergent part of the field. The erosion sites on the cathodes are found to be related to the propellant supply. The decreasing anode mass flow rate enlarges the range of erosion. These results are much helpful for the analysis of cathode erosion site location since they provide evidences of erosion mechanisms and point out the directions for further research.

Far-Field Plume Characterization of a 100-W Class Hall Thruster

The 100 W-class ISCT100-v2 Hall Thruster (HT) has been characterized in terms of far-field plume properties. By means of a Faraday Cup and a Retarding Potential Analyzer, both the ion current density and the ion energy distribution function have been measured over a 180 ∘ circular arc for different operating points. Measurements are compared to far-field plume characterizations performed with higher power Hall thrusters. The ion current density profiles remain unchanged whatever the HT input power, although an asymptotic limit is observed in the core of the plume at high discharge voltages and anode mass flow rates. In like manner, the ion energy distribution functions reveal that most of the beam energy is concentrated in the core of the plume [ − 40 ∘ ; 40 ∘ ] . Moreover, the fraction of low energy ion populations increases at large angles, owing to charge exchange and elastic collisions. Distinct plume regions are identified; they remain similar to the one described for high-power HTs. An efficiency analysis is also performed in terms of current utilization, mass utilization, and voltage utilization. The anode efficiency appears to be essentially affected by a low voltage utilization, the latter originating from the large surface-to-volume ratio inherent to low-power HTs. Experimental results also show that the background pressure clearly affects the plume structure and content.

High Fidelity Multi-Objective Design Optimization of a Downscaled Cusped Field Thruster

The Cusped Field Thruster (CFT) concept has demonstrated significantly improved performance over the Hall Effect Thruster and the Gridded Ion Thruster; however, little is understood about the complexities of the interactions and interdependencies of the geometrical, magnetic and ion beam properties of the thruster. This study applies an advanced design methodology combining a modified power distribution calculation and evolutionary algorithms assisted by surrogate modeling to a multi-objective design optimization for the performance optimization and characterization of the CFT. Optimization is performed for maximization of performance defined by five design parameters (i.e., anode voltage, anode current, mass flow rate, and magnet radii), simultaneously aiming to maximize three objectives; that is, thrust, efficiency and specific impulse. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the three design objectives and additional performance measures. The research indicates that the anode current and the Outer Magnet Radius have the greatest effect on the performance parameters. An optimal value for the anode current is determined, and a trend towards maximizing anode potential and mass flow rate is observed.

A type of cylindrical Hall thruster with a magnetically insulated anode

In this paper, a type of magnetically insulated anode structure is proposed for the design of a low-power cylindrical Hall thruster. The magnetic field distribution in the channel is guided by the magnetically insulated anode, altering the intersection status of the magnetic field line passing through the anode and wall. Experimental and simulation results show that a high potential is formed near the wall by the magnetically insulated anode. As the ionization moves towards the outlet, the energy and flux of the ions bombarding the channel wall can be reduced effectively. Due to the reduction in the bombardment of the wall from high-energy ions, the thrust and specific impulse greatly increase compared with those of the non-magnetically insulated anode. For anode mass flow rates of 0.3 and 0.35 mg s−1 and discharge voltages in the 100–200 V range, the thrust can be increased by more than 33% and the anode efficiency can be improved by more than 7%. Meanwhile, the length of the sputtering area is clearly reduced. The starting position of the sputtering area is in front of the magnetic pole, which can effectively prolong the service life of the thruster.

高周波プラズマ電子源を用いたホールスラスタの作動特性

Radio frequency (RF) plasma cathodes have been researched and developed as an electron source to remove the drawbacks with dispenser hollow cathodes from ion and Hall thrusters. In this paper, we evaluate the electron-emission performance of a RF plasma cathode and the operating characteristics of the combination of a 1-kW class anode-layer-type Hall thruster and the RF plasma cathode. For the RF plasma cathode, an anode current of 3.5A is obtained at an RF power of 140W, a xenon mass flow rate of 0.2mg/s. The anode current is sufficiently high to operate the Hall thruster. Normal ion beam extraction and neutralization for the combination were experimentally confirmed. The anode efficiency of the RF plasma cathode is comparable to that of a hollow cathode. Based on this result, the feasibility of the RF plasma cathode for the Hall thrusters in ground tests was confirmed. However, because of the high RF power consumption, the total efficiency of the RF plasma cathode is lower than that of the hollow cathode at anode mass flow rates of 2.0mg/s and 2.9mg/s. This result indicated that for space-related applications, it is important to reduce the power consumption of RF plasma cathodes.

Azimuthal Spoke Propagation in Hall Effect Thrusters

Spokes are azimuthally propagating perturbations in the plasma discharge of Hall effect thrusters (HETs) that travel in the E x B direction. The mechanisms for spoke formation are unknown, but their presence has been associated with improved thruster performance in some thrusters motivating a detailed investigation. The propagation of azimuthal spokes are investigated in a 6 kW HET by using high-speed imaging and azimuthally spaced probes. The spoke velocity is determined from high-speed image analysis using three methods with similar results. The spoke velocity for three discharge voltages (300, 400, and 450 V) and three anode mass flow rates (14.7, 19.5, and 25.2 mg/s) are between 1500 and 2200 m/s across a range of magnetic field settings. The spoke velocity is inversely dependent on magnetic field strength for lower B-fields and asymptotes at higher B-fields. Spoke velocities calculated from the probes are consistently higher by 30% or more. An empirically approximated dispersion relation of ω <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">α</sup> = v <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">ch</sub> <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">α</sup> k <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">θ</sub> <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">α</sup> - ω <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">ch</sub> <sup xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">α</sup> where α ≥ 1 yields a characteristic velocity that matches the ion acoustic speed for N5 eV electrons which exist in the near-anode and near-field plume regions of the discharge.

Performance of a Helicon Hall Thruster Operating with Xenon, Argon, and Nitrogen

The helicon Hall thruster is a two-stage thruster that was developed to investigate whether a radio frequency ionization stage can improve the overall efficiency of a Hall thruster operating at high thrust and low specific impulse. This paper describes an experiment that measured the single-stage and two-stage performance of the helicon Hall thruster operating at anode mass flow rates of xenon at 100–200 V discharge voltages, and also for of argon at 300 V, and of nitrogen at 200 V. The helicon Hall thruster performance during operation with argon and nitrogen is characterized by low beam divergence efficiency and low propellant utilization efficiency. During two-stage operation, the thrust of the helicon Hall thruster marginally increased with radio frequency power, but the propulsive efficiency and thrust-to-power both decreased with increasing radio frequency power. Probe diagnostics suggest that gains were realized by a slight increase in propellant efficiency, but that the rate of increase was not sufficient to overcome the increase in power.

Interior and Near-Wall Ion Velocity Distribution Functions in the H6 Hall Thruster

This paper presents the results of a two-axis laser-induced fluorescence velocimetry study of singly charged xenon in the interior and near the channel walls of the H6 Hall thruster. The thruster has a nominal operating discharge power of 6 kW and has been operated for before testing. It is operated under seven conditions spanning discharge voltages of 150–600 V and anode mass flow rates of . The mean velocity of the near-wall ion populations varies from 0.5 to with an angle of incidence relative to the wall surface that varies from 30 to 80 deg. In general, the mean velocity and angle of incidence are strongly correlated with the discharge voltage and weakly correlated with the anode mass flow rate. There is strong correlation between the axial locations where high energy (greater than a few tens of electron volts) ions exist and where erosion is discernible in the channel profile. This evidence supports the notion that erosion is negligible below a certain energy threshold. Potential application of this velocimetry technique for qualitative analysis of Hall thruster channel wall lifetime is also discussed.

Experimental Investigations of a Krypton Stationary Plasma Thruster

Stationary plasma thrusters are attractive electric propulsion systems for spacecrafts. The usual propellant is xenon. Among the other suggested propellants, krypton could be one of the best candidates. Most studies have been carried out with a Hall effect thruster previously designed for xenon. The ATON A-3 developed by MSTU MIREA (Moscow) initially defined for xenon has been optimized for krypton. The stable high-performance ATON A-3 operation in Kr has been achieved after optimization of its magnetic field configuration and its optimization in different parameters: length and width of the channel, buffer volume dimensions, mode of the cathode operation, and input parameters. For a voltage of 400 V and the anode mass flow rate of 2.5 mg/s the anode efficiency reaches 60% and the specific impulse reaches 2900 s under A-3 operating with Kr. The achieved performances under operation A-3 with Kr are presented and compared with performances obtained with Xe.

홀추력기 개념 설계를 위한 설계 공간 탐색

기존의 홀추력기 개념 설계는 비용이 많이 소요되며 과거의 데이터에 기반한 실험적 방법에 근거하고 있다. 본 연구에서는 이러한 홀추력기의 설계 과정을 향상시키기 위해 최근 제안된 수치적 방법에 근거하여 설계 목적에 적합한 설계 공간을 설정하고 이에 대한 설계 공간 탐색을 수행하였다. 설계 공간 탐색의 결과를 통해 기본적으로 주어진 설계 공간에 대한 성능 범위를 결정하였으며 성능간의 관계를 분석하였다. 세부적인 결과 분석을 통해 홀추력기의 주요한 설계 변수로는 양극에서의 질량유량과 방전전압임을 도출하였다. Current design process for the Hall effect thruster has relied on expensive experimental method based on the limited historical data. In this study, a proper design space for the Hall effect thruster is chosen and associated design space exploration is conducted based on a recently proposed numerical method in order to improve current design process. According to the results of the design space exploration, performance envelope is determined for the given design space and the correlations between performance metrics are analyzed. Further analysis shows that main factors in performances for the Hall effect thruster are the anode mass flow rate and the discharge voltage.

Self-consistent, one-dimensional analysis of the Hall effect thruster

This work achieves a self-consistent, robust, one-dimensional, macroscopic numerical model of the Hall effect thruster (HET) by dividing the discharge into two domains: (1) a collisionless electron diffusion region and (2) a collisional dominant electron diffusion region. Each region is modeled by governing equations based on the dominant physical characteristics. The solution for the discharge channel is created by matching the regional solutions at a common interface. The point validation for the model shows that the predicted HET performance is within 8% of the measured value. Furthermore, the proposed analysis is quite efficient in terms of a numerical perspective with a conventional PC wall time of 22 s for the nominal operating condition of the SPT-100. In addition, this paper shows that the numerical experiments for anomalous coefficients provide sufficient accuracy for performance metrics, closely matching the experimental data with correct values of the anomalous coefficients over a wide range of anode mass flow rates.

Neutral Flow Evolution in a Six-Kilowatt Hall Thruster

Laser-induced fluorescence velocimetry measurements obtained from the interior of a 6-kW Hall thruster using the Xe I 6s 3=2 02 ! 6p 3=2 2 transition at 823.4 nm (vacuum) and the Xe I 6s0 1=2 01 ! 6p0 3=2 2 transition at 834.9 nm (vacuum) are presented. The thruster is operated under seven conditions, with discharge voltages ranging from 150 to 600 V and anode mass flow rates ranging from 10 to 30 mg=s. Velocimetry results along the channel centerline show that the neutral propellant leaves the anode at a bulk axial velocity of 100 m=s and accelerates to 300–400 m=s by the exit plane. The temperature of the neutrals starts out at 1000–1600 K near the anode but cools down to 500–800 K near the exit plane. The anode mass flow rate appears to have a bigger influence on the bulk velocity and temperature of the neutrals than the discharge voltage. Radial sweeps across the channel exit plane revealed the presence of near-wall boundary layers approximately a fewmillimeters thick. The experimental results are compared with MONACO and HPHall-2 simulation results. The comparison suggests that the neutral-flow acceleration is due to a combination of wall thermalization, hydrodynamic, and ionization effects.