Power Hall Thrusters Research Articles

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.

Trends in mass utilization of a magnetically shielded Hall thruster operating on xenon and krypton

The trends in mass utilization with increasing discharge voltage and current are investigated for a magnetically shielded Hall thruster operating on xenon and krypton. A 9 kW class shielded thruster is operated with discharge voltages from 300 to 600 V and discharge currents from 15 to 30 A on xenon and krypton. Experimental measurements of discharge current, thrust, anode efficiency, and ion velocity as a function of axial position are used to calibrate a multi-fluid 2D Hall thruster code at all operating conditions. The results of these calibrated simulations are employed to interrogate the plasma properties inside the thruster channel. A simplified 0D model for mass utilization evaluated on spatial averages of the simulated plasma parameters is employed to interpret the response of this efficiency mode with power for each propellant. It is found that with both higher voltage and current, mass utilization increases for both gases and their relative gap in this efficiency decreases. This can be attributed to the higher plasma densities and ionization rate coefficients at high voltage, and solely to higher plasma densities at high current. The driving factors for the increase in mass utilization are examined in the context of its nonlinear response to internal plasma properties. The behavior of mass utilization is also discussed in context of the gap in overall efficiency between the propellants. Finally, the implications of these results for improving the performance of high power Hall thrusters operating on krypton are examined.

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.

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.

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.

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%.

500-A LaB6 Hollow cathode for high power electric thrusters

NASA continues to develop higher power ion and Hall thrusters for deep space human and robotic missions that require higher thrust and specific impulse than available in current thrusters. The NASA XR-100 nested Hall thruster successfully operated at 100 kW of power; driven by a JPL LaB6 hollow cathode that produced up to 250 A of discharge current in the thruster and 25–350A in the cathode test facility. The next generation of this LaB6 hollow cathode has been designed and built at JPL and tested at up to 500 A of discharge current. Extending the current to this high level was challenging, with iterations in the design necessary to provide adequate thermal isolation of the insert to lower the heater power required to ignite, to achieve sufficient keeper discharge current to aid in the startup, and to maintain the cathode orifice plate at acceptable temperatures during very high current operation. The 2nd generation version of the cathode successfully achieved 500 A of discharge current operation with low high-energy ion content. Discussions of the plasma instabilities observed in the plume, the plasma penetration into the insert region, and cathode life are provided.

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.

Use of an accelerated testing method to characterize inner front pole cover erosion in a high power Hall thruster

Erosion characteristics on the cover of the inner front pole in a 12.5 kW Hall thruster were measured over a wide range of operating conditions in tests of 6–14 h duration using an accelerated test method and a very sensitive, radioactive tracer-based erosion diagnostic. The operating points included the nominal 300–600 V conditions on a constant 20.8 A throttle curve and conditions at other currents spanning the throttling envelope and with varying magnetic field strength, facility pressure, and discharge voltage oscillation amplitude. Surprisingly, the results show that the highest wear rates occur at the lowest voltages and currents. The wear rates were insensitive to discharge voltage ripple but increased monotonically with magnetic field strength, particularly near the inner radius of the pole cover. The inner region was also sensitive to facility pressure, showing lower rates at a higher pressure level. Separate experiments in which the energy distributions of ions generated by the hollow cathode were measured suggest that the cathode plume may be a source of energetic ions responsible for some of the erosion trends, in addition to ions originating in the thruster plume.

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.

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.

The far-field plasma characterization in a 600 W 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 a 600 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 at 300 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.

Thrust performance, propellant ionization, and thruster erosion of an external discharge plasma thruster

It is quite a challenge to design low power Hall thrusters with a long lifetime and high efficiency because of the large surface area to volume ratio and physical limits to the magnetic circuit miniaturization. As a potential solution to this problem, we experimentally investigated the external discharge plasma thruster (XPT). The XPT produces and sustains a plasma discharge completely in the open space outside of the thruster structure through a magnetic mirror configuration. It eliminates the very fundamental component of Hall thrusters, discharge channel side walls, and its magnetic circuit consists solely of a pair of hollow cylindrical permanent magnets. Thrust, low frequency discharge current oscillation, ion beam current, and plasma property measurements were conducted to characterize the manufactured prototype thruster for the proof of concept. The thrust performance, propellant ionization, and thruster erosion were discussed. Thrust generated by the XPT was on par with conventional Hall thrusters [stationary plasma thruster (SPT) or thruster with anode layer] at the same power level (∼11 mN at 250 W with 25% anode efficiency without any optimization), and discharge current had SPT-level stability (Δ &amp;lt; 0.2). Faraday probe measurements revealed that ion beams are finely collimated, and plumes have Gaussian distributions. Mass utilization efficiencies, beam utilization efficiencies, and plume divergence efficiencies ranged from 28 to 62%, 78 to 99%, and 40 to 48%, respectively. Electron densities and electron temperatures were found to reach 4 × 1018 m−3 (∂ne/ne = ±52%) and 15 eV (∂Te/Te = ±10%–30%), respectively, at 10 mm axial distance from the anode centerline. An ionization mean free path analysis revealed that electron density in the ionization region is substantially higher than the conventional Hall thrusters, which explain why the XPT is as efficient as conventional ones even without a physical ionization chamber. Our findings propose an alternative approach for low power Hall thruster design and provide a successful proof of concept experiment of the XPT.

Performance Validation and Plume Divergence of the CAM200 Low Power Hall Thruster

Performance Validation and Plume Divergence of the CAM200 Low Power Hall Thruster

Innovation in the development of plasma propulsion devices in Israel

In this paper we review plasma propulsion development approach which focuses on innovation. We then bring the example of the state of Israel in general, and Rafael in particular, and show how it has adopted an innovative approach to develop a low power Hall thruster and a low current cathode. To present one special test-case of innovation we elaborate upon the development process of a heaterless hollow cathode that was developed at Rafael. In particular, by presenting the cathode characterization and wear test results we demonstrate that the heaterless cathode developed has a sufficiently wide operational range and may operate continuously for 1,500 hours without any measurable degradation in performance.

Perturbations induced by electrostatic probe in the discharge of Hall thrusters.

Emissive and Langmuir probes are two widely used plasma diagnostic techniques that, when used properly, give access to a wide range of information on the plasma's ions and electrons. We show here that their use in small and medium power Hall thrusters produces large perturbations in the discharge characteristics. Potential measurements performed by both probes and non-invasive Laser Induced Fluorescence (LIF) spectroscopy highlight significant discrepancies in the discharge profile. This phenomenon is observed both in the 200 W and the 1.5 kW-class thrusters. In order to have a better understanding of these perturbations, ion velocity distribution functions are acquired by LIF spectroscopy at different positions in the smaller thruster, with and without the probes. Emissive probes are shown to produce the biggest perturbation, shifting the acceleration region upstream. The probe insertion is also shown to have significant effect on both the average discharge current, increasing it by as much as 30%, and its harmonic content in both amplitude and spectrum. These perturbations appear as the probe tip passes a threshold located between 0 and 5 mm downstream of the thruster exit plane.

The Development of CAM200 - Low Power Hall Thruster

The Development of CAM200 - Low Power Hall Thruster

Compact high-speed reciprocating probe system for measurements in a Hall thruster discharge and plume

A compact high-speed reciprocating probe system has been developed in order to perform measurements of the plasma parameters by means of electrostatic probes in the discharge and the plume of a Hall thruster. The system is based on a piezoelectric linear drive that can achieve a speed of up to 350 mm/s over a travel range of 90 mm. Due to the high velocity of the linear drive the probe can be rapidly moved in and out the measurement region in order to minimize perturbation of the thruster discharge due to sputtering of probe material. To demonstrate the impact of the new system, a heated emissive probe, installed on the high-speed translation stage, was used to measure the plasma potential and the electron temperature in the near-field plume of a low power Hall thruster.