Propellant Utilization Efficiency Research Articles

Kinetic investigation of discharge performance for Xe, Kr, and Ar in a miniature ion thruster using a fast converging PIC-MCC-DSMC model

Abstract The present work develops a full particle-based model that couples the particle-in-cell plus Monte Carlo collision (PIC-MCC) simulation for plasma dynamics and the direct simulation Monte Carlo (DSMC) method for neutral dynamics in a synergistic iterative manner. This new model overcomes the slow convergence issue in the conventional direct coupling approach caused by the disparity of the time scales between the plasma and neutral dynamics. This model is applied to simulate the behavior of xenon (Xe) and its potential alternatives, krypton (Kr) and argon (Ar), in the discharge chamber of a miniature direct current (DC) ion thruster. The results show that a stable discharge is difficult to achieve for Kr and Ar under the operating conditions optimal for Xe. While increasing the discharge voltage can effectively improve the stability of discharge for Kr and Ar, other common strategies such as changing the magnetic field strength, propellant flow rate, and cathode current are not successful. The propellant utilization efficiency and discharge efficiency are affected by both discharge voltage and propellant flow rate. A maximum utilization efficiency and an optimal discharge efficiency are observed for all three propellants, with the values decreasing in the order of Xe, Kr, and Ar. Moreover, the discharge voltage corresponding to the optimal efficiency is inversely proportional to the square root of the propellant mass, indicating that the ion diffusional loss to the wall, rather than the ionization energy, is the dominant factor affecting the discharge performance for alternative propellants in a miniature DC thruster.

Improving the specific impulse of Hall thrusters using a wide channel design

This paper proposes increasing the width of the discharge channel to improve the specific impulse of Hall thrusters. In our study, we significantly increased the achievable channel width by optimizing the magnetic circuit topology, and the design freedom of the Hall thruster was improved. A prototype of a wide-channel high-specific-impulse Hall thruster was designed and experimentally verified. The effects of channel width on the discharge process and the characteristics of an aft-loaded magnetic field Hall thruster were investigated. The high-intensity luminous region did not diffuse radially with an increase in the channel width, and the plasma was always concentrated near the channel centerline. A wider channel is more conducive for improving the propellant utilization efficiency, thereby effectively improving the specific impulse of the Hall thruster. By optimizing the discharge channel width, the upper limit of the discharge voltage was increased, and the maximum specific impulse was increased from 3170 to 3794 s, a relative improvement of 20%. The effect of the specific impulse improvement at the same discharge voltage was analyzed and discussed. The results of this study are significant for the design of high-specific-impulse Hall thrusters.

Thrust measurements of a waveguide electron cyclotron resonance thruster

Direct thrust measurements are performed on a circular waveguide electron cyclotron resonance (ECR) thruster working at 5.8 GHz using a pendulum thrust balance with mechanically amplified displacement. Thrust levels between 1 and 3.5 mN are found for power levels in the range of 60–350 W and xenon flow rate between 2 and 8 SCCM. A maximum thrust efficiency of 3.5% is reached at 2 SCCM and 60 W. Plasma plume diagnostics are used to estimate the thruster partial efficiencies to understand the main losses, and to perform a comparative analysis between directly and indirectly measured thrust. Results suggest that the low energy conversion efficiency and propellant utilization efficiency (<6.4% and < 53%, respectively) are the key factors spoiling the ECR thruster performance. Finally, retarding potential analyzer measurements show the presence of energetic electrons with energy tail up to about 300 eV.

Optimizing a Hall thruster with aft-loaded magnetic field by aft-loading design of gas flow

Electric propulsion systems with high-specific impulses have replaced chemical propulsion systems in many space missions. The aft-loaded magnetic field technology of Hall thrusters moves the ionization and acceleration processes downstream of the channel by varying the magnetic field distribution. Consequently, ion bombardment of the channel walls is inhibited, and the service life of the Hall thrusters is substantially increased. However, neutral gas leaks because of the rapid diffusion near the channel outlet, where strong ionization occurs. Thus, the propellant atoms are ionized insufficiently, resulting in a decline in the performance. Therefore, an aft-loading design for the neutral gas flow field was developed to optimization distribution of the neutral gas in this study. A 1.5 kW Hall thruster was verified experimentally. The experimental results show that the aft-loaded neutral gas flow field significantly improved the ionization and acceleration processes. The propellant utilization efficiency increased by 4–8% within the discharge voltage range of 300–900 V. The ion energy distribution moved toward the high-energy direction significantly, promoting an overall improvement in the performance, with a thrust increase of 7.3–11.5 %. The anode efficiency increased from 49–57 % to 58–61 %. The results of this study are significant for the design of high-performance, long-life Hall thrusters.

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.

Influence Mechanism of Magnetic Field and Wave Modes on Helicon Plasma Thruster

The mechanism of energy deposition on a helicon plasma thruster under a focused magnetic field is different with a uniform magnetic field, and energy coupling efficiency has greatly improved. This paper studies the spatial distribution of power deposition under different magnetic field configurations. The results show that the magnetic field configuration can change the spatial distribution of energy deposition. A focused magnetic field configuration can couple more energy on the antenna downstream and reduce energy loss during plasma transport, which leads to improved propellant utilization efficiency and enhanced thruster performance. Then this paper studies the energy coupling efficiency and energy distribution characteristics for ionization and acceleration under different W modes, including W1 and W2. The results show that, in the W1 mode, the energy is mainly focused on ion acceleration, resulting in higher ion beam energy. In contrast, the W2 mode is focused on propellant ionization, resulting in greater efficiency of propellant utilization. The W2 mode demonstrates a higher energy coupling efficiency and has superior thruster performance compared to the W1 mode, with a thrust increase of about 1.6 times. The paper proposes several suggestions to improve the thrust-to-weight ratio and specific impulse of a helicon plasma thruster, which provides theoretical support for engineering applications.

Magnetic field optimization of hall thruster with large height-radius ratio for high specific impulse operation

Hall thrusters with large height-radius ratio not only have incalculable application values in reducing the volume and weight of thrusters, but also have the potential advantages of higher discharge performance and longer service life. However, the lower propellant density in the main ionization zone and the higher electron temperature in the channel aggravate the loss of propellant and current under high voltage, and significantly reduce the discharge efficiency under high specific impulse mode. To improve the discharge performance of Hall thrusters with large height-radius ratio under high voltage, an optimization scheme of internally loaded magnetic field was proposed in this work. The simulation results show that under the internally loaded magnetic field, both the ionization zone and the acceleration zone move toward the inside of the channel. Although the ion loss on the walls increases, the higher propellant density at the channel upstream greatly promotes the increase of ionization rate and significantly improves the propellant utilization efficiency. The second zone crossed by magnetic field lines in the channel can be established by the internally loaded magnetic field, which enhances the magnetic field intensity on the inner and outer walls, and reduces the electron temperature near the channel outlet significantly. So that the axial conduction of electrons is effectively restrained and the current utilization efficiency is greatly improved. With the introduction of internally loaded magnetic field, the total efficiency of HEP-1350PM can be increased by 7.2% at 400 V. Moreover, the performance optimization effect brought by the internally loaded magnetic field will be gradually amplified with the increase of discharge voltage, which makes the Hall thruster with large height-radius ratio expected to achieve high-efficiency discharge under higher specific impulse.

Accuracy of using metastable state measurements in laser-induced fluorescence diagnostics of xenon ion velocity in Hall thrusters

Laser-induced fluorescence measurements of singly-charged xenon ion velocities in Hall thrusters typically target metastable states due to lack of available laser technology for exciting the ground state. The measured velocity distribution of these metastable ions are assumed to reflect the ground state ion behavior. However, this assumption has not been experimentally verified. To investigate the accuracy of this assumption, a recently developed xenon ion (Xe II) collisional-radiative model is combined with a 1D fluid model for ions, using plasma parameters from higher fidelity simulations of each thruster, to calculate the metastable and ground state ion velocities as a function of position along the channel centerline. For the HERMeS and SPT-100 thruster channel centerlines, differences up to 0.5 km s−1 were observed between the metastable and ground state ion velocities. For the HERMeS thruster, the difference between the metastable and ground state velocities is less than 150 m s−1 within one channel length of the channel exit, but increases thereafter due to charge exchange (CEX) that reduces the mean velocity of the ground state ions. While both the ground state ions and metastable state ions experience the same acceleration by the electric field, these small velocity differences arise because ionization and CEX directly into these states from the slower neutral ground state can reduce their mean velocities by different amounts. Therefore, the velocity discrepancy may be larger for thrusters with lower propellant utilization efficiency and higher neutral density. For example, differences up to 1.7 km s−1 were calculated on the HET-P70 thruster channel centerline. Note that although the creation of slow ions can influence the mean velocity, the most probable velocity should be unaffected by these processes.

Development and validation of an iodine plasma model for gridded ion thrusters

Iodine is emerging as an attractive alternative propellant to xenon for several electric propulsion technologies due to its significantly lower cost and its ability to be stored unpressurized as a solid. Because of the more complex reaction processes and energy-loss channels in iodine plasmas however, as well as the historical lack of reliable collision cross-section data, the development of accurate theoretical and numerical models has been hindered. Using recently calculated theoretical cross-sections, we present an iodine plasma model and perform a comparison with experimental data obtained from an iodine-fuelled gridded ion thruster. The model is in reasonable agreement with experimental measurements of the ion beam current, propellant mass utilization efficiency, and ion beam composition, and is able to quantitatively and qualitatively reproduce system behaviour as the input mass flow rate and RF power are varied. In addition, both the model and experiment show that the use of iodine can lead to a performance enhancement when compared with xenon. This occurs because of the combination of different iodine reaction processes, collision cross-section values, and inelastic energy thresholds which result in lower collisional energy losses, as well as an increased antenna-plasma power transfer efficiency for thrusters using a radio-frequency inductive coil.

The effect of magnet core size on the performance of a cylindrical Hall thruster with the near-anode cusp magnetic field

Effect of magnet core size on the performance of a cylindrical Hall thruster with the near-anode magnetic field has been investigated. Experiments found that extending magnetic core and rising discharge voltage show similar effects on the ionization process. When the magnetic core ends flush with the anode and discharge voltage is fixed at 240 V, the upstream ionization related with the electron E × B drift is hardly established under an approximately axial magnetic field and the ionization region is located at the near-axial magnetic mirror field, the thruster manifests the highest propellant utilization efficiency (as high as 1.43), the lowest current utilization efficiency (0.54) and beam divergence efficiency (0.52). When the magnet core extends towards the thruster exhaust, the enhancement of upstream ionization results in significant enhancements in the current utilization efficiency (by 14.6%) and the beam divergence efficiency (by 27.1%). However, the reduction of magnetic mirror ionization also led to a reduction of propellant utilization efficiency (by 32.1%). Therefore, the anode efficiency varies non-monotonically (climbing up and then declining) and has an optimal solution: shorter magnet core performs higher anode efficiency in low voltage condition and longer magnet core performs higher anode efficiency in high voltage condition.

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.

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

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

A study on the ionization mechanisms in a miniaturized cylindrical Hall thruster

Ionization process plays a crucial role in the plasma discharge of cylindrical Hall thruster. In this paper, the ionization characteristics for the two types of magnetic field configurations involving the opposite polarity configuration and same polarity configuration are studied by adopting a movable anode and then the ionization mechanism is analyzed by the ion energy characteristics decomposing method. Experimental results show that when the anode is flush with the inner wall, the dominant ionization regions are both located in the channel center for these two configurations. While as the anode moves upstream, the ionization region keeps constant for the same polarity configuration. While another ionization could take place actively in the expanding annular region for the opposite polarity configuration, which could be the main cause for the enhancement of propellant utilization efficiency and the significant narrowing of the IEDF curves with the bimodal distribution. Additionally, this ion energy distribution function is decomposed by extracting the characteristic function, the calculated results indicate that the proportion of the ion current produced by this additional ionization mechanism reaches 72%, which indicates that this additional ionization acting as the dominant ionization mechanism could be the leading cause for the jet-like plume.

Effects of the peak magnetic field location on discharge performance of a 100-W Hall thruster with large gradient magnetic field

Traditional Hall thrusters can ionize gas inside the channel and accelerate ions outside it by moving the peak magnetic field out of the channel, which improves the thruster performance. In this study, the effects of the peak magnetic field location on the discharge characteristics of a 100-W Hall thruster with large-gradient magnetic field were investigated by changing the ratio of the magnetic field intensity at the channel exit to the peak magnetic field intensity (Brexit/Brmax). The experimental results show that the discharge performance decreases significantly when Brexit/Brmax < 1. Moreover, the PIC numerical simulation and neutral gas simulation results demonstrate that the ionization rate inside the channel reduces and the ionization proportion outside the channel increases when Brexit/Brmax decreases, thereby reducing the performance. When Brexit/Brmax = 1, under the anode flow of 5.5 sccm and discharge voltage of 250 V, the thrust, anode specific impulse, propellant utilization efficiency, and anode efficiency of the 100-W thruster were 6.4 mN, 1210 s, 68.9%, and 38.5%, respectively, which represent respective improvements of 16.2%, 16.4%, 13.0%, and 30.0% (relative value) compared with the condition when Brexit/Brmax = 0.9. These results provide a reference for the performance optimization of 100-W Hall thrusters with a large-gradient magnetic field.

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

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

Determining energetic characteristics and selecting environmentally friendly components for solid rocket propellants at the early stages of design

This paper has investigated the possibility to theoretically calculate a value of the specific impulse for highly energetic compositions using only two parameters – the heat of the reaction and the number of moles of gaseous decomposition reaction products. Specific impulse is one of the most important energetic characteristics of rocket propellant. It demonstrates the level of achieving the value of engine thrust and propellant utilization efficiency. Determining the specific impulse experimentally is a complex task that requires meeting special conditions. For the stage of synthesis of new promising components, the comparative analysis of energetic characteristics, forecasting the value of specific impulse, especially relevant are calculation methods. Most of these methods were first developed to determine the energetic characteristics of explosives. Since explosives and rocket propellants in many cases have similar energy content and similar chemical composition, some estimation methods can be used to assess the specific impulse of solid rocket propellant. The specific impulse has been calculated for 45 compositions based on environmentally friendly oxidizers (ammonium dinitramide, hydrazinium nitroformate, hexanitrohexaazaisowurtzitane) and polymer binders polybutadiene with terminal hydroxyl groups, glycidylazide polymer, poly-3-nitratomethyl-3-methyloxetane). It was established that the estimation data obtained correlate well with literary data. Deviation of the derived values of the specific impulse from those reported in the literature is from 0.4 % to 1.8 %. The calculation results could be used for preliminary forecasting of energetic characteristics for highly energetic compositions, selecting the most promising components, as well as their ratios.

Neutral atom density measurements of xenon plasma inside a μ10 microwave ion thruster using two-photon laser-induced fluorescence spectroscopy

This paper reports measurements of the xenon ground state and excited state densities inside a μ10 microwave ion thruster. This thruster exhibits a 40% thrust enhancement upon changing from the waveguide to the discharge chamber propellant injection mode. In the present work, the associated mechanism was quantitatively evaluated using two-photon laser induced fluorescence (TALIF) spectroscopy to monitor the thruster waveguide. The 834.7 nm emission from excited state xenon was investigated with a 224.3 nm dye laser to excite the Xe I 5p61 S0 6pʹ [3/2]2 state, compared with the emission without the laser. The resulting data confirm that the neutral density exhibits a linear relationship with the propellant flow rate in the cold gas and ionized state, while the ion acceleration decreases the neutral density by the same order of magnitude as the propellant utilization efficiency is changed. As the propellant flow rate increases, the collisions of neutrals that generate excited states occur in the waveguide and, when this process plateaus, the ground state emission suddenly increases. Propellant injection from the discharge chamber is evidently effective at suppressing collisions with electrons in the waveguide that generate excited states and that potentially interfere with microwave propagation.

Improving the Hall Thruster Global Efficiency over a Wide Flow-Rate Range

In this study, a 1.35 kW HEP-100X Hall thruster is used to explore effective methods for improving the global efficiency of a Hall thruster across a wide range of flow rates, from the perspective of optimizing the channel configuration. Experimental and simulation results indicate that an appropriate increase in the length of the straight segment in the variable cross section channel can increase the neutral gas density as well as move the high-gas-density zone toward the channel outlet, thus increasing the ionization rate and moving the main ionization zone outward. At a low flow rate, the propellant utilization efficiency and anode efficiency can be improved by 2.4 and 6.1%, respectively. In addition, compared with a traditional Hall thruster of the same power with a uniform cross section channel, the HEP-100X Hall thruster equipped with the optimal channel exhibits significantly improved performance, and the lower limit of the anode efficiency is increased from 43.3 to 46.6% across a wide range of flow rates. This study resolves the issue of low efficiency of a Hall thruster at a low flow rate and also provides an effective solution for efficient discharge of the Hall thruster across a wide range of flow rates.

Water and xenon ECR ion thruster—comparison in global model and experiment

Gridded ion thrusters are one of the most commonly used types of electric propulsion, and alternative propellants have been studied for miniature ion thrusters to meet the demand of propulsion systems for micro-/nano-satellites. Water is a candidate as an alternative non-pressurized propellant for a CubeSat thruster. It is consistent with the CubeSat concept of short-term and low-cost development. In this paper, the characteristics of a miniature water ion thruster were compared with those of a xenon one using a global model and experiments. The dependence of the performance on the mass flow rate and the input microwave power was examined, and the effects of dissociation and doubly charged ions were directly measured by a quadrupole mass spectrometer. The estimates on the model were compared against experimental results for both propellants, and the performance of the thruster operating on xenon propellant was compared to the performance operating on water propellant. In the comparison between the estimates and the experimental results, the two differences were discussed: the one between water and xenon and the other from the experimental result in both cases. A performance decrease in the propellant utilization efficiency and the specific impulse cannot be avoided when using water as a propellant in an ion thruster. However, the ion production cost did not increase, and it showed the capability of water ion thruster for CubeSat application taking advantage of safety, low cost, non-pressurized system, and human-friendliness of water when used as a propellant.

Numerical simulation study on the influence of channel geometry on discharge characteristics of low-power magnetically shielded Hall thrusters

By adjusting the relative position of the magnetic screen and the discharge channel, two types of 200 W low-power magnetically shielded Hall thrusters with three different Brexit/Brmax ratios (the ratio of the radial magnetic field magnitude at the channel exit to the maximum radial magnetic field magnitude, taken at the channel mean diameter) using straight wall channel and chamfered wall channel geometries are designed. The discharge characteristics of the plasma under the two types of Hall thrusters are studied using the particle-in-cell (PIC) simulation, and the influence laws of the channel geometries on the thrust, specific impulse, and anode efficiency of the 200 W Hall thrusters are verified by the corresponding experimental testing. Both the results indicate that compared with the straight wall channel geometry of 200 W magnetically shielded Hall thrusters, the chamfered wall channel geometry will reduce the neutral gas density in the discharge channel, causing the propellant utilization efficiency to be lower, thus decreasing the ionization rate, ultimately reducing the performance of thruster. It explains and clarifies the reason for the dark gap phenomenon and the lower performance of the low-power magnetically shielded Hall thrusters.