Performance investigation of a low-power Hall thruster fed on iodine propellant

The common propellants used for electric thrusters, such as xenon and krypton, are rare, expensive,and difficult to acquire. Solid iodine attracts much attention with the advantages of low cost,extensive availability, low vapor pressure, and ionization potential. The performance of a lowpower iodine-fed Hall thruster matched with a xenon-fed cathode is investigated across a broad range of operation conditions. Regulation of the iodine vapor's mass flow rates is stably achieved by using a temperature control method of the iodine reservoir. The thrust measurements are finished utilizing a thrust target during the tests. Results show that thrust and anode-specific impulse increase approximately linearly with the increasing iodine mass flow rate.At the nominal power of 200 W class, iodine mass flow rates are 0.62 and 0.93 mg/s, thrusts are7.19 and 7.58 m N, anode specific impulses are 1184 and 826 s, anode efficiencies are 20.8%and 14.5%, and thrust to power ratios are 35.9 and 37.9 m N/k W under the conditions of 250 V,0.8 A and 200 V, 1.0 A, respectively. The operating characteristics of iodine-fed Hall thruster are analyzed in different states. Further work on the measurements of plasma characteristics and experimental optimization will be carried out.

Growth mechanism and characteristics of electron drift instability in Hall thruster with different propellant types

The existence of a significant electron drift instability(EDI) in the Hall thruster is considered as one of the possible causes of the abnormal increase in axial electron mobility near the outlet of the channel. In recent years, extensive simulation research on the characteristics of EDI has been conducted, but the excitation mechanism and growth mechanism of EDI in linear stage and nonlinear stage remain unclear. In this work, a one-dimensional PIC model in the azimuthal direction of the thruster near-exit region is established to gain further insights into the mechanism of the EDI in detail, and the effects of different types of propellants on EDI characteristics are discussed. The changes in axial electron transport caused by EDI under different types of propellants and electromagnetic field strengths are also examined. The results indicate that EDI undergoes a short linear growth phase before transitioning to the nonlinear phase and finally reaching saturation through the ion Landau damping. The EDI drives a significant ion heating in the azimuthal direction through electron–ion friction before entering the quasi-steady state, which increases the axial mobility of the electrons. Using lighter atomic weight propellant can effectively suppress the oscillation amplitude of EDI, but it will increase the linear growth rate, frequency, and phase velocity of EDI. Compared with the classical mobility, the axial electron mobility under the EDI increases by three orders of magnitude, which is consistent with experimental phenomena. The change of propellant type is insufficient to significantly change the axial electron mobility. It is also found that the collisions between electrons and neutral gasescan significantly affect the axial electron mobility under the influence of EDI, and lead the strength of the electric field to increase and the strength of the magnetic field to decrease, thereby both effectively suppressing the axial transport of electrons.

Influence of acceleration stage electrode voltage on the performance of double-stage Hall effect thruster with adjustable zero magnetic point

The configuration of electrode voltage and zero magnetic point position has a significant impact on the performance of the double-stage Hall effect thruster. A 2D-3V model is established based on the two-magnetic peak type double-stage Hall thruster configuration, and a particle-in-cell simulation is carried out to investigate the influences of both acceleration electrode voltage value and zero magnetic point position on the thruster discharge characteristics and performances.The results indicate that increasing the acceleration voltage leads to a larger potential drop in the acceleration stage, allowing ions to gain higher energy, while electrons are easily absorbed by the intermediate electrode, resulting in a decrease in the anode current and ionization rate. When the acceleration voltage reaches 500 V, the thrust and efficiency are maximized, resulting in a 15%increase in efficiency. After the acceleration voltage exceeds 500 V, a potential barrier forms within the channel, leading to a decrease in thruster efficiency. Further study shows that as the second zero magnetic point moves towards the outlet of the channel, more electrons easily traverse the zero magnetic field region, participating in the ionization. The increase in the ionization rate leads to a gradual enhancement in both thrust and efficiency.

Influence of the position relationship between the cathode and magnetic separatrix on the discharge process of a Hall thruster

Previous studies have shown that there is an obvious coupling relationship between the installation location of the external cathode and the magnetic separatrix in the plume region of a Hall thruster.In this paper,the particle-in-cell simulation method is used to compare the thruster discharge process under the conditions of different position relationships between the cathode and the magnetic separatrix.By comparing the distribution of electron conduction,potential,plasma density and other microscopic parameters,we try to explain the formation mechanism of the discharge difference.The simulation results show that the cathode inside and outside the magnetic separatrix has a significant effect on the distribution of potential and plasma density.When the cathode is located on the outer side of the magnetic separatrix,the potential above the plume region is relatively low,and there is a strong potential gradient above the plume region.This potential gradient is more conducive to the radial diffusion of ions above the plume,which is the main reason for the strong divergence of the plume.The distribution of ion density is also consistent with the distribution of potential.When the cathode is located on the outer side of the magnetic separatrix,the radial diffusion of ions in the plume region is enhanced.Meanwhile,by comparing the results of electron conduction,it is found that the traiectories of electrons emitted from the cathode are significantly different between the inner and outer sides of the magnetic separatrix.This is mainly because the electrons are affected by the magnetic mirror effect of the magnetic tip,which makes it difficult for the electrons to move across the magnetic separatrix.This is the main reason for the difference in potential distribution.In this paper,the simulation results of macroscopic parameters under several conditions are also compared,and they are consistent with the experimental results.The cathode is located on the inner side of the magnetic separatrix,which can effectively reduce the plume divergence angle and improve the thrust.In this paper,the cathode moves from R=50 mm to R=35 mm along the radial direction,the thrust increases by 3.6 mN and the plume divergence angle decreases by 23.77%.Combined with the comparison of the ionization region and the peak ion density,it is found that the main reason for the change in thrust is the change in the radial diffusion of ions in the plume region.

Experimental investigation of the polarityswitching process with different bipolar ionic liquid thruster operating frequencies

The bipolar ionic liquid thruster employs ionic liquid as a propellant to discharge positively and negatively charged high-energy particles under an alternating current(AC)power source,effectively suppressing electrochemical reaction and ensuring charge neutrality.Determining an optimal AC supply power source frequency is critical for sustained stable thruster operation.This study focuses on the emission characteristics of the ionic liquid thruster under varied AC conditions.The AC power supply was set within the frequency range of 0.5-64 Hz,with eight specific frequency conditions selected for experimentation.The experimental results indicate that the thruster operates steadily within a voltage range of±1470 to±1920 V,with corresponding positive polarity current ranging from 0.41 to 4.91μA and negative polarity current ranging from−0.49 to−4.10μA.During voltage polarity switching,an emission delay occurs,manifested as a prominent peak signal caused by circuit capacitance characteristics and a minor peak signal resulting from liquid droplets.Extended emission test was conducted at 16 Hz,demonstrating approximately 1 h and 50 min of consistent emission before intermittent discharge.These findings underscore the favorable impact of AC conditions within the 8-16 Hz range on the self-neutralization capability of the ionic liquid thruster.

Power transfer efficiency in an air-breathing radio frequency ion thruster

Due to a series of challenges such as low-orbit maintenance of satellites, the air-breathing electric propulsion has got widespread attention. Commonly, the radio frequency ion thruster is favored by low-orbit missions due to its high specific impulse and efficiency. In this paper, the power transfer efficiency of the radio frequency ion thruster with different gas compositions is studied experimentally, which is obtained by measuring the radio frequency power and current of the antenna coil with and without discharge operation. The results show that increasing the turns of antenna coils can effectively improve the radio frequency power transfer efficiency, which is due to the improvement of Q factor. In pure N_2 discharge,with the increase of radio frequency power, the radio frequency power transfer efficiency first rises rapidly and then exhibits a less steep increasing trend. The radio frequency power transfer efficiency increases with the increase of gas pressure at relatively high power, while declines rapidly at relatively low power. In N_(2)/O_(2) discharge, increasing the N_(2) content at high power can improve the radio frequency power transfer efficiency, but the opposite was observed at low power. In order to give a better understanding of these trends, an analytic solution in limit cases is utilized, and a Langmuir probe was employed to measure the electron density. It is found that the evolution of radio frequency power transfer efficiency can be well explained by the variation of plasma resistance, which is related to the electron density and the effective electron collision frequency.

Performance of pulsed plasma thruster at low discharge energy

As the size of satellites scales down, low-power and compact propulsion systems such as the pulsed plasma thruster(PPT) are needed for stabilizing these miniature satellites in orbit. Most PPT systems are operated at 2 J or more of discharge energy. In this work, the performance of a PPT with a side-fed, tongue-flared electrode configuration operated within a lower discharge energy range of 0.5-2.5 J has been investigated. Ablation and charring of the polytetrafluoroethylene propellant surface were analyzed through field-effect scanning electron microscopy imaging and energy-dispersive X-ray spectroscopy. When the discharge energy fell below 2 J, inconsistencies occurred in the specific impulse and the thrust efficiency due to the measurement of the low mass bit. At energy ≥2 J, the performance parameters are compared with other PPT systems of similar configuration and discussed in depth.

Study on Characteristics of a High-Precision Cold Gas Micro Thruster

In order to improve the reliability of the spacecraft micro cold gas propulsion system and realize the precise control of the spacecraft attitude and orbit, a micro-thrust, high-precision cold gas thruster is carried out, at the same time due to the design requirements of the spacecraft, this micro-thrust should be continuous working more than 60 minutes, the traditional solenoid valve used for the thrusts can’t complete the mission, so a long-life micro latching valve is developed as the control valve for this micro thruster, because the micro latching valve can keep its position when it cuts off the outage. Firstly, the authors introduced the design scheme and idea of the thruster. Secondly, the performance of the latching valve and the flow characteristics of the nozzle were simulated. Finally, from the experimental results and compared with the numerical study, it shows that the long-life micro cold gas thruster developed in this paper meets the mission requirements.

Three-dimensional particle simulation of ion thruster plume flows with EX-PWS

Ion thruster plumes from a multi-thruster array of different working configurations are simulated by a hybrid fluid-particle software. The particle in cell method is employed to model the transports of ions. The direct simulation Monte Carlo method is used to model momentum and charge exchange （CEX） collisions. The software is based on unstructured grids which make it easy to handle with complex geometry. The results of chamber simulation are compared with experimental data in ion current density and number density, which show good agreements. The maximum difference of current density along the thruster centerline is less than 9.30%. The interaction effects of plumes when multiple thrusters are operating in vacuum are predicted. Distributions of single charged xenon ions are significantly different in the near-field plume flow, however, merge into one in the far downstream region. Moreover, the interaction effect on the spatial distribution of CEX xenon ions is displayed as well.

GEO Satellite Thruster Configuration and Optimization

A thruster configuration and optimization method for GEO satellite is proposed in this paper.All thrusters are installed on the back panel of satellite,which is not only compatible with the design of satellite subdivision structure,but also able to provide a large installation room for the payloads and components carried by the satellite.Chemical thruster and electric thruster were selected because they have complementary advantages and can effectively reduce propellant consumption.The working mode and tasks of the thruster were analyzed in detail for thruster configuration,which was optimized in order to minimize propellant consumption.Lastly,a GEO satellite thruster was configured and optimized using this method,with results showing that the method is feasible and effective,and consequently has value for engineering applications.

Electric Thrusters Redundancy Configuration Strategy Study for All Electric Propulsion Platform Station-Keeping

Electric propulsion is used for all electric propulsion satellites to perform the orbit transfer,attitude control and station-keeping tasks. Generally electric propulsion subsystem contains 4 thrusters. But if one thruster fails in the beginning of satellite lifetime,other thrusters will undertake all the firing tasks. The firing time will be 2 to 3 times of thrusters without failure. Thus it may go beyond the allow ed lifetime of thruster. This paper puts forward two thruster redundancy configuration solutions with 6 thrusters to solve this problem. Two layout configurations and their corresponding station-keeping strategies are simulated and compared. The results show that the maximum firing time of both layout configurations can meet the lifetime limitation. This solution is a good reference for all electric propulsion satellites design.

Effect of Pressure Level on the Performance of an Auto-Initiated Pulsed Plasma Thruster

Pulsed plasma thrusters （PPT） are micro-propulsion devices used in satellites for station keeping. Conventionally the plasma discharge in a PPT is initiated by a spark plug. The primary objective of the present work was to develop and characterize a PPT that does not need a spark plug to initiate the plasma discharge. If the spark plug is eliminated, the size of the thrusters can be reduced and arrays of such thrusters can be manufactured using micro electro mechanical systems （MEMS） techniques, which can provide tremendous control authority over the satellite positioning. A parallel rail thruster was built and its performances were characterized inside a vacuum chamber to elucidate the effect of vacuum level on the performance. The electrical performance of the thruster was quantified by measuring the voltage output from a Rogowski coil, and the thrust produced by the developed thruster was estimated by measuring the force exerted by the plume on a light weight pendulum, whose deflection was measured using a laser displacement sensor. It was observed that the thruster can operate without a spark plug. In general, the performance parameters such as thrust, mass ablation, impulse bit, and specific impulse per discharge, would increase with the increase in pressure up to an optimum level due to the increase in discharge energy as well as the decrease in the total impedance of the plasma discharge. The thrust efficiency is found to be affected by the discharge energy.

Particle Simulation of Three-Grid ECR Ion Thruster Optics and Erosion Prediction

Ion optics is a critical component of ion thrusters. A two-dimensional axisymmetric model is developed to study the characteristics of three-grid electron cyclotron resonance ion thruster optics. The code is based on a particle-in-cell combined with the Monte Carlo collision method to simulate ion dynamics and charge-exchange processes in the grid region. The simulation results show that the mode can give a reasonable estimate of the physics characteristics of the ion optics. The design of the ion optics satisfies the requirement of preventing electron backstreaming. Charge-exchange ions can cause damage to the grids, especially to the accelerator grid. ＇Barrel＇ erosion can increase the accelerator grid aperture radius at a rate of 1.91~ 10-11 m/s, while the decelerator grid plays an important role in reducing ＇pits-and-grooves＇ erosion.

Weak thruster fault detection for AUV based on stochastic resonance and wavelet reconstruction

When the bi-stable stochastic resonance method was applied to enhance weak thruster fault for autonomous underwater vehicle(AUV), the enhancement performance could not satisfy the detection requirement of weak thruster fault. As for this problem, a fault feature enhancement method based on mono-stable stochastic resonance was proposed. In the method, in order to improve the enhancement performance of weak thruster fault feature, the conventional bi-stable potential function was changed to mono-stable potential function which was more suitable for aperiodic signals. Furthermore, when particle swarm optimization was adopted to adjust the parameters of mono-stable stochastic resonance system, the global convergent time would be long. An improved particle swarm optimization method was developed by changing the linear inertial weighted function as nonlinear function with cosine function, so as to reduce the global convergent time. In addition, when the conventional wavelet reconstruction method was adopted to detect the weak thruster fault, undetected fault or false alarm may occur. In order to successfully detect the weak thruster fault, a weak thruster detection method was proposed based on the integration of stochastic resonance and wavelet reconstruction. In the method, the optimal reconstruction scale was determined by comparing wavelet entropies corresponding to each decomposition scale. Finally, pool-experiments were performed on AUV with thruster fault. The effectiveness of the proposed mono-stable stochastic resonance method in enhancing fault feature and reducing the global convergent time was demonstrated in comparison with particle swarm optimization based bi-stochastic resonance method. Furthermore, the effectiveness of the proposed fault detection method was illustrated in comparison with the conventional wavelet reconstruction.

PIC simulation of plasma properties in the discharge channel of a pulsed plasma thruster with flared electrodes

Plasma in the discharge channel of a pulsed plasma thruster(PPT) with flared electrodes is simulated by a self-developed two-dimensional code. The fully particle-in-cell method with Monte Carlo collision is employed to model the particle movement and collisions and investigate the plasma properties and acceleration process. Temporal and spatial variations of the electron density distribution and the ion velocity between electrodes are calculated and analyzed in detail.The computational results of the electron number density, which is in the order of 1023 m-3,show good agreements with experimental results of a PPT named ADD SIMP-LEX. The ion velocity distributions along the center line of the channel lead to a comprehensive understanding of ions accelerated by electromagnetic field. The electron distributions of PPT with discharge voltages varying from 1300 to 2000 V are compared. The diffusion of electrons presents strong dependency on discharge voltage and implies higher degree of ionization for higher voltage.

Study of the key factors affecting the triple grid lifetime of the LIPS-300 ion thruster

In order to ascertain the key factors affecting the lifetime of the triple grids in the LIPS-300 ion thruster,the thermal deformation,upstream ion density and component lifetime of the grids are simulated with finite element analysis,fluid simulation and charged-particle tracing simulation methods on the basis of a 1500 h short lifetime test.The key factor affecting the lifetime of the triple grids in the LIPS-300 ion thruster is obtained and analyzed through the test results.The results show that ion sputtering erosion of the grids in 5 kW operation mode is greater than in the case of 3 kW.In 5 kW mode,the decelerator grid shows the most serious corrosion,the accelerator grid shows moderate corrosion,and the screen grid shows the least amount of corrosion.With the serious corrosion of the grids in 5 kW operation mode,the intercept current of the acceleration and deceleration grids increases substantially.Meanwhile,the cold gap between the accelerator grid and the screen grid decreases from 1 mm to 0.7 mm,while the cold gap between the accelerator grid and the decelerator grid increases from 1 mm to 1.25 mm after 1500 h of thruster operation.At equilibrium temperature with 5 k W power,the finite element method（FEM）simulation results show that the hot gap between the screen grid and the accelerator grid reduces to 0.2 mm.Accordingly,the hot gap between the accelerator grid and the decelerator grid increases to 1.5 mm.According to the fluid method,the plasma density simulated in most regions of the discharge chamber is 1?×?10;-8?×?10;m;.The upstream plasma density of the screen grid is in the range 6?×?10;-6?×?10;m;and displays a parabolic characteristic.The charged particle tracing simulation method results show that the ion beam current without the thermal deformation of triple grids has optimal perveance status.The ion sputtering rates of the accelerator grid hole and the decelerator hole are 5.5?×?10;kg s;and 4.28?×?10;kg s;,respectively,while after the thermal deformation of the triple grids,the ion beam current has over-perveance status.The ion sputtering rates of the accelerator grid hole and the decelerator hole are 1.41?×?10;kg s;and 4.1?×?10;kg s;,respectively.The anode current is a key factor for the triple grid lifetime in situations where the structural strength of the grids does not change with temperature variation.The average sputtering rates of the accelerator grid and the decelerator grid,which were measured during the 1500 h lifetime test in5 k W operating conditions,are 2.2?×?10;kg s;and 7.3?×?10;kg s;,respectively.These results are in accordance with the simulation,and the error comes mainly from the calculation distribution of the upstream plasma density of the grids.

Catalytic bed slenderness ratio and ADN/methanol ratio for decomposition and combustion characteristics within ammonium dinitramide(ADN)-based green aerospace thruster

The decomposition and combustion characteristics of ammonium dinitramide (ADN) based non-toxic aerospace propellant are analytically studied to determine the effects of catalytic bed structure (slenderness ratio) and operation parameters (mass fraction ratio of ADN/CH3OH) on the general performance within the ADN-based thruster. In the present research, the non-equilibrium temperature model is utilized to describe the heat transfer characteristics between the fluid phase and solid phase in the fixed bed. We determined the fluid resistance characteristics in the catalytic bed by experiments involving the method of pressure-mass. We have done the simulation study based on the available results in the literature and found the complex physical and chemical processes within the ADN thruster. Furthermore, an optimized catalytic bed slenderness ratio was observed w让h a value of 1.75 and the mass fraction ratio of 5.73 significantly influenced the propellant performance. These results could serve as a reference to explore the combustion characteristics within the thruster and the preparation of future propellants.

Numerical research of a 2D axial symmetry hybrid model for the radio-frequency ion thruster

Since the high efficiency discharge is critical to the radio-frequency ion thruster(RIT), a 2D axial symmetry hybrid model has been developed to study the plasma evolution of RIT. The fluid method and the drift energy correction of the electron energy distribution function(EEDF) are applied to the analysis of the RIT discharge. In the meantime, the PIC-MCC method is used to investigate the ion beam current extraction character for the plasma plume region. The beam current simulation results, with the hybrid model, agree well with the experimental results, and the error is lower than 11%, which shows the validity of the model. The further study shows there is an optimal ratio for the radio-frequency(RF) power and the beam current extraction power under the fixed RIT configuration. And the beam extraction efficiency will decrease when the discharge efficiency beyond a certain threshold(about 87 W). As the input parameters of the hybrid model are all the design values, it can be directly used to the optimum design for other kinds of RITs and radio-frequency ion sources.

Application and development of the pulsed plasma thruster

The application and development of pulsed plasma thrusters(PPTs)in recent years are reviewed in this paper.The advantages of PPTs are discussed.The schematics,propulsion performance parameters and key physical processes of PPTs are described.Some representative PPT products and flight systems developed in recent years are presented to show the performance of the PPT.Studies about how electrode structures,discharge circuits,propellant materials,energy discharge method,propellant feed method,ignition method and number of thruster heads influence the PPT performance are presented and analyzed.The ignitor design method,ignition process and propellant carbonization are introduced to discuss the reliability and lifetime issues in PPTs.The modeling methods of the discharge circuit,as well as ablation,ionization and acceleration in PPTs are presented.Finally,the application of PPTs in the future is analyzed and some suggestions for PPT development are proposed.

Nonlinear Dynamic Characteristics of the Vectored Thruster AUV in Complex Sea Conditions

The mobility of the vectored thruster AUV in different environment is the important premise of control system design. The new type of autonomous underwater vehicle (AUV) equipped with rudders and vectored thrusters which are combined to control the course is studied. Firstly, Euler angles representation and quaternion method are applied to establish six-DOF kinematic model respectively, then Newton second law and Lagrangian approach are used to deduce the vectored thruster AUV’s nonlinear dynamic equations with six degrees of freedom (DOF) respectively in complex sea conditions based on the random wave theory according to the structural and kinetic characteristics of the vectored thruster AUV in this paper. The kinematic models and dynamic models based on different theories have the same expression and conclusion, which shows that the kinematic models and dynamic models of the vectored thruster AUV are accurate. The Runge-Kutta arithmetic is used to solve the dynamic equations, which not only can simulate the motions such as cruise and hover but also can describe the vehicle’s low-frequency and high-frequency motion. The results of computation show that the mobility of the vectored thruster AUV in interference-free environment and the integrated signals including low-frequency motion signal and high-frequency motion signal in environmental disturbance accord with practical situation, which not only solve the problem of especial singularities when the pitch angle θ = ±90° but also clears up the difficulties of computation and display of the coupled nonlinear motion equations in complex sea conditions. Moreover, the high maneuverability of the vectored thruster AUV equipped with rudders and vectored thrusters is validated, which lays a foundation for the control system design.