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Effects from ion thruster plume have long raised concerns. However, little work has systematically analyzed the influencing parameters in the ion thruster plume simulation. This paper analyzes LIPS 200 ion thruster plume simulations about the influencing parameters. The numerical simulations are carried out by a hybrid particle-in-cell (PIC) method and the direct simulation Monte Carlo (DSMC) method. The PIC method is employed for the plasma dynamics, and the DSMC method is used for collisions. Simulation results were compared in detail to obtain the variation of the results with the parameters. Besides, experimental data were compared to simulation results to optimize the parameters. Finally, with these researches, the optimal parameters are obtained.

The low-pressure gas in the vacuum plume produced by the chemical thrusters contaminates the spacecraft when adsorbed on the low-temperature surface. To provide theoretical support for further research on gaseous plume pollutants, the adsorption isotherms of low-pressure H2O were measured by a quartz crystal microbalance (QCM) at temperatures ranging from 233 to 273 K. The measured isotherms are similar to the type-I and type-II isotherms and have been correlated by various models (e.g., the Langmuir, Dubinin-Radushkevich, Brunauer-Emmett-Teller (BET), and universal models). It shows that the universal model has a great advantage in predicting the adsorption at a specific temperature point in our study. To estimate the adsorption at the continuous temperature range, the critical parameters of the multi-Langmuir model were expressed in semiempirical formulas. Since the normalized isotherms of H2O at different temperatures converge well, a simplified multi-Langmuir (SML) model was proposed. The experimental results at the temperature and pressure ranges we explored are consistent with the results predicted by the SML model, suggesting that the SML model is more suitable and convenient to predict the low-pressure adsorption of H2O for a continuous low-temperature range. Moreover, the low-pressure adsorption behaviors of H2O and CO2 on the low-temperature surface are compared and discussed.

The convective heat transfer of the plume of a 120-N thruster is investigated experimentally and numerically. Numerical results agree well with experimental results in that there is a nonlinear decrease in convective heat transfer with an increasing cone angle. It is also found that convective heat transfer decreases with increasing distance from the thruster outlet. Furthermore, the convective heat transfer of the plume mainly concentrates within a 35° cone angle and the heat flux decreases to the same order as solar radiation at the Earth’s surface when the cone angle exceeds 60°. The results of the study will help improve spacecraft design.

The interaction between the high-energy particles in the plume and the spacecraft surfaces will produce interference torque that affects the operating state of the spacecraft in orbit. The thrust of an electric propulsion system is quite small, so it is difficult to be measured directly. The Vacuum Plume Laboratory (VPL) measured the LIPS-200 type ion thruster plume force in an order of 10 − 3 N using a fully elastic micro thrust measuring device. In this paper, the particle in cell (PIC) method and the direct simulation Monte Carlo (DSMC) method are employed to analyze the three-dimensional plasma environments under specified experimental conditions. The Maxwell model is used to calculate the plume force on a 300 mm diameter plate. Simulation results of the plume force give good agreements with the experimental data. Moreover, the effects of the 300 mm diameter aluminium plate on the flow field in vacuum conditions are analyzed. The results show that the number density of atoms is greatly increased before the plate, which has a further impact on the distribution of charge exchange ions (CEX). The enhanced CEX ions moving towards the solar battery panels or sensitive optical components may cause possible damage or interference, which should be avoided by the designers.

Ionic liquid electrospray thrusters with high specific impulse, high thrust accuracy and low thrust noise are very promising for space gravitational wave detection missions. The plume, which may lead to surface charging of solar panels and sensitive spacecraft components is a great concern for the applications of electrospray thruster. Therefore, this paper investigates the plume neutralization process of the ionic liquid electrospray thruster through fully kinetic particle-in-cell simulations. The unipolarity operation mode is firstly simulated and compared with the experimental measurements. The bipolar operation mode is analyzed by considering the premixing and the separation of positive and negative ion beams. At the same time, the effect of beam spacing on the plume characteristics is investigated. The results show that the plume neutralization of the ionic liquid electrospray thruster is achieved by the spatial and temporal oscillations of the ion beams. In the horizontal direction, the spatial oscillations are caused by the different mass and hence velocities of positive and negative ions. In the vertical direction, the spatial oscillations are mainly because of the non-zero beam spacing. The temporal oscillations may be related to the tradition plasma oscillation. As the beam spacing increases, the oscillation amplitude of the horizontal electric potential curve changes scarcely; however, in the vertical direction, the oscillation amplitude of the curve increases. The ion temperature goes up with the beam spacing and the deviation of the temperature of beam ions does not exceed 15 eV in the horizontal direction but exceeds 100 eV in the vertical direction. Moreover, the plume divergence half angle and the beam spacing are positively correlated, suggesting that the ionic electrospray thrusters with positive and negative polarity need to be placed as close as possible in the spacecraft.

A coupled NS-DSMC method possessing adapted-interface and two-way coupling features is studied to simulate the plume impingement effects of space thrusters. The continuum-rarefied interface is determined by combining KnGL and Ptne continuum breakdown parameters. State-based coupling scheme is adopted to transfer information between continuum and particle solvers, and an overlapping grid technique is investigated to combine structured-grid NS code and Cartesian-grid DSMC code to form the coupled solver. Flow problem of a conical thruster plume impinging on a cone surface is simulated using the coupled solver, and the simulation result is compared with experimental data, which proves the validity of the proposed method. Plume flow while the ascent stage of lunar module lifting off in lunar environment is also computed by using the present coupled NS-DSMC method to demonstrate its capability. The whole flow field from combustion chamber to the vacuum environment is obtained, and the result reveals that special attention should be paid to the plume aerodynamic force at the early stage of launching process.

Pressure-sensitive paint (PSP) technique was employed to experimentally investigate the aerodynamic force effect of vacuum plume in this study. The characterization and comparison for two types of PSP were firstly conducted in an air pressure range from 0.05 to 5000 Pa. The PSPs were prepared using PtTFPP as the active dye and different binders, i.e., polymer-ceramic (PC) and poly(1-trimethylsilyl-1-propyne) [poly(TMSP)]. The static calibrations showed that PtTFPP/poly(TMSP) had a higher pressure sensitivity and a lower temperature dependency compared to PtTFPP/PC in this pressure range. The pressure distributions of a single and two interacting plumes impinging onto a flat plate model were measured using PSP technique. The experimental data were compared to numerical solutions that combined both the computed fluid dynamics (CFD) and direct simulation Monte Carlo (DSMC) methods. Remarkable agreements were achieved, demonstrating the feasibility and accuracy of the numerical approach. Finally, the aerodynamic force effect of interacting plumes at different separation distances was investigated numerically.

In spacecrafts such as high-altitude rockets and lunar landers, multiple thrusters or engines operate simultaneously with a small separation distance. Thus, the plume interaction may cause a different three-dimensional flow field with a single plume or plumes separated by a large distance. To study the flow field of the plume interaction, a hydrogen/oxygen sample thruster with a thrust of 60 N with the same propellant, mixing ratio, and shrunken bell-nozzle profile as an actual engine was first designed and tested in the Vacuum Plume Effects Experimental System (PES). The pressure field of the plume was determined using a pitot-tube array. The experimental data were compared with the results from a numerical solution that combined both computational fluid dynamics (CFD) and direct simulation Monte Carlo (DSMC) analyses, and a favorable agreement was obtained, demonstrating the application ability of the CFD-DSMC method to predict the plume flow field. Subsequently, two identical sample thrusters were modeled with a separation distance ranging from 1.0 to 3.0 times the diameter of the nozzle exit and a deflection angle ranging from 0° to 20°, and the plume interactions were numerically investigated.

This paper introduces a newly developed vacuum Plume effects Experimental System(PES) used for plume effect tests of rocket engines and vacuum heat tests of satellites. The design level, manufacturing technique, and testing capabilities of the PES have reached a highly advanced level at home and abroad. The PES mainly consists of a vacuum chamber, vacuum acquisition system, nitrogen system, helium system, and parameter measurement system. A breakthrough was obtained on the Large Scale Cryo-Pumping System, which was based on a combined liquid nitrogen and liquid helium heat sink. An internal cryopump with a limiting temperature of 4.2 K and an efficient absorption area of 305 m 2 was developed. The absorption capability of the cryopump was above 7×10 7 L/s. Vacuum plume tests were performed in the temperature ranges of ambient temperature, liquid nitrogen, and liquid helium. The experimental results showed that the plume test capability of PES is higher than that of similar foreign equipment STG and CHAFF-4. For 2 g/s and 117 N rocket engines, the dynamic vacuum degree of environment was 8.0×10 -4 Pa(approximately 137 km height) and 1.1×10 -2 Pa(approximately 106 km height), respectively.

When the plume of electric thrusters, which is the plasma flow, collides with a surface of spacecraft, a force is exerted on the surface, causing the spacecraft to be flipped or deorbited. The force is generated by the plasma–surface interactions, which can be described using the momentum accommodation coefficient (MAC). As far as we know, only the MAC where the plasma flow is assumed as the combination of singly charged ions, electrons, and neutral gases has been studied. However, the plasma flow of electric thrusters contains not only singly charged ions but also multiply charged ions. In addition, the plasma–surface interactions are usually accompanied by the charged-exchange (CEX), sputtering, and plasma sheath, all of which influence the momentum transfer. Therefore, in this study, the MAC in plasma–surface interactions is developed by quantifying the effect of the multiply charged ions, CEX, sputtering, and plasma sheath. The formulas of the normal momentum accommodation coefficient (NMAC) and tangential momentum accommodation coefficient are derived, respectively, and then, a measurement method of the MAC is proposed with the derived formulas. Moreover, a NMAC was measured when the plasma flow, which was jetted by an ion thruster with an acceleration voltage of 1000 V, impacted an immersed surface vertically, and the measured NMAC is approximated as 1.0. The effect degree and mechanism of the aforementioned factors on the MAC were discussed. Finally, the force exerted on the immersed surface was simulated using the measured NMAC, and the results are in good agreement with the experimental results.

Attitude-control engine plume will generate a low pressure area on spacecraft surface. The low pressure gas adsorbed on cryogenic surface could cause contamination effect and affect the performance of spacecraft. Low pressure adsorption theory at low surface temperature is conducive to analyze contamination effect. Therefore, it is necessary to investigate the low pressure adsorption theory. In this paper, an experimental system was designed for low gas pressure and low surface temperature control. Low-pressure carbon dioxide adsorption equilibrium isotherms less than 100Pa on a gold-plated quartz crystal microbalance were experimentally measured at nine surface temperatures from 213 K to 273 K. The experimental results showed that each isotherm was consistent with type I isotherm. However, the maximum adsorption capacities of the isotherms under different temperatures were different and decreased exponentially with the increase of surface temperature. This phenomenon indicates that multilayer adsorption occurred in these experimental conditions and the traditional Langmuir model was not suitable for the isotherms in this experiment. Therefore, a Multi-Langmuir isotherm model was proposed based on the isotherms to describe carbon dioxide adsorption isotherm under low pressure and low surface temperature. The error between the Multi-Langmuir model and the experimental results is basically less than 20%.

In vacuum environment, the exhaust flow of attitude control thrusters would expand freely and produce the plume, which possibly causes undesirable contamination, aerodynamic force and heating effects to the spacecraft. Plume work station (PWS) is developed by Beihang University (BUAA) for numerically simulating the vacuum plume and its effects. An approach which combines the direct simulation Monte Carlo (DSMC) method and difference solution of Navier–Stokes (N–S) equations is applied. The internal flows in nozzles are simulated by solving the NS equations. The flow parameters at nozzle exit are used as the inlet boundary condition for the DSMC calculation. Experimental studies are carried out in a supersonic low density wind tunnel which could simulate the 60–80 km altitude environment to investigate the plume and its effects. To demonstrate the capability of PWS, numerical simulations are performed for the vacuum plume of several typical attitude control thrusters. The research results are of great help for the engineering design.

When lunar modules land on the Moon, dust impingement on the lunar module components and deposition on the thermal and optical surfaces would cause many serious problems. The emphasis of this research is on simulating the interaction of rocket plume and lunar dust using the direct simulation Monte Carlo (DSMC) method. The method is extended to model the movement and collision stages of rarefied plume gas and dust particles, including three collisional mechanisms: molecule–molecule, molecule–particle and particle–particle collisions. The reflection of gas molecules on the particle surface is computed by an indirect approach based on the coordinate transformation. Neighboring-cell contact detection scheme is applied to check for contacts between all possible particle pairs. The simulation results show that the acceleration of dust particle is mostly determined by the parameters of plume field. In the computational regions with larger gas density and velocity the particles can be accelerated to larger velocity and convected to higher temperature.

A momentum flux measuring instrument (MFMI), which ranges from 1 mN∼300 N, is designed to measure the parameters related to the momentum flux of exhaust plume of attitude and orbit control thrusters covered electric and chemical thrusters. A MFMI with the variable-range provides an efficient and economy way to study the exhaust plume of both electric and chemical thrusters. The flexural pivots, replaceable strain gauge, replaceable target plate, and force arm with the variable length make it possible. The designed MFMI with the non-displacement measurement can reduce the influence of pipelines and test lines, especially the elastic force and friction due to displacement. The designed MFMI system with the range of 0∼100 mN and 0∼15 N was calibrated separately. The target indirect measurement method is a simple and economical option for measuring the thrust of electric thrusters, and is employed to measure the thrust of the LIPS-200 ion thruster using the designed MFMI. Research indicates that the sputtering contributes to the momentum transfer of plasma-surface interactions. In addition, the high-speed charged ions from the thruster will experience a charge-exchange (CEX) collision with the slow neutral background. Therefore, the influence of the sputtering and CEX on target indirect measurement method is derived in detail, and then the measured thrust is corrected. The measured average thrust, which is 40.1 mN, demonstrates that the designed MFMI is effective. The designed MFMI can also be employed as a thrust stand by directly mounting the thrusters on torsional arm, and a maximum of three thrusters can be installed on the MFMI at the same time.

The plume flow of space thrusters can cause series of harmful effects to spacecraft by impinging on some sensitive surfaces. To prevent the plume impingement effects that a 60N thruster can do to the backflow region, a plume protect cover is intended to be assembled to the thruster. The flow field from the combustion chamber of the thruster to the vacuum environment contains both continuum and rarefied flow regimes, which makes it difficult to model this flow phenomenon by using one single method. The coupled NS/DSMC method is one of the best choices for modeling the continuum-rarefied transitional flows. In present study, a coupled NS/DSMC solver is developed by the authors and is used to estimate the effect of the plume protector assembled to the 60N thruster. The present coupled NS/DSMC solver is constructed with the NS solver NozzleFlow and DSMC solver PWS. Interface location is determined by combining two continuum breakdown parameters. One is the KnGL parameter and the other is the thermal non-equilibrium parameter Ptne. State-based coupling method is adopted to transfer information between the continuum and particles methods. The continuum-rarefied interfaces and gas parameters of the vacuum plume flow field of the thruster are obtained by conducting coupled calculation, which shows that the plume protector can decrease the mass flux density in the back flow region by two orders. The present coupled NS/DSMC method can be a useful tool for the vacuum plume effects analysis.

During the second period of China “Tanyue” Project, the explorer will softland on the moon. The cushion engines are used to decelerate the explorer and reduce the impact on the lunar ground. It is necessary to study its plume effects on the explorer component. The self-developed PWS (Plume WorkStation) software based on direct simulation Monte Carlo (DSMC) method is used to simulate the plume effects of two 150 N engines. Due to the complex structure of the explorer, PWS uses a decoupling method to treat the boundary mesh, which mainly interacts with simulation particles, and has no relation with the computational grids. After the analytical expressions of plane surfaces and curved surfaces of each boundary block are given, the particle position within or without the boundary blocks can be easily determined. Finally the 3D plume field of two 150 N engines is simulated. The pressure, temperature and velocity distributions of plume field are clearly presented by three characteristic slices. The aerodynamic effects on the explorer bottom, the landfall legs and antenna are separately shown. The compression influence on the plume flow of four landfall legs can be observed.