Hall Effect Thrusters (HETs) are nowadays widely used for satellite applications because of their efficiency and robustness compared to other electric propulsion devices. Computational modelling of plasma in HETs is interesting for several reasons: it can be used to predict thrusters’ operative life; moreover, it provides a better understanding of the physical behaviour of this device and can be used to optimize the next generation of thrusters. In this work, the discharge within the accelerating channel and near-plume of HETs has been modelled by means of an axisymmetric hybrid approach: a set of fluid equations for electrons has been solved to get electron temperatures, plasma potential and the discharge current, whereas a Particle-In-Cell (PIC) sub-model has been developed to capture the behaviour of neutrals and ions. A two-region electron mobility model has been incorporated. It includes electron–neutral/ion collisions and uses empirical constants, that vary as a continuous function of axial coordinates, to take into account electron–wall collisions and Bohm diffusion/SEE effects. An SPT-100 thruster has been selected for the verification of the model because of the availability of reliable numerical and experimental data. The results of the presented simulations show that the code is able to describe plasma discharge reproducing, with consistency, the physics within the accelerating channel of HETs. A small discrepancy in the experimental magnitude of ions’ expansion, due probably to boundary condition effects, has been found.

Radio-frequency and Helicon Plasma Thrusters have emerged as viable electric propulsion systems due to their high plasma density, thrust density, and useful life. Helicon Plasma Thruster (HPT) is a very attractive technology because it could use many propellants and does not require hollow cathodes or grids, overcoming their associated critical erosion problem and extending the thruster’s lifetime to some tens of thousands of hours. Despite the fact that high-power HPTs have reached 30% efficiency in laboratory configurations, sophisticated numerical models are required for a deeper understanding of the main plasma phenomena and for the preliminary design to increase the very low HPT’s efficiency (3–7%) typical of the low-power class thrusters. The paper focuses on the development of a model for the low-medium power range (50–2000 W) of HPTs design. Starting from Lafleur’s model, it has been improved with the hypothesis of neutral gas being expelled at the real thruster’s wall operative temperature (300–600 K) in place of the more frequent laboratory temperature assumption (300 K). This hypothesis affects total thrust and specific impulse by about 10%. A parametric analysis of the slenderness ratio (chamber length-to-radius) has been conducted. The results showed that slender configurations lead to higher efficiencies. Downstream from the numerical model validation, a tool for the global design has been built with the Particle Swarm Optimization (PSO) technique that leads to optimal thruster configuration. This tool has been used to design a 4 mN HPT tuning the PSO in order to minimize the dimensions and the weight according to the assigned mission constraints (i.e., power, thrust, and weight). A total efficiency of 10.4% results.

The erosion of the accelerating chamber walls is one of the main factors limiting the operational life of Hall effect thrusters (HETs), and it is mainly related to the sputtering of ceramic walls due to the impacting energetic ion particles. The erosion phenomenon is investigated by means of a numerical model that couples the plasma model HYPICFLU2, used for evaluating the local distributions of ion energies and incidence angles, and a sputtering model specific for the xenon–Borosil pair, which is the most used in HETs application. The sputtering yield model is based on the measurements by Ranjan et al. that are improved with a linear factor to include wall temperature effect, recently studied by Parida et al. The experimental eroded profiles of SPT100 walls are selected as benchmark. The results show that there is a decrease in erosion speed with time, in accordance with experimental measurements, but the model underestimates, by about 50–60%, the erosion at the channel exit, which suggests a stronger dependence of sputter yield on surface temperature. Thus, the need for new experimental measurements of sputtering in the range of impact energy, angle, and wall temperature, respectively, of 10–250 eV, 0–85°, 30–600 °C, arises.

Gridded ion thrusters (GITs) are an established technology that, by covering a wide range of power class, allows one to accomplish a lot of space mission types. Many analysis tools and analytical models describing the physics of GITs are present in the open literature, while there is a lack of tools for preliminary design, considering the mission requirements (i.e., thrust or power). Thus, in this work, a tool that takes as input thrust or power and that combines analytical formulas, describing GITs’ physics; a curve-fitting approach, exploiting data from different ion thrusters present in the open literature; and an FEMM (finite element method magnetics) simulation has been developed and validated against known medium-low-power (<5 kW) gridded ion thrusters (e.g., NSTAR, XIPS, ETS-8). Some of the main outputs of the developed tool are its specific impulse, efficiencies, voltages, and propellant flow rate. The results obtained by the tool have been in good agreement with the real performance and working parameters of the thrusters selected for the validation, obtaining an average error of less than 5–10%. The tool has been also compared with a tool proposed in the literature as a possible design tool, which makes use of a simple macroscopic plasma-source simulation (SMPS) code with a genetic algorithm (GA) and obtains slightly more accurate results on average. Finally, the tool has been exploited for the design of a very low-power GIT (100 W) that is able to produce 2 mN of thrust, as the interest of the scientific community in miniaturizing electric engines has recently grown because they could enable new space missions.

Orifice Hollow Cathodes are electric devices necessary for the functioning of common plasma thrusters for space applications. Their reliability mainly depends on the success of a spacecraft’s mission equipped with electric propulsion. The development of plasma models is crucial in the evaluation of plasma properties within the cathodes that are difficult to measure due to the small dimensions. Many models, based on non-linear systems of plasma equations, have been proposed in the openiterature. These are solved commonly by means of iterative procedures. This paper investigates the possibility of solving them by means of the Particle Swarm Optimization method. The results of the validation tests confirm the expected trends for all the unknowns; the confidence bound of the discharge current as a function of mass flow rate is very narrow (2 ÷ 5) V); moreover, the results match very well the experimental data except at theowest mass flow rate (0.08 mg/s) and discharge current (1A), where the computations underpredict the discharge current to the utmost by 40%. The highest data dispersion regards the plasma density in the emitter region (±20% of the average value) and the wall temperatures (±50 K with respect to the average values) of the orifice and insert; those of the others variables are very tiny.