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.

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.

Orificed Hollow Cathodes are electric devices necessary for the functioning of common plasma thrusters for space applications. From their reliability mainly depends the success of a spacecraft’s mission equipped by 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 system of plasma equations, have been proposed in the open literature. These are solved commonly by means of iterative procedures. This paper investigates the possibility to solve them by means of 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 function of mass flow rate is very narrow (2÷5V), moreover the results match very well the experimental data except at the lowest mass flow rate (0.08mg/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 (± 50K respect to the average values) of orifice and insert; those of the others variables are very tiny.

Gridded Ion Thruster (GIT) is a versatile space propulsion device that can be used for a wide range of space missions being characterized by a high trust-to-power ratio and a relatively high specific impulse. In recent years, there has been a significant increase in academic and scientific interest in interplanetary and deep space exploration that has shifted research resources toward boosting the performances of GITs. One promising concept is to separate the extraction process of the ions from their acceleration (Dual Stage). Some researches focused the attention on basic experiments to assess the functioning principle, some tried to simply model its Optical System (OS) and only few of them used a numerical approach to verify it. In this work a global plasma model for both the Direct Current (DC) ionization chamber and OS is proposed. The model combines analytical formulas, describing the plasma physics into ionization chamber, a linear optic model for OS, and a curve-fitting approach, to get performances and working parameters. As input it requires the diameter of the engine and the thrust or power. An educated guess is made by considering the validity of the thrust/diameter-total efficiency curve, obtained for conventional GIT, due to the lack of these data for Dual Stage Gridded Ion Thrusters (DSGIT). Despite this hypothesis, the results obtained by the model are in good agreement with the performances and working parameters of preliminary designed high power (25-50 kW) thrusters, available in open literature and selected for the validation, obtaining for the most differences less than 5 %. Finally the model has been applied to exploit the advantage of DS approach by designing a 1kW thruster: the results are very promising because a ten times smaller thruster than a conventional GIT, could have hundred times higher thrust and power density and an increment of the efficiency and the specific impulse by 25 %.

Magnetoplasmadynamic thrusters represent a promising Electric Propulsion technology for future space missions; however, their optimization is hampered by the lack of accurate performance models in the emerging regime of low power (0.1 T), where traditional formulations prove inadequate. In this work, a new semi-empirical model for predicting the thrust and discharge voltage of argon-fed MPD thrusters was developed and validated. Starting from state-of-the-art physical models, multi-factorial correction factors were introduced to account for the coupled effects of discharge current (8–180 A), mass flow rate (3–21 mg/s), and applied magnetic field (up to 0.6 T). The model was calibrated and validated using a comprehensive and homogeneous collection of experimental data from the literature. A comparative analysis demonstrates that the corrected model significantly reduces prediction errors (0–9%) compared to reference models available in the literature (8–50%). In particular, the model exhibits remarkably superior accuracy in both the Self-Field and Applied-Field regimes, overcoming the main limitations of previous formulations and providing more robust estimates across a wide operational envelope. The developed model constitutes a reliable and physically consistent tool for the analysis and preliminary design of low-power, argon-fed magnetoplasmadynamic thrusters, enabling more effective optimization for this class of propulsion systems.