Currently, there is a need for dynamic space missions based on small satellites. These missions can be supported by propulsion systems with thrust-vectoring capabilities. This capability can be realized based on electrodeless plasma thrusters (EPTs). EPTs stand out for their versatility, offering adjustable thrust characteristics and fewer components, making them ideal for small satellites. However, their efficiency remains below optimal levels, largely due to complexities in plasma acceleration. This research aims to better understand dominant acceleration mechanisms in EPTs by studying ion energy distribution function changes based on exhaust orifice diameter and power variations. The total power supplied to the thruster varies in the range of 24 to 40 W, and the exhaust diameter varies in the range from 6.5 to 10.5 mm. It was found that the ion velocity does not change as a function of the diameter of the exit aperture. This indicates the insignificance of the mechanism of the gas-dynamic acceleration of plasma in EPTs with a small form factor and supports recent views that the main contribution to the acceleration of particles in EPT is made by electromagnetic effects. The findings could help refine EPT designs, enhancing their overall effectiveness and reliability for future space missions.

The efficient operations of orbital satellite constellations require propulsion systems capable of continuously performing maneuvers. These maneuvers include not only the orbit phasing, the altitude alterations and the attitude control, but also collision-avoidance and deorbiting maneuvers. These maneuvers require the generation of thrust in multiple directions. The thrust-vectoring capability can be achieved by changing the spacecraft orientation beforehand. However, this approach may lead to inefficient propellant management. Another approach to control the thrust vector direction is the use of propulsion systems with thrust-vectoring capability. One of such propulsion systems is the thruster with close ring-shaped gas discharge chamber providing capability to generate thrust in two and more directions. Such geometry of the gas discharge chamber requires the thorough studies. In this work, the starting modes of the thruster with close ring-shaped gas discharge chamber utilizing krypton and argon are studied. It has been established that the presence of a toroidal magnetic field decreases the ignition thresholds of RF discharge by 32.4%. It is found that the energy consumed for discharge ignition in multichannel gas discharge chambers can be based on their geometry. Under certain configurations of the external magnetic field, a decrease in the discharge ignition thresholds is observed due to magnetized electrons circulation in-and-out of the gas discharge chamber.

Electrodeless plasma thrusters with thrust vector control are proven technologies for increasing maneuvrability of small satellites. In this study, the thruster with a closed ring-shaped gas discharge is experimentally investigated. The thruster architecture allows for three-dimensional thrust vector control using toroidal magnetic coils and magnetic nozzles, which forms the basis of a multidirectional electrodeless plasma thruster. Electron temperature and density inside the discharge chamber are determined using optical emission spectroscopy. The experiments are conducted using argon as the propellant at flow rates ranging from 10 to 210 sccm and at a power of 40 W. The results show a clear decrease in electron temperature and an increase in electron density with increasing volume flow rate. The electron temperature ranged from 5.32 to 3.40 eV, and the electron density increased from 2.11∙1017 m-3 to 2.85∙1017 m-3 in the experiment, with flow rates varying from 50 to 210 sccm, respectively. Quantitatively, the average relative deviation between the electron density predicted by the theoretical model and that determined experimentally is 9.8 %. The experiments revealed that the thruster with a closed ring-shaped gas discharge chamber exhibits plasma densities at low powers, making it a promising technology for multidirectional propulsion systems for small satellites.

The desire to use space in the most rational and efficient way to address contemporary challenges leads to the necessity of creating multi-purpose space missions capable of solving a wide range of diverse tasks. This creates a demand for propulsion systems that can provide high maneuverability for modern and future spacecraft. One potential solution to increase the maneuverability of satellites is the use of electrodeless plasma thrusters with magnetic thrust-vectoring (MTVEPT). Their simple design and acceptable thrust-to-power characteristics can improve the cost-effectiveness of a space mission, increase its reliability and operational lifetime, and enable the required orbital maneuvers. This paper presents an experimental study on the ignition thresholds of a radiofrequency discharge in an electrodeless plasma thruster utilizing argon. The study is conducted over a gas flow rate range of 20 to 210 sccm, with solenoid currents from 0 to 5 A, for two magnetic field directions and two diameters of the exhaust orifice, which is varied using a diaphragm. It is found that a 93% relative reduction in the channel diameter leads to an average twofold decrease in the discharge ignition threshold, reaching a minimum value of 2.5 × 10³ V/m at a flow rate of 100 sccm. This can be used to reduce the thruster’s power consumption for the repetitive discharge ignitions when the propellant reserves are limited. Furthermore, four distinct discharge ignition regions are identified, depending on the solenoid current. The existence of a minimum threshold electric field for the discharge ignition of 4.0 × 10³ V/m is demonstrated for a multidirectional electrodeless plasma thruster without changing the discharge channel geometry within the studied parameter range, occurring at a solenoid current of I = 2 A.