Search Results (82)

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. Full article

This paper has the goal of exploring the potential of electromagnetic propulsion systems based on tethers to create artificial Sun-synchronous orbits for remote sensing satellites, as well as performing station-keeping maneuvers and de-orbiting of the satellite after the end of its useful life. To create artificial Sun-synchronous orbits, the force is applied to keep the longitude of the ascending node with the same angular velocity of the apparent motion of the Sun around the Earth, which is the definition of a Sun-synchronous orbit. These orbits are very important for remote sensing satellites, because in these orbits the satellite passes by a given point at the same time, helping in analyzing the data collected. The use of electrodynamic tethers can extend the regions of Sun-synchronous orbits, both in terms of inclination and semi-major axis. To perform the de-orbiting of the satellite, the same tether can apply a force in the opposite direction of the motion of the satellite, so reducing its energy and decreasing the semi-major axis until the satellite crashes into the atmosphere of the Earth. This is very important to avoid increasing the presence of space debris in space, a very serious problem nowadays. For the station-keeping maneuvers, we just need to use the appropriate control laws, from time to time, to correct any errors in the Keplerian elements. A significant advantage of employing an electrodynamic tether over traditional thrusters is that it does not require consumption of fuel. The study assumes that a current can flow in both directions through the tether, so interacting with the magnetic field of the Earth to create the Lorentz force. The possibility of using electrodynamic tethers with autonomous charge generation, to avoid dependence on plasma densities and other external factors, is considered. The results presented here help in space and planetary science, since they give more options for remote sensing satellites, which are a key element in planetary science. Full article

The development and execution of prospective inner and outer space missions require focusing on the use of many small space vehicles operating in swarms with multiple informational, navigational, and mission-oriented interactions among themselves. Such missions involve providing communication and surveillance services, facilitating distributed material production in space, and conducting research expeditions to explore the resources and environments of new worlds. The cornerstone technology for operating distributed space systems is propulsion. Among a range of propulsion technologies—from using pressurized cold gases to implementing laser beams to generate thrust—certain methods stand out for application in small spacecraft. This paper provides a summary of space-operated propulsion, emphasizing the reasons for the more frequent adoption of one technology over another. The discussion on propulsion trends is complemented by examining the physical, engineering, production, operational, and societal rationale behind these choices. The findings reinforce the trend toward transitioning to fully electric satellites. This review serves as a means for reevaluating global propulsion trends and guiding the future development of inner and outer space propulsion-assisted economies effectively. Full article

Magnetoplasmadynamic thrusters (MPDTs) are becoming increasingly viable as electric propulsion (EP) technology for space missions, yet their complex plasma behaviour, intricate thrust-generation process, and nonlinear multi-physics thrust–field interactions prove difficult for conventional modelling approaches, including empirical techniques. Traditional empirical modelling shortcomings include failure to predict accurately across wide operational regimes. This paper introduces a physically interpretable, artificial intelligence (AI)-powered thrust model for Applied-Field Magnetoplasmadynamic Thrusters (AF-MPDTs), developed using symbolic regression (SR) to address the gap between data-driven prediction and physics-based understanding. The proposed method, an alternative to traditional black box AI methods, incorporates physics-aware composite-term operators, ensuring that the resulting analytical expressions are bounded by known physical behaviours while retaining the flexibility to discover previously overlooked nonlinear couplings. A comprehensive dataset of AF-MPDTs undergoes rigorous preprocessing to ensure dimensional consistency and noise robustness. The SR model then evolves candidate equations, balancing predictive accuracy with interpretability through Tree-Structured Parzen Estimator (TPE) optimisation. The results, closed-form surrogate correlations with 95.98% of accuracy as goodness of fit, root mean square error of 0.0199, mean absolute error of 0.0143, and mean absolute percentage error reduction of 28.91% against the benchmark model in the literature. A post-discovery protocol for numerical robustness and physical consistency is implemented, with Shapley Additive Explanations (SHAP) providing insight into the influence of each composite-term in the developed correlation, followed by a numerical robustness and physical consistency validation using a Monte Carlo (MC) envelope. A StabilityScore is calculated for all developed correlations, enabling explicit accuracy–complexity–stability comparisons. In doing so, we demonstrated that SR can systematically recover known physical relationships—such as the scaling of thrust with discharge current and applied magnetic field—while proposing interpretable higher-order corrections that improve fit quality. The resulting SR-based thrust models not only achieve competitive accuracy relative to state-of-the-art numerical and empirical methods but also offer more explainable and interpretable results capable of revealing compact formulations that capture essential acceleration mechanisms with transparency. Overall, this paper, using SR, advances explainable AI (XAI) methodologies capable of generating trustworthy, analytically transparent models for next-generation electric propulsion systems. Full article

Propulsion systems aboard small satellites assisting dynamic space missions at the proximity of deep space natural objects may face challenges in long-term non-serviceable operations, achieving thrust vector direction control, and adapting to severe environmental conditions. The proposed propulsion solution involves using a pulsed plasma thruster with multiple spark plugs for uniform ignition and thrust vector control, enhancing reliability and efficiency. Key advantages of the use of such an approach include minimal power consumption, an efficient volume utilization, and enhanced reliability through redundant ignition points realized within a single thruster head. Experimental validation demonstrates the effectiveness of the proposed architecture, confirming uniform ignition patterns. Also, the results of the experimental investigation qualitatively demonstrate the capability of thrust vector control by selectively discharging the distributed spark plugs. It can be supposed that this approach supports the viability of small satellites in dynamic space missions, promising dynamic, resilient, and reusable proliferated space systems for development of deep space economies. Full article

During ion thruster operation, electromagnetic waves propagating through the plasma plume undergo absorption and refraction effects. This paper presents a graphics processing unit (GPU) parallel ray tracing (RT) algorithm for inhomogeneous media to analyze plasma plume-induced perturbations on the radiation characteristics of a satellite reflector antenna, substantially improving computational efficiency. This algorithm performs ray path tracing in the plume, with the vertex and central rays in each ray tube assigned to dedicated GPU threads. This enables the parallel computation of electromagnetic wave attenuation, phase, and polarization. By further applying aperture integration and the superposition principle, the influence of the plume on the far-field antenna radiation patterns is efficiently analyzed. Comparison with serial results validates the accuracy of the algorithm for plume calculation, achieving approximately 319 times speed-up for 586,928 ray tubes. Within the 2–5 GHz frequency range, the plume causes amplitude attenuation of less than 3 dB. This study provides an efficient solution for real-time analysis of plume-induced interference in satellite communications. Full article

The study presents an upgraded design and the results of experimental investigations of a solid-propellant pulsed plasma thruster (PPT) in which graphite simultaneously serves as both the propellant and the ignition element. The proposed configuration comprises a planar parallel system of copper electrodes and a graphite initiating electrode equipped with an electromagnetic discharge-triggering mechanism. Experimental tests were conducted under vacuum conditions of approximately 10⁻⁵ Torr at an energy-storage capacitor voltage of 800–1400 V. Discharge current amplitudes of up to 3.16 kA were recorded at a single-pulse energy of up to 4.41 J. The measured impulse bit was about 17.1 μN $·$s, and the plasma jet exhaust velocity reached 11.1 km/s. Spectroscopic analysis of the plasma confirmed the presence of characteristic carbon emission lines, thereby indicating the active participation of the graphite propellant in the formation of the plasma plume. The present work continues previous research on PPTs with graphite electrodes and is aimed at further miniaturization of the earlier developed design. The primary objective of the study is the experimental validation of the proposed discharge concept in a planar parallel electrode configuration while preserving the key thrust and energy performance characteristics of the thruster. Full article

We study thrust production in a single-fluid magnetohydrodynamic (MHD) thruster with Hall-type coaxial geometry and show how velocity–field alignment and magnetic topology set the operating regime. Starting from the momentum equation with anisotropic conductivity, the axial Lorentz force density reduces to $f_z={\sigma }_{\theta z}{E}_z{B}_r(\chi -1)$, with the motional-field ratio $\chi \equiv \left(u{B}_r\right)/E_z$. Hence, net accelerating force ($f_z>0$) is achieved if and only if the motional electric field $E_m=u{B}_r$ exceeds the applied axial bias $E_z$ ($\chi >1$), providing a compact, testable design rule. We separate alignment diagnostics (cross-helicity $h_c=\mathit{u}·\mathit{B}$) from the thrust criterion ($\chi$) and generate equation-only axial profiles for $\chi \left(z\right)$, $j_{\theta }\left(z\right)$, and $f_z\left(z\right)$ for representative parameters. In a baseline case $(E_z=150\mathrm{V}{\mathrm{m}}^{-1},{\sigma }_{\theta z}=50\mathrm{S}{\mathrm{m}}^{-1},u_0=12\mathrm{km}{\mathrm{s}}^{-1},B_{r0}=0.02\mathrm{T},L=0.10\mathrm{m})$, the $\chi >1$ band spans $\approx 21.2\%$ of the channel; a lagged correlation peaks at $\Delta z^{★}\approx 8.82\mathrm{mm}$$(C_{HU}=0.979)$, and ${\int }_0^L{f}_z\mathrm{d}z$ is slightly negative—indicating that enlarging the $\chi >1$ region or raising ${\sigma }_{\theta z}$ are effective levers. We propose a reproducible validation pathway (finite-volume MHD simulations and laboratory measurements: PIV, Hall probes, and thrust stand) to map $f_z$ versus $\chi$ and verify the response length. The framework yields concrete design strategies—$B_r\left(z\right)$ shaping where u is high, conductivity control, and modest $E_z$ tuning—supporting applications from station-keeping to deep-space cruise. Full article

The paper presents experimental results for a modified pulsed plasma thruster (PPT) with solid propellant, using a coaxial anode–cathode design. Graphite from pencil leads served as propellant, and a tungsten trigger electrode was tested to reduce carbonization effects. Experiments were performed in a vacuum chamber at 0.001 Pa, employing diagnostics such as discharge current/voltage recording, power measurement, ballistic pendulum, time-of-flight (TOF) method, and a Faraday cup. Current and voltage waveforms matched an oscillatory RLC circuit with variable plasma channel resistance. Key discharge parameters were measured, including current pulse duration/amplitude and plasma channel formation/decay dynamics. Impulse bit values, obtained with a ballistic pendulum, reached up to 8.5 μN·s. Increasing trigger capacitor capacitance reduced thrust due to unstable “pre-plasma” formation and partial pre-discharge energy loss. Using TOF and Faraday cup diagnostics, plasma front velocity, ion current amplitude, current density, and ion concentration were determined. Tungsten electrodes produced lower charged particle concentrations than graphite but offered better adhesion resistance, minimal carbonization, and stable long-term performance. The findings support optimizing trigger electrode materials and PPT operating modes to extend lifetime and stabilize thrust output. Full article

The development of dynamic missions of small satellites requires the development of efficient, compact, and reliable propulsion systems (PSs). This paper investigates a propellant storage and supply system (PSSS), utilizing alternative solid-state propellants in the form of wire. To establish the background to the suggested solutions implemented in the proposed system, two types of comparative analysis were performed. The first one compared different types of propellant management system designs while the second juxtaposes a variety of propellants. It is shown that the solid-state systems for small satellite operations are advantageous, while the selection of propellants should be focused on safe operations and operational requirements. The principle of operation and structural design of the proposed wire-based solid-state propellant management system are discussed, including the assessment of its engineering realization. The strategies to mitigate the potential problems with the system’s operations such as propellant unwanted deposition and corrosive effects are suggested. An example of using the proposed system is provided, which considers a deep space dynamic mission case. The proposed PSSS architecture is dedicated to increasing the energy efficiency, resilience to environmental factors, and suitability for small satellite platforms, including that of the CubeSat format. Full article

Very low Earth orbit (VLEO) satellites are confronted with the challenge of orbital decay caused by thin atmospheres, and the volume and power limitations of micro satellites further restrict the application of traditional electric propulsion systems. In response to the above requirements, this study proposes an innovative scheme of radio frequency plasma micro-thrusters based on magnetic nozzle acceleration technology. By optimizing the magnetic nozzle configuration through the system, the plasma confinement efficiency was significantly enhanced. Combined with the mixed working medium (5 sccm Xe + 10 sccm air), the thrust reached 1.7 mN at a power of 130 W. Experiments show that the configuration of the magnetic nozzle directly affects the plasma beam morphology and ionization efficiency, and a multi-magnet layout can form a stable trumpet-shaped plume. The air in the mixed working medium has a linear relationship with the thrust gain (60 μN/sccm), but xenon gas is required as a “seed” to maintain the discharge stability. The optimized magnetic nozzle enables the thruster to achieve both high thrust density (13.1 μN/W) and working medium adaptability at a power level of hundreds of watts. This research provides a low-cost and miniaturized propulsion solution for very low Earth orbit satellites. Its magnetic nozzle-hybrid propellant collaborative mechanism holds significant engineering significance for the development of air-aspirating electric propulsion technology. Full article

A collaborative inter-laboratory study was conducted to characterise the performance of the novel 250 W External Discharge Plasma Thruster (XPT) with a channel-less Hall effect-type thruster designed to address lifetime limitations and lower-power efficiency challenges in conventional Hall effect thrusters. This study aimed to validate performance measurements across different facilities and thrust stands, investigating potential facility effects on thrust characterisation. Performance testing was conducted both at the University of Surrey using a torsional thrust balance and at the University of Southampton with a double inverted pendulum thrust stand, providing independent verification of the thrust and efficiency metrics. The comparison highlighted the importance of cross-facility testing with differing background pressures, calibration methods, and thrust balance types. These differences provide valuable insights, ensuring more robust and reliable low-power thruster characterisation. The XPT thruster demonstrated consistent performance across both the University of Surrey and University of Southampton facilities, with thrust levels ranging from 1.60 mN to 11.8 mN, specific impulses from 327 s to 1067 s, and anode efficiencies up to 11%. Higher anode voltages and mass fluxes at Southampton enabled extended operational envelopes, revealing performance plateaus at elevated powers, particularly for flow rates above 8 sccm. Cross-facility testing highlighted facility-dependent influences, with Southampton achieving a higher thrust and specific impulse at lower flow rates (5–6 sccm) due to increased anode currents, while discrepancies between test sites of up to 25% were observed at higher flow rates (8–10 sccm) and powers above 200 W. Characterisation identified an optimal operating range at 200 W of anode power with a mass flux below 8 sccm. This work underscores the importance of inter-laboratory validation in electric propulsion testing and provides insights into the best practices for assessing next-generation Hall effect-type thrusters. Full article

This research paper provides a conceptualization of a new type of plasma thruster based on screw-pinch physics and on the magnetic mirror concept. The article proposes a method to size a screw-pinch with a non-uniform axial magnetic field as a plasma thruster and to estimate its propulsive performance. The results obtained show that the plasma thruster is suitable for space missions inside the Earth’s sphere of influence and for space transportation of small satellites. Full article

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. Full article

The current study, as part of a PhD project on the design of a helicon thruster, aims to provide a rational methodology for the design of the helicon thruster’s main component, i.e., the helicon antenna. A helicon thruster is an innovative electrodeless plasma thruster that works by exciting helicon waves in a magnetized plasma, and its antenna is capable of producing a uniform, low-temperature, high-density plasma. A magnetic nozzle is used to accelerate the exhaust plasma in order to generate a propulsive thrust. In this paper, we consider a simple helicon antenna, specifically the Nagoya type-III antenna. We consider a common experimental setup consisting of a quartz tube with finite length containing a uniform magnetized plasma and a Nagoya type-III antenna placed at the tube centre. Considering previous studies on helicon waves theory, we compare three different design methods, each based on simplifying different modelling assumptions, and evaluate the predictions of these models with results from full-wave 3D simulations. In particular, we concentrate on deriving a rational design method for the helicon antenna length, given the dimension of the quartz tube and the desired target plasma parameters. This work aims to provide a practical and fast method for dimensioning the antenna length, useful for initializing more accurate but computationally heavier full-wave simulations in 3D geometry or simply for a rapid prototyping of the helicon antenna. These results can be useful for the development of a helicon thruster but also for the design of a high-density radiofrequency plasma source. Full article