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This review deals with the selection of the electric propulsion system (EPS) for the internationally developed and designed, primary nuclear-electric space tug International Nuclear Power and Propulsion System (INPPS). INPPS is scheduled for interplanetary missions to Mars and Jupiter moon Europa missions by the end of decade 2020. Regarding specific technical and mission parameters preselected electric thruster (ET) types, developed by international companies and institutions, are analysed, evaluated and investigated for a possible application as propulsion system (PS), the so-called CET (Cluster of Electric Thrusters). It is analysed whether solely electric thrusters, combined in an adequate CET, enable the envisaged interplanetary missions—robotic and astronautic/crewed with the INPPS flagship.Thruster clusters with strategic consortium considerations are analysed as a feasible PS of the INPPS. The studied CET consists of the following: (a) only European ETs, (b) combination of German and European ETs, (c) Japanese and European ETs or at least (d) Japanese, European and US thrusters. The main results are (1) Robotic and crewed INPPS mission to Mars/Europa are realizable with EPS only (no chemical propulsion is needed), (2) that every CET, except (c) of only Japanese and part of European thrusters, is capable to perform the main part of envisaged INPPS flagship mission orbit to Mars, back to Earth and to Jupiter/Europa moon.

AbstractThe results of studies on the use of solid-state infrared lasers in combination with special targets for obtaining primary electrons in ionization chambers of plasma-ion thrusters are presented. Such thrusters being equipped with free molecular air intakes for using the surrounding atmosphere as a propellant for air-breathing electric thrusters, which are highly efficient for long-term maintenance of spacecraft in ultra-low orbits, providing significant advantages in Earth remote sensing and telecommunications. It is shown that the method of electron emission proposed can be an alternative to the current-heated cathodes used today, significantly increasing their lifetime.

To address the issues of excessive observation equipment and discontinuous observation points in tethered discrete profiling observations, a mobile profiling observation platform with mooring and cable-traversing capabilities has been designed. This platform is propelled by electric thrusters and incorporates a mooring mechanism that clamps onto the tether to achieve clinging positioning. A dynamic simulation model of the mobile platform was established to analyse the thrust required for traversing the cable and the clamping force and torque needed during mooring. The study reveals that the magnitude of horizontal flow velocity is the primary factor influencing the thruster thrust and clamping torque, both of which increase with higher flow velocities. The designed cable-traversing mobile observation platform provides a methodological approach for conducting profiling observations on tethered marine observation platforms.

In the actual working process of electric thrusters based on high-voltage electric fields, the discharge breakdown phenomenon is universal and complex, and such phenomena will have a significant impact on the thruster structure, working state and spacecraft system. In order to study the interpolar discharge breakdown characteristics of ionic liquid electrospray thrusters, a basic electrospray model and test system were constructed, and the change curves of discharge characteristic parameters such as breakdown voltage, threshold current, breakdown voltage frequency and so on in the range of 7×10 −3 ~10 5 Pa with air pressure and transmitter inner diameter were obtained, and the air pressure range that the electrospray model could work in normally was calibrated. The results show that the breakdown voltage characteristic curve of the electrospray model has typical minimum characteristics, and the minimum values all appear around 80 Pa. Lowering the air pressure below 10 −2 Pa can effectively increase the breakdown threshold between the poles and the emission current, thereby obtaining a larger voltage regulation range, and when the air pressure is reduced to 7×10 −3 Pa, the breakdown can reach more than 3200 V. The 60-μm inner diameter emitter performed better in the discharge experiment, and the breakdown threshold, emission current and operating area range were better than the slightly larger inner diameter emitter under the same working conditions.

Due to its high specific impulse, low toxicity, contamination, safety, and cost, krypton has gradually replaced other propellants as the primary choice for electric thrusters on deep-space exploration and large-orbit transfer missions with long-range, long-duration, and large-thrust requirements. However, due to the cryogenic propellant itself having a low boiling point, controlling the pressure change of krypton cryogenic propellant during the process of long-term cryogenic storage has emerged as one of the most important issues that must be resolved and proven to achieve non-destructive storage of cryogenic propellant. This paper examines the necessity and viability of a cryogenic gas propellant liquefaction storage scheme and quantifies the effectiveness of tank pressure control of the supercooled liquid krypton injection mixing scheme to solve the problem at present. Its purpose is to offer scheme examples for the design of a krypton thruster propellant storage unit and the in-tank pressure control during the long-term application process.

Hall thrusters are one of the most successful electric thrusters for space application that has been developed until now. The Plasma Physics Laboratory of the University of Brasília (UnB) has been developing a Permanent Magnet Hall Thruster (PHALL) for the Brazilian Space Program since 2004. Recently we have achieved important experimental results satisfying our initial goals of generating a force above 40 mN with powers around 620 W. We will discuss in this article possible applications of this thruster to nano and microsatellites with powers above 50 W. Meanwhile, a complete description is given of our present and future installations where the new thruster will be tested; taking advantage of our new 1.5 m diameter vacuum chamber (the old chamber had 0.5 m in diameter), which intends to test our thruster in the most realistic conditions, including mounting and testing on a 3U CubeSat structure, which is where we intend to start testing our thruster in a real mission in space.

This paper studies the performance of an interplanetary CubeSat equipped with a continuous-thrust primary propulsion system in a heliocentric mission scenario, which models a nodal flyby with a potential near-Earth asteroid. In particular, the mathematical model discussed in this work considers a small array of (commercial) miniaturized electric thrusters installed onboard a typical CubeSat, whose power-generation system is based on the use of classic solar panels. The paper also discusses the impact of the size of thrusters’ array on the nominal performance of the transfer mission by analyzing the trajectory of the CubeSat from an optimization point of view. In this context, the propulsive characteristics of a commercial electric thruster which corresponds to a iodine-fueled gridded ion-propulsion system are considered in this study, while the proposed procedure can be easily extended to a generic continuous-thrust propulsion system whose variation in thrust magnitude and specific impulse as a function of the input electric power is a known analytic function. Using an indirect approach, the paper illustrates the optimal guidance law, which allows the interplanetary CubeSat to reach a given solar distance, with the minimum flight time, by starting from a circular (ecliptic) parking orbit of assigned radius. The mission scenario is purely two-dimensional and models a rapid nodal flyby with a near-Earth asteroid whose nodal distance coincides with the solar distance to be reached.

In a worldwide scenario which sees an increasing number of small satellite launches, novel mission concepts may be unlocked providing the spacecrafts with the very precise and rapid maneuvering capability that electric thrusters cannot guarantee. In this context, chemical thrusters appear to be a possible solution. This work aimed to experimentally study and solve the problem of ignition for 10 N hybrid rockets based on hydrogen peroxide. Firstly, the study analyzed the performance of a monopropellant engine capable of functioning as a hybrid injection system. In particular, the effects of the liquid mass injected, the initial temperature, and the supply pressure on the pulsed engine performance were experimentally investigated. The injected mass showed a greater impact on the performance with respect to the starting chamber temperature and injection pressure. This thruster also showed a good potential for space applications. In the second part of the work, the objective was to find an ignition procedure that reduced propellant consumption and eliminated the need for a glow plug. This is important because the electrical power consumption in real applications significantly affects other subsystems and is undesirable for chemical engines. Different ignition procedures were tested to emphasize their respective advantages and disadvantages, and the findings indicated that the concept of pulsed preheating is feasible with only a small amount of propellant consumption, while substantially decreasing the ignition duration from approximately 45 min to a maximum of just 3 min. Finally, similar ignition procedures were adopted using different fuels. The results showed that PVC and ABS, under the same operating conditions, ignite more easily than HDPE, which requires an oxidizer consumption approximately double that of the other two fuels. Considerations about the effect of chamber pressure and oxidizer mass flow rate on engine ignition were also included.

Electric thrusters are used for in-space propulsion of spacecraft for a combination of practical and economic reasons. These devices that use electric power to accelerate the mass flow of the propellant at exhaust velocities are one or two order of magnitude faster than those achieved in conventional chemical propulsion. This feature results in significant propellant savings that allows longer mission times and heavier payloads, however, the thrusts achieved are lower for important electric power consumptions. The different electric thrusters that are currently used in Europe for both in-orbit corrections and also for orbit raising to the operational orbit from the separation stage of the launcher will be introduced. This propulsion system also serves to expel these vehicles from their orbit at the end of their useful life to control the increase of space debris in Earth orbit. Additionally, the characteristics of new low-power electric engines with thrust levels in the range 0.1-10 mN and electric power consumptions below 500 W will be discussed. These are required for flight formation, orbital corrections and end-of-life disposal in the new constellations of small satellites in low Earth orbit intended for planetary internet coverage and interactive television services.

This paper presents an overview of current world developments in the field of low-energy plasma propulsion systems for small micro - and nanoclass spacecraft. A promising direction of thrusters for devices of nanoclass are propulsion ion propulsion. The use of electrical energy and high-frequency electromagnetic radiation eliminates the problem of significant power losses, because at this level, the laws of optics (reflection and focusing of waves by metal surfaces and dielectrics) and electrical conductivity (energy transfer through metal conductors) work. This fact is an advantage of plasma installations over thermal ones.