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This paper aims to suggest a new thermonuclear space propulsion and electric generator for aerospace.

Methods of thermonuclear physics are used for research.

The paper applies, develops and researches mini‐sized Micro‐AB thermonuclear reactors for space propulsion and space power systems. These small engines directly convert the high‐speed charged particles produced in the thermonuclear reactor into vehicle thrust or vehicle electricity with maximum efficiency. The simplest AB‐thermonuclear propulsion offered allows spaceships to reach speeds of 20,000‐50,000 km/s (1/6 of light speed) for fuel ratio 0.1 and produces a huge amount of useful electric energy. The offered propulsion system permits flight to any planet of the solar system in a short time and to the nearest non‐Sun stars by E‐being or intellectual robots during a single human life period.

Technical limitations may be apparent.

The theory of this propulsion and electric generator is developed and possibilities researched.

The paper seeks to propose and analyze a new electrostatic ramjet space engine.

The upper atmosphere (85‐1,000 km) is extremely dense in ions (millions per cubic cm). The interplanetary medium contains positive protons from the solar wind. A charged ball collects the ions (protons) from the surrounding area and a special electric engine accelerates the ions to achieve thrust or decelerates the ions to achieve drag. The thrust may have a magnitude of several Newtons. If the ions are decelerated, the engine produces a drag and generates electrical energy. The theory of the new engine is developed.

It is shown that the proposed engine driven by a solar battery (or other energy source) cannot only support satellites in their orbit for a very long time but can also work as a launcher of space apparatus. The latter capability includes launch to high orbit, to the Moon, to far space, or to the Earth's atmosphere (as a return thruster for space apparatus or as a killer of space debris). The proposed ramjet is very useful in interplanetary trips to far planets because it can simultaneously produce thrust or drag and large electric energy using the solar wind.

Two scenarios, launch into the upper Earth atmosphere and an interplanetary trip, are simulated and the results illustrate the excellent possibilities of the new concept.

This paper aims to provide an overview of current and historical arcjet development. The reviewed arcjets are considered with respect to both design and thruster relevant parameters. Correspondingly, the paper enables the identification of adequate design criteria and of the probable thruster parameters.

The approach consists of a database for thruster relevant parameters in conjunction with relevant operational requirements (such as type of propellant) and specific design criteria (such as e.g. propellant injection systems).

The synopsis of both operational parameter and respectively assigned design allows for the derivation of development approaches for arcjets under given high level requirements such as power regime.

The paper is a general review. However, its strength is in the synthesis of the arcjet classification, the functional evidence of design criteria and the application scenario.

Not only basic but also specific design criteria are analyzed and evaluated leading to a recommendation feature of the paper with respect to the overall design of adequate arcjets.

Within the scope of the Clean Space initiative, new applications and scenarios from the operation of arcjets arise enabling EOL phases of spacecraft that fulfill respective levels of debris mitigation and, in addition, the requirements concerning the adequately adapted re-entry of spacecrafts that are at end of life.

The paper is a general review. However, its strength is in the synthesis of the arcjet classification, the functional evidence of design criteria and the application scenario.

Interstellar gas passing through the solar system may effect the interplanetary gas, planetary atmospheres and satellite orbits. Interaction of the interstellar and interplanetary gases is considered; a solar system corona may be formed.

The author's aim is to offer a revolutionary method – the non‐rocket transfer of energy and thrust into Space with distances of millions of kilometers.

The author develops the theory and makes the computations.

The method is more efficient than transmission of energy by high‐frequency waves.

The method may be used for space launch and for acceleration of spaceships and probes to very high speeds, up to relativistic speed by current technology.

The research presented contains prospective projects which illustrate the possibilities of the suggested method.

Field emission electric propulsion (FEEP) thruster is a type of electric propulsion based on space‐proven indium liquid metal ion sources. The lifetime of FEEP thruster limits its application in space. A better method to improve its lifetime is the reduction of emission current. This paper aims to discuss the minimum emission current of operating FEEP thrusters.

In this work, theoretical models, including fluid‐flow model and ion formation model, are analyzed. Current densities of these models are discussed and the minimum emission current is calculated.

There are few equilibrium states under low emission current conditions. However, the minimum emission current is the only stable state at which the FEEP thruster can operate.

This analysis is mainly based on the needle indium FEEP, which is compared indirectly with experiments of gallium.

This paper attempts to help designers choose appropriate electric parameters to improve the lifetime of FEEP.

By introducing and analyzing theoretical models, this paper calculates the minimum emission current for stable operation of FEEP.

“Microscale light sails” (MLS) are simultaneously manufactured and launched as a matter‐beam from a proposed Lunar facility. Lunar aluminum would be refined for the feedstock of this “thin film beam generator”. A battery of linear, aluminum‐vapor, rocket engines make up the first stage of a “laser cooled thermal beam”. After a supersonic expansion, the condensing sheets of AlI atoms undergo light‐force mediated cooling, guidance, and compression. The individual, partly condensed sheets are brought together at sufficiently low energy to form the core of the thin film. MLS‐swarms can become either the reaction‐mass for a deep space, beam‐propulsion transportation network, the constituents of an orbital space‐mirror or an interstellar, laser‐driven probe, or simply be used as raw building material for outer space structures. An articulation of the beam generator may manufacture solar cells and other kinds of thin‐films from space resources.

Author's preference for use of term magnetofluidmechanics instead of magnctohydromechanics; résumé of main principles; applications of the science in astronautics.

To suggest a new revolutionary electrostatic linear engine.

Methods of the electrostatic physic are used for research.

Theory of this engine is developed and its possibilities researched.

This engine gives a big thrust (up 4 × 10⁵ N/m²), uses a high voltage electricity and light wires.

This engine can be used as a linear engine (accelerator), as a strong space launcher, as a high speed delivery system for space elevator, Earth‐Moon cable transport space, for an electrostatic levitation train, as a conventional high voltage rotating engine, as an electrostatic electric generator weapon (high speed gun), and so on. Theory of engine applications was developed and it shows powerful possibilities in space, transport and military industries. The projects are computed and show the good potential of the offered new concept.

Succeeds in proposing a new revolutionary electrostatic linear engine.

Thruster point assembly mechanism (TPAM) of the electric propulsion system allows to adjust the thrust vector, so that the thrust vector is directed to the satellite center of gravity (COG) during the satellite on-orbit working period. In this way the impact of disturbance torque caused by deviation of the thrust vector from the satellite COG during thruster ignition can be decreased. Therefore, the control accuracy of satellite is influenced directly by the control accuracy of TPAM. On the other hand, the on-orbit application of TPAM is restricted to the on-orbit computer resource. Therefore, the purpose of this paper is to design a control strategy for TPAM, and the strategy should not only be able to control the TPAM precisely but also be easily implemented by the on-board computer.

First, the structure and work principle of TPAM are discussed, and the mathematical model based on D-H coordinate system is built for it. Then the fitting methods are utilized to design the control strategy of TPAM. Absolute position fitting-based control strategy and relative position fitting-based control strategy are designed, and the least squares algorithm is introduced for parameter selection.

Simulations and tests are provided for the TPAM. Compared with the state-of-the-art PD controller, the proposed control strategy shows smaller overshoot and more simple realization. The experiment results are matched with the simulation results and both the experiment and simulation results show the validity of the proposed control strategies.

The designed control strategies can be used for the TPAM of some satellite’s electric propulsion system.

The mathematical model of the TPAM based on D-H coordinate system is given. The absolute position fitting-based control strategy and relative position fitting-based control strategy are proposed. Compared with existing methods, the two control strategies have more simple structure and smaller amount of computations. Furthermore, the relative position fitting-based control strategy achieves high precision with simple structure.