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A method has been developed for the combined de-orbiting of large-size objects of space debris from low-Earth orbits using an electro-rocket propulsion system as an active de-orbiting means. A principal de-orbiting technique has been devised, which takes into consideration the patterns of using an electric rocket propulsion system in comparison with the sustainer rocket propulsion system. A procedure for determining the parameters of the de-orbiting scheme has been worked out, such as the minimum total speed and the time of the start of the de-orbiting process, which ensures its achievement. The proposed procedure takes into consideration the impact exerted on the process of the de-orbiting by the ballistic factor of the object, the height of the initial orbit, and the phase of solar activity at the time of the de-orbiting onset. The actual time constraints on battery discharge have been accounted for, as well as on battery charge duration, and active operation of the control system. The process of de-orbiting a large-size object of space debris has been simulated by using the combined method involving an electro-rocket propulsion system. The impact of the initial orbital altitude, ballistic coefficient, and the phase of solar activity on the energy costs of the de-orbiting process have been investigated. The dependences have been determined of the optimal values of a solar activity phase, in terms of energy costs, at the moment of the de-orbiting onset, and the total velocity, required to ensure the de-orbiting, on the altitude of the initial orbit and ballistic factor. These dependences are of practical interest in the tasks of designing the means of the combined de-orbiting involving an electric rocket propulsion system. The dependences of particular derivatives from the increment of a velocity pulse to the gain in the ballistic factor on the altitude of the initial orbit have been established. The use of these derivatives is also of practical interest to assess the effect of unfolding an aerodynamic sailing unit

A methodology for assessing the relative effectiveness of alternative options for building space object diversion systems has been improved. An algorithm for assessing the effectiveness of the system of removal of space objects from near-Earth orbit based on the method of integral assessment is given. It makes it possible to simplify the process of optimal choice of the method to divert space objects and determine efficiency in the early phases of the life cycle of rocket and space technology objects. The use of the appropriate toolset makes it possible to build a system for assessing the effectiveness of projects for the removal of space objects from low Earth orbits using various diversion methods (active, passive, combined). The analysis of defining world indicators of evaluation of objects of rocket and space technology based on regulations by international space agencies has been carried out. An indicator of the total integrated relative efficiency of projects of space object diversion systems from low Earth orbits has been proposed, which makes it possible to build the removal of passive, active, and combined methods for leveling the risks of space activities. It is argued that the selected combined system using an autophagic launch vehicle could reduce environmental losses and, as a result, reduce compensation payments to owners of space objects. The possibilities of building combined systems with reusable engines have been considered in order to reduce such indicators as the period of diversion and reduction of operating costs due to fuel economy.

Purpose Leading developers and providers in the modern space launch market note a splash in the development of ultralight launch vehicle (LV), driven by the growing demand for small satellites for large constellations in low Earth orbits. One of the promising ways to solve the problem of the quick launch of such satellites is to use a new type of ultralight launch vehicle with a plastic body. The project of such a launch vehicle was proposed by Oles Honchar Dnipro National University (Ukraine). Along with that, there is a need for appropriate research studies on the thermal resistance of the plastic shell, as the physical, mechanical and thermophysical characteristics of polymers significantly differ from traditional aerospace materials. The purpose of this study is to validate the design and ballistic parameters of such a launch vehicle in terms of providing an acceptable thermal environment at the atmospheric phase of the trajectory. Design/methodology/approach The workability of a new type of propulsion system is being investigated experimentally in bench conditions. To study the process of aerodynamic heating of a plastic shell, numerical modeling based on the integration of the flight dynamics and heat transfer equations is used. Findings Brief information about the design of a new type of ultra-light autophage launch vehicle with a plastic body is presented. A mathematical model for the movement of the launch vehicle at the atmospheric phase of the trajectory, and for the heating of the polyethylene body of the launch vehicle, taking into account the dynamic change in the atmospheric parameters is proposed. The influence of the motion trajectory on the thermal environment of the rocket body is investigated, rational motion trajectories and corresponding permissible g-loads are determined. Originality/value The fundamental possibility of using plastic (polyethylene) as a structural material and fuel for bodies of a new type of ultralight launch vehicles has been substantiated. It is shown that to ensure acceptable thermal conditions of a plastic body, it is necessary to use thermal insulation. It is proposed to use a polymeric Teflon coating as such thermal insulation. The results are important for the development of technologies for launching small satellites into orbit, as the use of plastic as the main structural material of the rocket body will significantly reduce the launch cost.