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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

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Работа посвящена проблеме космического мусора, который представляет серьезную угрозу деятельности человека в околоземном космическом пространстве. Выполнен краткий обзор известных технических средств и технологий для удаления космического мусора с низких околоземных орбит. На сегодняшний день основными путями борьбы с техногенным засорением околоземного пространства является использование активных и пассивных методов увода космического мусора. Активные методы позволяют обеспечить увод объектов мусора в гарантированные сроки, однако требуют значительных энергетических затрат. Пассивные методы являются более экономичными, но процесс более продолжителен и может не удовлетворять требованиям международной конвенции по уменьшению космического мусора. В работе рассмотрен новый комбинированный подход к уводу крупногабаритного космического мусора в плотные слои атмосферы Земли, где он прекратит своё существование. Данный метод предполагает совместное использование активного средства-реактивной двигательной установки-и пассивного средства-аэродинамического парусного устройства. Представлена концепция использования комбинированного метода для очистки околоземного пространства, определена область эффективного его использования. Данная область составляет высоты от 700 до 2500 км в зависимости от баллистического коэффициента. В качестве критерия эффективности предложено рассматривать относительное отклонение массы компонентов топлива, необходимых для обеспечения увода. Показана целесообразность и перспективность применения комбинированного метода увода. Результаты работы представляют практический интерес для технико-экономического обоснования и разработки мероприятий по использованию комбинированного метода очистки низких околоземных орбит от элементов крупногабаритного космического мусора. Ключевые слова: загрязнение околоземного пространства, космический мусор, комбинированный метод очистки.

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This study shows a thick plate, pointed on both sides, of variable elongation under action of subsonic and supersonic air flows. The relevance of the work is driven by the development of new aircraft that change their geometry during the flight, in particular, aircraft that change the relative elongation during the flight. The aim of this work is to determine the aerodynamic fields around the pointed plate, taking into account its variable elongation, and to determine the effect of variable elongation on the drag coefficient of the plate. Numerical simulation was used as a research tool. Mathematical modeling is based on Reynolds-averaged Navier-Stokes equations. The original software was used for the analysis; the computational algorithm is based on the finite volume method. The mathematical model and algorithm of numerical computation of subsonic and supersonic air flow acting on a pointed plate are presented. A pattern of gas flow acting on the plate is obtained at zero angle of attack for Mach numbers from 0,6 to 4. The dependence of the plate drag coefficient on Mach number and relative elongation is obtained. It is established that for a thick pointed plate the drag coefficient significantly depends on the elongation for both sound and supersonic flow. As a result of numerical experiments, the anomalous behavior of the drag coefficient in the transonic area at large elongations was found. It is shown that the decrease in elongation leads to an increase in the drag coefficient of the plate, which affects the aerodynamic characteristics of the aircraft of variable elongation.

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.