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.

Introduction. A constant increase in the amount of space debris already constitutes a significant threat to satellites in nearEarth orbits, starting with the trajectory of their launch vehicle injection. Problem Statement. Design and development of various modern methods of protection against space debris requires knowledge of the statistical characteristics of the distribution of the kinematic parameters of the simultaneous motion of a launch vehicle injecting satellite and a group of space debris objects in the area of its trajectory. Purpose. Development of a mathematical model of a launch vehicle rendezvous with a group of observable orbital debris while injecting a satellite into near-earth orbits with an altitude of up to 2100 km and an inclination from 45 to 90 degrees. Materials and Methods. The following methods are used in the research: analysis, synthesis, comparison, simulation modeling, statistical processing of experimental results, approximation, correlation analysis, and the least squares method. Results. The simultaneous motion of a launch vehicle and a group of space debris objects has been studied. The distributions of relative distance, relative velocity, angle of encounter, and moments of time of approach of a launch vehicle to a group of the observed space debris at a relative distance of less than 5 km have been obtained. The dependence of the average rendezvous concentration on the distribution of space debris across the average altitude of the orbit and the inclination of the target orbit of the launch vehicle has been determined. The dependence of the average probability of rendezvous in the launch on the inclination of the target orbit, the number of orbital debris, and the relative distance of the rendezvous has been determined. Conclusions. The obtained mathematical model of rendezvous of a launch vehicle with a group of observed orbital debris can be used while designing means of cleaning the near-Earth space and systems to protect modern satellite launch vehicles from orbital debris. In addition, the results of the research can be used to assess the impact of unobserved orbital debris on the flight of a launch vehicle.

This article is devoted ascent depreciation in an earth orbit of space crafts for different functions that allow to solve many problems connected with space activity, in particular, maintenance of wider use of space resources and more effective clearing of a near space from space debris. The last is particularly true for bulky space objects of a technogenic origin which can be deorbit from earth orbits by using the special means for deorbiting injection into the target orbit by launch vehicles. As the ascent of such means for deorbiting by orbital launch vehicle demands significant financial expenses, for their decreased application of easy autophagy launch vehicles with burned fuel tanks is offered. The basic making unit of such launch vehicles is an autophagy engine. The design of autophage engine design and the principle of its operation of the version with compulsory fuel supply to the combustion chamber are presented and the expediency of use of such an engine for small spaceships in case of the impossibility of application of rather heavy usual systems of storage and fuel supply is shown. However, now the weight of the accepted systems of compulsory fuel supply autophagy engines for spaceships remains significant for launch vehicles. As one of the alternative ways of a feed the engine of autophagy launch vehicle, the capability of submission by using the inertia of the most fuel charge is considered. The unit of the autophagy engine contains an evaporator and works in a pulse mode, supplying fuel supply between momentums at the moment of low pressure in the engine. The effect of retardation (easings), representing pressure decrease and engine thrusts in flight time is established. This effect is called by the forces of inertia acting on the sliding engine, and has экспоненциальную an easing picture. Considering the effect of retardation, a system of the equations for calculating kinematic and design data of the launch vehicle with the autophagy engine is offered. The received results of the research have scientific and practical value and can find wide applications in the design of schemes and deorbiting means.

Introduction. A constant increase in the amount of space debris already constitutes a significant threat to satellites in nearEarth orbits, starting with the trajectory of their launch vehicle injection. Problem Statement. Design and development of various modern methods of protection against space debris requires knowledge of the statistical characteristics of the distribution of the kinematic parameters of the simultaneous motion of a launch vehicle injecting satellite and a group of space debris objects in the area of its trajectory. Purpose. Development of a mathematical model of a launch vehicle rendezvous with a group of observable orbital debris while injecting a satellite into near-earth orbits with an altitude of up to 2100 km and an inclination from 45 to 90 degrees. Materials and Methods. The following methods are used in the research: analysis, synthesis, comparison, simulation modeling, statistical processing of experimental results, approximation, correlation analysis, and the least squares method. Results. The simultaneous motion of a launch vehicle and a group of space debris objects has been studied. The distributions of relative distance, relative velocity, angle of encounter, and moments of time of approach of a launch vehicle to a group of the observed space debris at a relative distance of less than 5 km have been obtained. The dependence of the average rendezvous concentration on the distribution of space debris across the average altitude of the orbit and the inclination of the target orbit of the launch vehicle has been determined. The dependence of the average probability of rendezvous in the launch on the inclination of the target orbit, the number of orbital debris, and the relative distance of the rendezvous has been determined. Conclusions. The obtained mathematical model of rendezvous of a launch vehicle with a group of observed orbital debris can be used while designing means of cleaning the near-Earth space and systems to protect modern satellite launch vehicles from orbital debris. In addition, the results of the research can be used to assess the impact of unobserved orbital debris on the flight of a launch vehicle.

Debris in Earth's orbit poses a significant threat to existing spacecraft and hinders the safe launch of new spacecraft. This paper discusses various methods of space debris removal from orbit. In particular, active, passive, and combined methods are considered. The possibility of using  suborbital rockets to    launch   the    removal    system   into  orbit  is  also considered. The purpose of this paper is to review existing methods and technical means of space debris de-orbiting and analyze and identify promising areas for the development of such systems. A patent analysis was conducted for the period from 2002 to 2022. Modern scientific works on various methods of space debris removal were also analyzed. The papers on the combined method of space debris removal were analyzed separately. Non-standard approaches to space debris removal are considered. In particular, the possibility of using suborbital launch vehicles to launch removal systems into orbit is considered. As a result of the work, the current state of development of methods and technical means for space debris removal was determined. It has been found that the most commonly patented and studied in scientific papers are technical means based on the passive removal method. At the same time, the active method receives less attention in scientific papers. Because of this, it has been established that the study of passive and combined methods and the development of technical means based on these methods are promising areas. It was found that the combined method is promising, but little was studied, and the basic requirements for systems using suborbital launch vehicles were also identified. This work allowed us to summarize the current state of the problem of space debris management and identify the most promising areas for the development of space debris removal systems.