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The problem of forming a nominal optimal program for controlling the direction of the low-thrust vector of a spacecraft for a significant change in the elements of the spatial orbit is considered. A scheme for the joint change of the elements of the orbit within the framework of a simple structure for controlling the thrust vector on the orbit is proposed. [ABSTRACT FROM AUTHOR]

The problem of near Earth space debris is studied in this article. A special spacecraft debris collector, equipped with electro rocket engine of low-thrust, for large space debris disposal is introduced. The mass model of spacecraft debris collector, one-off or reusable, is obtained. Ballistic scheme, spacecraft debris collector orbital transfer from parking orbit to space debris orbit, its disposal in Earth atmosphere and reusable spacecraft return to the parking orbit are studied. For the introduced criteria of transport transfer efficiency and assumption about consistency of acceleration from thrust, analytical relations for spacecraft design parameters calculation are obtained. The design parameters calculation results are shown in a generalized form. [ABSTRACT FROM AUTHOR]

The problem of electro rocket engine type and its main design parameters optimization for spacecraft located at low circular Earth orbit is studied. The efficiency of chemical (impulse) and electro rocket engine of low-thrust is analyzed. Introduced criteria of efficiency – payload mass maximization, introduced assumption that engine exhaust velocity is constant. In accordance with criteria and assumption, the analytical solution for spacecraft design parameters of spacecraft is obtained. Also, analytical solution for area of efficiency of electro rocket low-thrust engine is obtained. Calculation of spacecraft design parameters equipped with electro rocket low-thrust engine and area of its efficiency applying is carried out. [ABSTRACT FROM AUTHOR]

The problem of fragment of space debris located at geostationary orbit disposal is studied. For this purpose, a special spacecraft debris collector is used. The main part of disposal operation – rendezvous transfer with fragment of space debris - is studied. Relative motion of a spacecraft debris collector about the fragment of space debris is considered in an orbital cylindrical reference frame. Secular and periodic components of relative motion are determined in an implicit form. Spacecraft debris collector motion control is carried out by its transversal acceleration from thrust switching. Radial thrust component is equal to zero. Boundary conditions for rendezvous transfer are formulated in terms of secular and periodic components of motion. The Motion control program for secular components contained two active areas and passive areas between them. Active areas duration were determined as functions of boundary conditions and passive area duration (free parameter). An analytical solution for the determined control program is formulated in an implicit form. Satisfaction of boundary conditions for periodic components is carried out by passive area duration and rendezvous start moment determination. Equations to determine these parameters are formulated. Numerical simulation is carried out. [ABSTRACT FROM AUTHOR]

the aim of this work is to study the possibility of the disposal of the space debris in the dense layers of the Earth's atmosphere. We use information about the fragments of the space debris location that stored in NORAD databases. There two types of the orbits of the space debris are explored: the high elliptical orbit and the circular orbits. The algorithm of the operation of the utilization of the fragments of the space debris in the dense layers of the Earth's atmosphere using the special spacecraft debris collector is composed. The spacecraft debris collector must be equipped with electro propulsion engines with low-thrust. The mathematical model of the plane motion of the spacecraft debris collector, in the form of the system of osculating elements, is composed. Problem of the descent of the fragments of the space debris reduced to the problem of the reducing of the height of the perigee of elliptical orbit. Obtained the locally - optimal control program of the orientation of the thrust vector. Using the L.S. Pontryagin's maximum principle, we solved the boundary value problem of the optimal control (decrease) of the perigee of the elliptical orbit. The results of the parametric study of the problem in the space of the initial orbits and the acceleration traction of the electro propulsion engines is shown. Numerical simulation of the space debris reduction from the low circular orbits is performed. Determined the number of revolutions passed a fragment of the space debris around the Earth during braking due to aerodynamic drag. Calculated the estimates of the gains of the characteristic velocity to perform the operation. [ABSTRACT FROM AUTHOR]