Non-Propellant Eddy Current Brake and Traction in Space Using Magnetic Pulses

The safety of on-orbit satellites is threatened by space debris with large residual angular velocity and the space debris removal is becoming more challenging than before. This paper explores the non-contact despinning and traction problem of high-speed rotating targets and proposes an eddy current brake and traction technology for space targets without any propellant consumption. The principle of the conventional eddy current brake is analyzed in this article and the concept of eddy current brake and traction without propellant is put forward for the first time. Secondly, according to the key technical requirements, a brand-new structure of a satellite generating artificial magnetic field is designed accordingly. Then the control mechanism of eddy current brake and traction without propellant is analyzed qualitatively by simplifying the model and conditions. Then, the magnetic pulse control method is proposed and analyzed quantitatively. Finally, the feasibility of the technology is verified by the numerical simulation method. According to the simulation results, the eddy current brake and traction technology based on magnetic pulses can make the angular speed of target decrease linearly without propellant during the process. This technology has huge advantages compared with conventional eddy current brake technology in terms of efficiency and reduced propellant consumption.

Analysis of the advantages of an aerodynamic compensator in contactless space debris removal

A modified scheme of the known technology for contactless space debris removal, which is called Ion Beam Shepherd, is considered. This scheme uses an aerodynamic compensator in order to reduce the propellant consumption of the additional electrojet thruster of the shepherd spacecraft. The thruster serves to compensate the spacecraft motion caused by the action of the main electrojet thruster, whose ion plume “brakes” the space debris object. The aerodynamic compensator significantly increases the spacecraft cross-sectional area compared to the space debris object one. This fact, together with the aerodynamic perturbations acting in the direction perpendicular to the orbital plane, calls for estimating the propellant consumption of the control system thruster to maintain the required position of the spacecraft relative to the space debris object in that direction. The goal of this article is to identify the advantages of using the aerodynamic compensator in space debris removal from low Earth orbits using the Ion Beam Shepherd technology. The tasks of the study are to estimate the reduction in the cost of the momentum of the additional electrojet thruster during contactless space debris object de-orbiting due to the use of the aerodynamic compensator and the additional cost of the momentum of the thruster of the spacecraft – space debris object relative position control system to correct deviations perpendicular to the orbital plane. Using a number of simplifying assumptions, integral estimates of these costs are obtained. Using these cost estimates, it is shown that the use of an aerodynamic compensator is advantageous in terms of the cost of the saved electrojet thruster propellant (xenon) regardless of the type of the spacecraft control system thruster.

Optimal control of propellant consumption during vertical lifting of rocket in homogeneous atmosphere using regularized solution of integral equation of the first kind

The article is devoted to the study of the optimal control of propellant consumption during vertical lifting of rocket in homogeneous atmosphere using regularized solution of integral equation of the first kind. The problem of lifting of a rocket into desired height along optimal trajectory in the view of minimal consumption of propellant leads to solving the set of differential and integral equations. Problem of optimal control of propellant consumption during lifting of rocket in homogeneous atmosphere is solved using regularized solution of integral equation of the first kind which is solution of corresponding Euler equation on discrete time net. Influence of the regularization parameter and some additional parameters on precision of discreted problem is investigated. Considered algorithm is summed up easily to the case of non-homogeneous atmosphere by introducing dependence of the ballistic coefficient on altitude of flight and to problem of putting spacecraft into determined orbit and suborbital flights by setting desired altitude and velocity and modifying of motion equations.

Autonomous relay control system for satellite formation operating in low Earth orbit

A relay control algorithm is proposed for the satellite formation consisting of a «passive» (virtual) spacecraft moving in an unperturbed circular orbit, and several “active” spacecraft maneuvering relative to it. The purpose of control is to keep each active spacecraft on its finite path relative to the virtual spacecraft defining the center of formation with as low as possible propellant consumption. To describe the motion of maneuvering spacecraft relative to the center of formation, modified Hill–Clohessy–Wiltshire equations are used accounting for the earth's flattening and aerodynamic drag. The main emphasis of this paper is on the study of the existence of stable limit cycles and determination of their domains of attraction, as well as the search for values of the relay control system parameters ensuring a minimum propellant consumption for maintaining a proper dynamic formation behavior. Key words: relay control, point transformation method, formation flying, limit cycle, phase plane.

Transfers with a rendezvous lasting one to two orbits between near-circular coplanar orbits

The paper discusses orbit transfers involving spacecraft rendezvous which belong to the class of coplanar non-intersecting orbits of a spacecraft and a space station. The duration of the rendezvous is assumed to be limited to two orbits, because for longer durations there is a known optimal solution algorithm, where phasing is achieved through the optimal orbit-to-orbit transfer between coplanar orbits. The proposed programs include in the final leg of the transfer a three-impulse rendezvous program lasting one orbit, which was determined using the method of splitting the impulse burns for the optimal orbit-to-orbit transfer. The phasing needed to achieve the phase difference required at the start of the three-impulse rendezvous program is attained through maneuvering during the previous leg of the transfer and the splitting of the orbit transfer pulses, not resulting in an increased propellant consumption. The structure of the optimal rendezvous program as function of its duration was determined and computing formulas were obtained. The range of phase differences at the start of maneuvering was determined, within which the characteristic velocity of the rendezvous is equal to the characteristic velocity of the orbit-to-orbit transfer. The paper presents simulation results for “quick" rendezvous profiles that use the proposed programs. Key words: spacecraft, orbital station, «quick» rendezvous, orbit transfer, rendezvous program.

Solving the problem of minimax closed-loop control of liquid-propellant launch vehicle fuel consumption control

The paper provides mathematical formalization and a method of solving the problem of minimax (guaranteed) closed-loop terminal control of fuel consumption of a liquid-propellant launch vehicle power plant. The initial discrete-continuous nonlinear model of the controlled object is linearized along the given reference phase path and is approximated by a linear discrete-time multistep dynamical system. The approximated system includes the state vector, the control vector and the disturbance vector that defines the error of formation of the approximated model. Taking into account the geometrical constrains of control and disturbance vectors in the approximated system, we formulate the main problem of minimax closed-loop terminal control of propellant consumption of the launch vehicle’s propulsion system. This problem consists in solving a number of auxiliary tasks of minimax open-loop terminal control. To solve each of these tasks we use an instrument of development and analysis of generalized attainability domains of the approximated linear discrete dynamical system. These techniques are implemented by modifying the general recurrent algebraic method. To solve the problems under consideration we propose an approach and an appropriate numerical algorithm that is reduced to the implementation of a finite sequence of only one-step algebraic and optimization operations. The efficiency of the proposed approach to solving the problem under consideration is demonstrated and verified by a computer simulation example. This simulation example consists in controlling the process of propellant consumption for “Soyuz-2.1b” launch vehicle’s third stage propulsion system.