الفهرس 
CHAPTER I INTRODUCTIONOnly 14 pages are availabe for public view
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Abstract
Growing the number of space debris has recently become a topic of eminent concern. Therefore, orbital debris removal has become a very grave issue to prevent loss of spacecraft due to debris collision, it is very essential to address the aggregate risk, which need an efficient way to remove or avoid collision with operating satellites. This work develops an analytical method for orbital decay using Electrodynamics tether system (EDT) as a technique of active debris removal to avoid collision with operational satellites. Gaussian form of Lagrange Planetary equations used to evaluate the orbital motion of EDT with environmental perturbations of electrodynamic force and aerodynamic drag. Differential equations for the induced voltage-current across EDT are derived and solved with boundary conditions determined by mission objectives and duration of the deorbit. The change in electric current in the EDT due to its plasma environment and thermal conditions are considered. Analyses of different parameters of EDT dynamics with variations in the mass, type of the materials of the wire of tether, and the tether length were studied to classify the range of eccentricity of the elliptical orbit for possibility of deorbit the desired missions. Applied the current model of the EDT find that the orbit of a satellite deorbited will become elliptical in near polar orbits due to the higher-order perturbation Earth’s magnetic field and Lorentz force. Comparison between air drag only EDT for orbital decaying time in case of equatorial and near polar orbit are introduced.