Abstract: This paper investigates the ejection of tethered satellite systems. For the deployment of a tethered satellite from an elliptical orbit, the initial angle and velocity of the ejection that minimizes final errors are evaluated, and the dynamic equations are simplified with variable transformation. The method of orbital transfer is utilized for the deployment of a tethered satellite from a circular orbit. The parameters of orbit transformation are given, and the true anomaly is calculated at the end of transforming orbit. In many literatures, it is supposed that the mass of the spacecraft is far greater than that of the payload, so it is further supposed that the center of mass of the tethered satellite system coincides with that of the spacecraft. Those assumptions are only correct for that the spacecraft is massive, such as a space shuttle or a space station. In this paper, there is no assumption that the mass of the spacecraft is far greater than that of the payload, so the conclusions are more general. The numerical simulation results are presented and show that the method is really effective.