In this project, an algorithm has been proposed to detect a satellite’s maneuver by comparingthe orbital elements observed from the two line element data and the orbital elements propagatedwith the help of Simplified perturbations models. A set of TLE data for an object orbiting Earthcontains a specific set of orbital elements. Simplified perturbation are utilized to propagate theorbital velocity and position vector of the same object. By comparing the results obtained fromboth the methods, the maneuvers of a satellite are detected. This project outlines the workingmethodology and the implementation of the algorithm developed to detect the maneuvers. Thefunctioning of the technique is assessed with reference to two case studies for which the maneuverhistory is available by following the approach employed by Kelecy et al. (2007). The same methodis implemented to detect the orbit controlling maneuvers as well as the fine control maneuvers. Theresults derived from the analysis indicate that the maneuvers which has the magnitude of even aslow as cm/s has been detected when the detection parameters are calibrated properly.

The number of man-made space objects is dramatically growing nowadays. The continuous monitoring and studying of these objects are necessary to keep their number under control and ensure safe space operations. With respect thereto, international guidelines recommend decongesting the most populated space regions from satellites arriving at the end of their operational lifetime by performing post-mission disposal strategies. In general, a satellite is considered to be functional if it is still performing periodic manoeuvres to stay within the orbital operation configuration. This study presents a promising method to detect historical in-plane manoeuvrers of satellites on a geostationary orbit (GEO). Since a manoeuvrer changes the orbital state of the spacecraft, its effect can be detected by comparing the observed data to a reference evolution. In this case, the  model is represented by the dynamical model STELA  based on a semi-analytical theory. The observed data is provided by the public American space object catalogue. The Two-line element (TLE) database contains the orbital state of each tracked object, however, not all six orbital parameters are interesting to study in terms of in- plane manoeuvrers. The evolution of the longitude and of the eccentricity vector is immediately affected by a manoeuvre that changes the shape or the size of an orbit. Within the longitude analysis, the manoeuvre epoch is estimated by focusing on the manoeuvre strategy. An operational spacecraft usually performs a manoeuvre as soon as the longitude motion threatens to violate the operational deadband. Consequently, the longitude evolution follows a parabolic motion. Two polynomial curves of second degree are laid over the observation: the first curve is derived from a simplified dynamical model and the second curve is directly obtained through a Least Squares (LS) fitting method. The discrepancy between the LS and physical fitted parabolas gives an indication on the quality of the input data, that is to say, of the TLEs. The detected manoeuvre epoch must be companioned by a confidential parameter that denotes the time range around the estimated epoch in which the manoeuvre is expected to have happened. The manoeuvre interval is then forwarded to the eccentricity analysis where the manoeuvrer epoch is estimated more precisely by studying the divergence between the observed and expected eccentricity vector evolution. The latter is propagated with STELA after having estimated the area-to-mass ratio that is needed in order to model the perturbation effects accurately upon which the performance of the dynamical reference model strongly depends. As soon as the observed eccentricity vector deviates significantly from the expected evolution, the epoch and the velocity ΔV of the manoeuvre can be recovered, too.