| dc.description.abstract | The orbital state of the satellite or celestial body is defined as its position and velocity vectors or orbit elements during a specific epoch. After knowing the initial state of the satellite, we can use an orbit propagator to propagate the future ephemeris of the satellite, which makes it an important function inside the navigation filter of an artificial satellite. The algorithm of an orbit propagator should consider not only the simplified Keplerian problem but also other forces in the real space environment: e.g. the aspherical gravitational forces from the Earth, drag, third-body effect, solar radiation pressure, etc. The non-Keplerian forces above are called orbit perturbations, which will make the real trajectory of the satellite differ from the ideal Keplerian orbit.
One should consider the accuracy and computational efficiency when it comes to the choice and design of an orbit propagator, especially if intended for onboard use. The effect of orbit perturbations can be simulated more accurately using a numerical integration approach, but will cost more computationally compared to the analytical solution. This thesis presents and examines the sensitivities of our self-made, MATLAB-based UPOP (UPperair Orbit Propagator) orbit propagator: from the tasks of coordinate transformations, integrator tolerance, the implementations of force models, and finally compares propagated trajectories with the raw flight data of FORMOSAT-5. Our goal is to create an orbit propagator which has the propagational error less than 10 km after 7-days of propagation without orbit determination data inputs. | en_US |