Abstract: A method for controlling orbiting spacecraft uploads a plurality of parameters to an onboard processor, the parameters including a plurality of quaternion elements describing an inertial to orbit reference frame, and including radial position, radial velocity and angular rate of the spacecraft. Parameters are input for processing one of a plurality of propagation algorithms, preferably calculating said propagation algorithms with the uploaded parameters, computing enhanced time of day vectors, and computing ground time of day angles from the enhanced time of day vectors for adjusting the payload pointing angle, such as an antenna pointing angle, of the satellite. Preferably, the calculating step comprises integrating an orbital state vector with a second order Runge-Kutta integration algorithm. In addition, enhanced time of day fault processing provides an internal check on integration accuracy and is checked every time of day update.