Satellite Orbit and Ephemeris Determination using Inter Satellite Links

Software DescriptionInter Satellite LinksThe following sections contain brief a description of the implementation and functionality ofthe main software components. Additionally, equations for some "remaining" topics likemeasurement errors and co-ordinate transformation are given.5.1 Orbit IntegrationThe orbit integrator has to compute the forces acting on the satellites and conduct a numericalintegration. The forces are fixed in different co-ordinate frames vary with time in an other.The main force, earth's gravity is fixed with respect to the terrestrial frame, whereas thirdbody attraction and solar radiation depend on the ephemeris of celestial body which can beexpressed easier in inertial co-ordinates.The computations are therefore performed in the inertial frame ECI-J2000. The accelerationof the rotating earth gravity field has therefore to be converted into inertial referencedacceleration for each computation epoch. The transformation matrix from the terrestrial frameto the inertial frame consists of four elements, sidereal angle (hour angle), precession,nutation and polar motion. Only the first three can be computed, although with somecomputational effort, directly. Polar motion , as well as the true length of day, has a randomcomponent and is predicted by the IERS (Bulletin A) and updated from measurements.Normally these earth rotation parameters are estimated within the orbit determination process.The software however does not account for polar motion and true length of day up to now, butimplementation is planned for the near future.The following figure shows the flow chart of an orbit propagator. Starting from a satelliteposition and velocity at a given time, the contributing forces are computed sequentially andintegrated numerically to derive the state at the next epoch. This process is repeated, thus atime series of satellite states is generated.Page 72R. Wolf