Vehicle Satelltie Orbit, Earth Orientation Parameters


I am simulating the trajectory of my vehicle as “Earth-Orbiting Spacecraft”, e.g. with Kepler parameters. However, I have some problems replicating the Kepler orbit from the ECEF coordinates and velocities defined in the receiver_antenna.csv.
Transformation from ECEF to ECI and computing the kepler elements yield oscullating elements in my case. Does Skydel model more than a two-body orbit when “Earth-Orbiting Spacecraft” is selcted? If no, what conventions do you follow to transform ECEF to ECI coordaintes and what is the source of the default Earth Orientation Parameters?

Best regards

The trajectory is based on the two-body propagator and doesn’t include any perturbations.The mismatch between Keplerian parameters obtained from the state vector (in ECI) and the Keplerian parameters used for orbit definition can be caused by:

  • Our two-model propagator implementation. I believe we have some simplification in the propagation model as our primary goal was to estimate the accuracy of the GNSS receiver relative to simulated trajectory expressed in ECEF. For more accurate models, we suggest importing the satellite’s orbit expressed in cartesian coordinates defined in ECEF or ECI reference frame.
  • In general, by differences in transformation convention from ECEF to ECI.
  • Approximation in velocity computation. In ideal case, the state vector should include instantaneous velocity at time t0 of transformation.