TLE and OMM support

In MDS 1.4.0, full TLE support is provided. Example to load a TLE are present and a currently experimental functionality to generate TLE data has been added.

Method of importing TLE and OMM

Import is done using SGP4.Net library for C#. The tool is used to load TLE data, propagate it to the current epoch using SGP4, which is then rotated from TEME to J2000 and further propagated using MDS propagator.

MDS fully supports direct TLE input as strings in the API; as well as requesting TLE's from Spacetrack and Celestrak by satellite ID's. The propagation can both be done by the internal propagator or MDS' integrated SGP4. An example of TLE propagation precision can be found in example_load_iss_tle.py. ISS.

Figure: TLE Received and loaded into MDS with a custom ISS model.

TLE generation and export

MDS also supports outputting TLE's if this is required to interface for ground station link scheduling. Due to the nature of the mean vs osculating elements, the precision of the output TLE is rather varied. Some fields require an initial TLE of the satellite, while others are optimized by minimising residuals between the current orbit and the generated TLE's orbit. TLE generation is done as follows:

  • Add or import from another TLE a spacecraft or multiple spacecrafts.
  • Allow the body to propagate, multiple days are recommended.
  • The earliest available and valid position/velocity from flight record of spacecraft and also current position and velocity are then used for the optimization.
  • The current osculating elements of the spacecraft from position and velocity are computed.
  • The mean elements are computed so that spacecraft trajectory, computed by SGP4 from TLE would match the propagated trajectory in MDS as closely as possible. The previously calculated osculating elements act as the initial guess. The list of elements is then computed:
    • mean motion
    • inclination
    • eccentricity
    • RAAN
    • Argument of perigee
    • Mean anomaly
  • the b* and mean motion derivatives are either left as 0 or, if the spacecraft was imported from a TLE, then they are copied from the original TLE.
  • Once the fit is complete, the data is returned as TLE or OMM.

Example TLE Data Generation Precision

Below is a table of the discrepancy between the original orbit in [km] and the orbits resulting from the exported TLEs:

Time [h] Circular Orbit High Inclination Highly Elliptical Retrograde
0.0000000 9.7978 0.3461 3.6519 7.6972
113.8500000 21.1207 3.8234 42.1615 38.9155
121.8700000 18.9673 4.0041 46.2163 25.7515