propagate_keplerian
Propagate satellite orbits using two-body Keplerian dynamics. Accept initial ECI Cartesian state or Keplerian elements, then propagate to target epoch or over a time range.
Instructions
Propagate a satellite orbit using two-body (Keplerian) analytical dynamics.
Initialize with either an ECI Cartesian state vector or Keplerian elements. Provide either target_epoch for a single point, or start_epoch + end_epoch for a range.
Args: epoch: Initial epoch (ISO string). state_eci: ECI Cartesian state [x,y,z,vx,vy,vz] in meters and m/s. elements_koe: Keplerian elements [a,e,i,RAAN,omega,M] (a in meters, angles in input_angle_format). input_angle_format: Angle format for input KOE elements ("degrees" or "radians"). target_epoch: Single target epoch (ISO string). start_epoch: Range start epoch (ISO string). end_epoch: Range end epoch (ISO string). step_seconds: Step size in seconds for range propagation (default 60). output_frame: Output coordinate frame (eci, ecef, gcrf, itrf, eme2000, koe_osc, koe_mean). angle_format: Angle format for KOE output ("degrees" or "radians").
Input Schema
| Name | Required | Description | Default |
|---|---|---|---|
| epoch | Yes | ||
| end_epoch | No | ||
| state_eci | No | ||
| start_epoch | No | ||
| angle_format | No | degrees | |
| elements_koe | No | ||
| output_frame | No | eci | |
| step_seconds | No | ||
| target_epoch | No | ||
| input_angle_format | No | degrees |