SimulateOrbit

This program integrates an orbit from a given force function (dynamic orbit). The force functions are given by forces. For computation of non-conservative forces a satelliteModel is needed. The integration method must be selected with propagator. Because the orbit data are calculated in the celestial reference frame (CRF) you need earthRotation to transform the force function from the terrestrial reference frame (TRF). The integration start and end time, as well as the sampling, are derived from the timeSeries option. It is possible to integrate the arc in reverse, where the initial conditions are assumed to be met at the end time of the timeSeries.

Name	Type	Annotation
outputfileOrbit	filename	orbit file to be written.
inputfileSatelliteModel	filename	satellite macro model
timeSeries	timeSeries	time points for simulated orbit epochs.
integrationConstants	choice
kepler	sequence
majorAxis	double	[m]
eccentricity	double	[-]
inclination	angle	[degree]
ascendingNode	angle	[degree]
argumentOfPerigee	angle	[degree]
meanAnomaly	angle	[degree]
GM	double	Geocentric gravitational constant
positionAndVelocity	sequence
position0x	double	[m] in CRF
position0y	double	[m] in CRF
position0z	double	[m] in CRF
velocity0x	double	[m/s]
velocity0y	double	[m/s]
velocity0z	double	[m/s]
file	sequence
inputfileOrbit	filename	only epoch at timeStart is used
margin	double	[seconds] used when finding initial epoch in orbitFile
propagator	orbitPropagator	orbit propagation method.
earthRotation	earthRotation
ephemerides	ephemerides
forces	forces	considered in orbit propagation.
reverse	boolean	start integration at last epoch in timeSeries, going backward in time.