Package org.orekit.propagation

Propagation

This package provides tools to propagate orbital states with different methods.

Propagation is the prediction of the evolution of an initial state. The initial state and the propagated states are represented in OREKIT by a SpacecraftState, which is a simple container for all needed information at a specific date : mass, kinematics, attitude, date, frame. The state provides basic interpolation features allowing to shift it slightly to close dates. For more accurate and farthest dates, several full-featured propagators are available to propagate the state.

Keplerian propagation

The KeplerianPropagator implements the Propagator interface, which ensures that we can obtain a propagated SpacecraftState at any time once the instance is initialized with an initial state. This extrapolation is not a problem with a simple EquinoctialOrbit representation: only the mean anomaly value changes.

Eckstein-Hechler propagation

This analytical model is suited for near circular orbits and inclination neither equatorial nor critical. It considers J2 to J6 potential coefficients correctors, and uses mean parameters to compute the new position. As the Keplerian propagator, it implements the Propagator interface.

Numerical propagation

It is the most important part of the OREKIT project. Based on Hipparchus integrators, the NumericalPropagator class realizes the interface between space mechanics and mathematical resolutions. If its utilization seems difficult on first sight, it is in fact quite clear and intuitive.

The mathematical problem to integrate is a seven dimension time derivative equations system. The six first equations are given by the Gauss equations (expressed in EquinoctialOrbit) and the seventh is simply the flow rate and mass equation. This first order system is computed by the TimeDerivativesEquations class. It will be instanced by the propagator and then be modified at each step (a fixed t value) by all the needed force models which will add their contribution, the perturbing acceleration.

The integrators provided by Hipparchus need the state vector at t0, the state vector first time derivate at t0, and then calculates the next step state vector, and ask for the next first time derivative, etc. until it reaches the final asked date.

Author:

Luc Maisonobe, Fabien Maussion, Pascal Parraud

Interface Summary

Interface	Description
AdditionalStateProvider	This interface represents providers for additional state data beyond SpacecraftState.
BoundedPropagator	This interface is intended for ephemerides valid only during a time range.
FieldAdditionalStateProvider<T extends RealFieldElement<T>>	This interface represents providers for additional state data beyond SpacecraftState.
FieldBoundedPropagator<T extends RealFieldElement<T>>	This interface is intended for ephemerides valid only during a time range.
FieldPropagator<T extends RealFieldElement<T>>	This interface provides a way to propagate an orbit at any time.
Propagator	This interface provides a way to propagate an orbit at any time.

Class Summary

Class	Description
AbstractPropagator	Common handling of Propagator methods for analytical propagators.
FieldAbstractPropagator<T extends RealFieldElement<T>>	Common handling of Propagator methods for analytical propagators.
FieldSpacecraftState<T extends RealFieldElement<T>>	This class is the representation of a complete state holding orbit, attitude and mass information at a given date.
PropagatorsParallelizer	This class provides a way to propagate simultaneously several orbits.
SpacecraftState	This class is the representation of a complete state holding orbit, attitude and mass information at a given date.