public interface AdditionalEquations
In some cases users may need to integrate some problemspecific equations along with classical spacecraft equations of motions. One example is optimal control in low thrust where adjoint parameters linked to the minimized Hamiltonian must be integrated. Another example is formation flying or rendezvous which use the ClohessyWhiltshire equations for the relative motion.
This interface allows users to add such equations to a numerical propagator
. Users provide the
equations as an implementation of this interface and register it to the propagator thanks to
its AbstractIntegratedPropagator.addAdditionalEquations(AdditionalEquations)
method. Several such objects can be registered with each numerical propagator, but it is
recommended to gather in the same object the sets of parameters which equations can interact
on each others states.
The additional parameters are gathered in a simple p array. The additional equations compute
the pDot array, which is the timederivative of the p array. Since the additional parameters
p may also have an influence on the equations of motion themselves that should be accumulated
to the main state derivatives (for example an equation linked to a complex thrust model may
induce an acceleration and a mass change), the computeDerivatives
method can return a double array that will be
added to the main state derivatives. This means these equations can be used as an
additional force model if needed. If the additional parameters have no influence at all on
the main spacecraft state, a null reference may be returned.
This interface is the numerical (read not already integrated) counterpart of
the AdditionalStateProvider
interface.
It allows to append various additional state parameters to any numerical propagator
.
AbstractIntegratedPropagator
,
AdditionalStateProvider
Modifier and Type  Method and Description 

double[] 
computeDerivatives(SpacecraftState s,
double[] pDot)
Compute the derivatives related to the additional state parameters.

String 
getName()
Get the name of the additional state.

String getName()
double[] computeDerivatives(SpacecraftState s, double[] pDot) throws OrekitException
When this method is called, the spacecraft state contains the main
state (orbit, attitude and mass), all the states provided through
the additional
state providers
registered to the propagator, and the additional state
integrated using this equation. It does not contains any other
states to be integrated alongside during the same propagation.
s
 current state information: date, kinematics, attitude, and
additional statepDot
 placeholder where the derivatives of the additional parameters
should be putOrekitException
 if some specific error occursCopyright © 20022015 CS Systèmes d'information. All rights reserved.