org.orekit.propagation.numerical.cr3bp

Class STMEquations

java.lang.Object

org.orekit.propagation.numerical.cr3bp.STMEquations

All Implemented Interfaces:

AdditionalEquations

public class STMEquations
extends Object
implements AdditionalEquations

Class calculating the state transition matrix coefficient for CR3BP Computation.

Since:

10.2

Author:

Vincent Mouraux

See Also:

"Dynamical systems, the three-body problem, and space mission design, Koon, Lo, Marsden, Ross"

Constructor Summary

Constructors

Constructor and Description
STMEquations(CR3BPSystem syst) Simple constructor.

Method Summary

Modifier and Type	Method and Description
double[]	computeDerivatives(SpacecraftState s, double[] dPhi) Compute the derivatives related to the additional state parameters.
String	getName() Get the name of the additional state.
RealMatrix	getStateTransitionMatrix(SpacecraftState s) Method returning the State Transition Matrix.
SpacecraftState	setInitialPhi(SpacecraftState s) Method adding the standard initial values of the additional state to the initial spacecraft state.

Methods inherited from class java.lang.Object

clone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, wait

Methods inherited from interface org.orekit.propagation.integration.AdditionalEquations

init

Constructor Detail

STMEquations

public STMEquations(CR3BPSystem syst)

Simple constructor.

Parameters:

syst - CR3BP System considered

Method Detail

setInitialPhi

public SpacecraftState setInitialPhi(SpacecraftState s)

Method adding the standard initial values of the additional state to the initial spacecraft state.

Parameters:

s - Initial state of the system

Returns:

s Initial augmented (with the additional equations) state

computeDerivatives

public double[] computeDerivatives(SpacecraftState s,
                                   double[] dPhi)

Compute the derivatives related to the additional state parameters.

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.

Specified by:

computeDerivatives in interface AdditionalEquations

Parameters:

s - current state information: date, kinematics, attitude, and additional state

dPhi - placeholder where the derivatives of the additional parameters should be put

Returns:

cumulative effect of the equations on the main state (may be null if equations do not change main state at all)

getName

public String getName()

Get the name of the additional state.

Specified by:

getName in interface AdditionalEquations

Returns:

name of the additional state

getStateTransitionMatrix

public RealMatrix getStateTransitionMatrix(SpacecraftState s)

Method returning the State Transition Matrix.

Parameters:

s - SpacecraftState of the system

Returns:

phiM State Transition Matrix