AdditionalEquations.java
/* Copyright 2010-2011 Centre National d'Études Spatiales
* Licensed to CS Systèmes d'Information (CS) under one or more
* contributor license agreements. See the NOTICE file distributed with
* this work for additional information regarding copyright ownership.
* CS licenses this file to You under the Apache License, Version 2.0
* (the "License"); you may not use this file except in compliance with
* the License. You may obtain a copy of the License at
*
* http://www.apache.org/licenses/LICENSE-2.0
*
* Unless required by applicable law or agreed to in writing, software
* distributed under the License is distributed on an "AS IS" BASIS,
* WITHOUT WARRANTIES OR CONDITIONS OF ANY KIND, either express or implied.
* See the License for the specific language governing permissions and
* limitations under the License.
*/
package org.orekit.propagation.integration;
import org.orekit.propagation.SpacecraftState;
import org.orekit.time.AbsoluteDate;
/** This interface allows users to add their own differential equations to a numerical propagator.
*
* <p>
* In some cases users may need to integrate some problem-specific 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 rendez-vous which use the Clohessy-Whiltshire
* equations for the relative motion.
* </p>
* <p>
* This interface allows users to add such equations to a {@link
* org.orekit.propagation.numerical.NumericalPropagator numerical propagator}. Users provide the
* equations as an implementation of this interface and register it to the propagator thanks to
* its {@link org.orekit.propagation.numerical.NumericalPropagator#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.
* </p>
* <p>
* The additional parameters are gathered in a simple p array. The additional equations compute
* the pDot array, which is the time-derivative 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 {@link #computeDerivatives(SpacecraftState, double[])
* computeDerivatives} method can return a double array that will be
* <em>added</em> 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.
* </p>
* <p>
* This interface is the numerical (read not already integrated) counterpart of
* the {@link org.orekit.propagation.AdditionalStateProvider} interface.
* It allows to append various additional state parameters to any {@link
* org.orekit.propagation.numerical.NumericalPropagator numerical propagator}.
* </p>
* @see AbstractIntegratedPropagator
* @see org.orekit.propagation.AdditionalStateProvider
* @author Luc Maisonobe
*/
public interface AdditionalEquations {
/** Get the name of the additional state.
* @return name of the additional state
*/
String getName();
/**
* Initialize the equations at the start of propagation.
*
* <p>
* This method will be called once at propagation start,
* before any calls to {@link #computeDerivatives(SpacecraftState, double[])}.
* </p>
*
* <p>
* The default implementation of this method does nothing.
* </p>
*
* @param initialState initial state information at the start of propagation.
* @param target date of propagation. Not equal to {@code
* initialState.getDate()}.
*/
default void init(final SpacecraftState initialState, final AbsoluteDate target) {
// nothing by default
}
/** Compute the derivatives related to the additional state parameters.
* <p>
* When this method is called, the spacecraft state contains the main
* state (orbit, attitude and mass), all the states provided through
* the {@link org.orekit.propagation.AdditionalStateProvider additional
* state providers} registered to the propagator, and the additional state
* integrated using this equation. It does <em>not</em> contains any other
* states to be integrated alongside during the same propagation.
* </p>
* @param s current state information: date, kinematics, attitude, and
* additional state
* @param pDot placeholder where the derivatives of the additional parameters
* should be put
* @return cumulative effect of the equations on the main state (may be null if
* equations do not change main state at all)
*/
double[] computeDerivatives(SpacecraftState s, double[] pDot);
}