KeplerianPropagator.java
/* Copyright 2002-2020 CS Group
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* 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,
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package org.orekit.propagation.analytical;
import java.util.Collections;
import java.util.Map;
import org.orekit.annotation.DefaultDataContext;
import org.orekit.attitudes.Attitude;
import org.orekit.attitudes.AttitudeProvider;
import org.orekit.data.DataContext;
import org.orekit.orbits.Orbit;
import org.orekit.orbits.OrbitType;
import org.orekit.orbits.PositionAngle;
import org.orekit.propagation.Propagator;
import org.orekit.propagation.SpacecraftState;
import org.orekit.time.AbsoluteDate;
import org.orekit.utils.TimeSpanMap;
/** Simple Keplerian orbit propagator.
* @see Orbit
* @author Guylaine Prat
*/
public class KeplerianPropagator extends AbstractAnalyticalPropagator {
/** Initial state. */
private SpacecraftState initialState;
/** All states. */
private TimeSpanMap<SpacecraftState> states;
/** Build a propagator from orbit only.
* <p>The central attraction coefficient μ is set to the same value used
* for the initial orbit definition. Mass and attitude provider are set to
* unspecified non-null arbitrary values.</p>
*
* <p>This constructor uses the {@link DataContext#getDefault() default data context}.
*
* @param initialOrbit initial orbit
* @see #KeplerianPropagator(Orbit, AttitudeProvider)
*/
@DefaultDataContext
public KeplerianPropagator(final Orbit initialOrbit) {
this(initialOrbit, Propagator.getDefaultLaw(DataContext.getDefault().getFrames()),
initialOrbit.getMu(), DEFAULT_MASS);
}
/** Build a propagator from orbit and central attraction coefficient μ.
* <p>Mass and attitude provider are set to unspecified non-null arbitrary values.</p>
*
* <p>This constructor uses the {@link DataContext#getDefault() default data context}.
*
* @param initialOrbit initial orbit
* @param mu central attraction coefficient (m³/s²)
* @see #KeplerianPropagator(Orbit, AttitudeProvider, double)
*/
@DefaultDataContext
public KeplerianPropagator(final Orbit initialOrbit, final double mu) {
this(initialOrbit, Propagator.getDefaultLaw(DataContext.getDefault().getFrames()),
mu, DEFAULT_MASS);
}
/** Build a propagator from orbit and attitude provider.
* <p>The central attraction coefficient μ is set to the same value
* used for the initial orbit definition. Mass is set to an unspecified
* non-null arbitrary value.</p>
* @param initialOrbit initial orbit
* @param attitudeProv attitude provider
*/
public KeplerianPropagator(final Orbit initialOrbit,
final AttitudeProvider attitudeProv) {
this(initialOrbit, attitudeProv, initialOrbit.getMu(), DEFAULT_MASS);
}
/** Build a propagator from orbit, attitude provider and central attraction
* coefficient μ.
* <p>Mass is set to an unspecified non-null arbitrary value.</p>
* @param initialOrbit initial orbit
* @param attitudeProv attitude provider
* @param mu central attraction coefficient (m³/s²)
*/
public KeplerianPropagator(final Orbit initialOrbit,
final AttitudeProvider attitudeProv,
final double mu) {
this(initialOrbit, attitudeProv, mu, DEFAULT_MASS);
}
/** Build propagator from orbit, attitude provider, central attraction
* coefficient μ and mass.
* @param initialOrbit initial orbit
* @param attitudeProv attitude provider
* @param mu central attraction coefficient (m³/s²)
* @param mass spacecraft mass (kg)
*/
public KeplerianPropagator(final Orbit initialOrbit, final AttitudeProvider attitudeProv,
final double mu, final double mass) {
super(attitudeProv);
// ensure the orbit use the specified mu and has no non-Keplerian derivatives
initialState = fixState(initialOrbit,
getAttitudeProvider().getAttitude(initialOrbit,
initialOrbit.getDate(),
initialOrbit.getFrame()),
mass, mu, Collections.emptyMap());
states = new TimeSpanMap<SpacecraftState>(initialState);
super.resetInitialState(initialState);
}
/** Fix state to use a specified mu and remove derivatives.
* <p>
* This ensures the propagation model (which is based on calling
* {@link Orbit#shiftedBy(double)}) is Keplerian only and uses a specified mu.
* </p>
* @param orbit orbit to fix
* @param attitude current attitude
* @param mass current mass
* @param mu gravity coefficient to use
* @param additionalStates additional states
* @return fixed orbit
*/
private SpacecraftState fixState(final Orbit orbit, final Attitude attitude, final double mass,
final double mu, final Map<String, double[]> additionalStates) {
final OrbitType type = orbit.getType();
final double[] stateVector = new double[6];
type.mapOrbitToArray(orbit, PositionAngle.TRUE, stateVector, null);
final Orbit fixedOrbit = type.mapArrayToOrbit(stateVector, null, PositionAngle.TRUE,
orbit.getDate(), mu, orbit.getFrame());
SpacecraftState fixedState = new SpacecraftState(fixedOrbit, attitude, mass);
for (final Map.Entry<String, double[]> entry : additionalStates.entrySet()) {
fixedState = fixedState.addAdditionalState(entry.getKey(), entry.getValue());
}
return fixedState;
}
/** {@inheritDoc} */
public void resetInitialState(final SpacecraftState state) {
// ensure the orbit use the specified mu and has no non-Keplerian derivatives
final double mu = initialState == null ? state.getMu() : initialState.getMu();
final SpacecraftState fixedState = fixState(state.getOrbit(),
state.getAttitude(),
state.getMass(),
mu,
state.getAdditionalStates());
initialState = fixedState;
states = new TimeSpanMap<SpacecraftState>(initialState);
super.resetInitialState(fixedState);
}
/** {@inheritDoc} */
protected void resetIntermediateState(final SpacecraftState state, final boolean forward) {
if (forward) {
states.addValidAfter(state, state.getDate());
} else {
states.addValidBefore(state, state.getDate());
}
stateChanged(state);
}
/** {@inheritDoc} */
protected Orbit propagateOrbit(final AbsoluteDate date) {
// propagate orbit
Orbit orbit = states.get(date).getOrbit();
do {
// we use a loop here to compensate for very small date shifts error
// that occur with long propagation time
orbit = orbit.shiftedBy(date.durationFrom(orbit.getDate()));
} while (!date.equals(orbit.getDate()));
return orbit;
}
/** {@inheritDoc}*/
protected double getMass(final AbsoluteDate date) {
return states.get(date).getMass();
}
}