FieldKeplerianPropagator.java
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* this work for additional information regarding copyright ownership.
* CS licenses this file to You under the Apache License, Version 2.0
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*
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* Unless required by applicable law or agreed to in writing, software
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package org.orekit.propagation.analytical;
import java.util.Collections;
import java.util.List;
import org.hipparchus.CalculusFieldElement;
import org.hipparchus.util.MathArrays;
import org.orekit.attitudes.AttitudeProvider;
import org.orekit.attitudes.FieldAttitude;
import org.orekit.attitudes.InertialProvider;
import org.orekit.orbits.FieldOrbit;
import org.orekit.orbits.Orbit;
import org.orekit.orbits.OrbitType;
import org.orekit.orbits.PositionAngle;
import org.orekit.propagation.FieldSpacecraftState;
import org.orekit.time.FieldAbsoluteDate;
import org.orekit.utils.FieldArrayDictionary;
import org.orekit.utils.FieldTimeSpanMap;
import org.orekit.utils.ParameterDriver;
/** Simple Keplerian orbit propagator.
* @see FieldOrbit
* @author Guylaine Prat
*/
public class FieldKeplerianPropagator<T extends CalculusFieldElement<T>> extends FieldAbstractAnalyticalPropagator<T> {
/** All states. */
private transient FieldTimeSpanMap<FieldSpacecraftState<T>, T> 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>
*
* @param initialFieldOrbit initial orbit
* @see #FieldKeplerianPropagator(FieldOrbit, AttitudeProvider)
*/
public FieldKeplerianPropagator(final FieldOrbit<T> initialFieldOrbit) {
this(initialFieldOrbit, InertialProvider.of(initialFieldOrbit.getFrame()),
initialFieldOrbit.getMu(), initialFieldOrbit.getA().getField().getZero().add(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>
*
* @param initialFieldOrbit initial orbit
* @param mu central attraction coefficient (m³/s²)
* @see #FieldKeplerianPropagator(FieldOrbit, AttitudeProvider, CalculusFieldElement)
*/
public FieldKeplerianPropagator(final FieldOrbit<T> initialFieldOrbit, final T mu) {
this(initialFieldOrbit, InertialProvider.of(initialFieldOrbit.getFrame()),
mu, initialFieldOrbit.getA().getField().getZero().add(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 initialFieldOrbit initial orbit
* @param attitudeProv attitude provider
*/
public FieldKeplerianPropagator(final FieldOrbit<T> initialFieldOrbit,
final AttitudeProvider attitudeProv) {
this(initialFieldOrbit, attitudeProv, initialFieldOrbit.getMu(), initialFieldOrbit.getA().getField().getZero().add(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 initialFieldOrbit initial orbit
* @param attitudeProv attitude provider
* @param mu central attraction coefficient (m³/s²)
*/
public FieldKeplerianPropagator(final FieldOrbit<T> initialFieldOrbit,
final AttitudeProvider attitudeProv,
final T mu) {
this(initialFieldOrbit, attitudeProv, mu, initialFieldOrbit.getA().getField().getZero().add(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 FieldKeplerianPropagator(final FieldOrbit<T> initialOrbit, final AttitudeProvider attitudeProv,
final T mu, final T mass) {
super(initialOrbit.getA().getField(), attitudeProv);
// ensure the orbit use the specified mu and has no non-Keplerian derivatives
final FieldSpacecraftState<T> initial = fixState(initialOrbit,
getAttitudeProvider().getAttitude(initialOrbit,
initialOrbit.getDate(),
initialOrbit.getFrame()),
mass, mu, null, null);
states = new FieldTimeSpanMap<>(initial, initialOrbit.getA().getField());
super.resetInitialState(initial);
}
/** 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 (may be null)
* @param additionalStatesderivatives additional states derivatives (may be null)
* @return fixed orbit
*/
private FieldSpacecraftState<T> fixState(final FieldOrbit<T> orbit, final FieldAttitude<T> attitude, final T mass, final T mu,
final FieldArrayDictionary<T> additionalStates,
final FieldArrayDictionary<T> additionalStatesderivatives) {
final OrbitType type = orbit.getType();
final T[] stateVector = MathArrays.buildArray(mass.getField(), 6);
type.mapOrbitToArray(orbit, PositionAngle.TRUE, stateVector, null);
final FieldOrbit<T> fixedOrbit = type.mapArrayToOrbit(stateVector, null, PositionAngle.TRUE,
orbit.getDate(), mu, orbit.getFrame());
FieldSpacecraftState<T> fixedState = new FieldSpacecraftState<>(fixedOrbit, attitude, mass);
if (additionalStates != null) {
for (final FieldArrayDictionary<T>.Entry entry : additionalStates.getData()) {
fixedState = fixedState.addAdditionalState(entry.getKey(), entry.getValue());
}
}
if (additionalStatesderivatives != null) {
for (final FieldArrayDictionary<T>.Entry entry : additionalStatesderivatives.getData()) {
fixedState = fixedState.addAdditionalStateDerivative(entry.getKey(), entry.getValue());
}
}
return fixedState;
}
/** {@inheritDoc} */
public void resetInitialState(final FieldSpacecraftState<T> state) {
// ensure the orbit use the specified mu and has no non-Keplerian derivatives
final FieldSpacecraftState<T> formerInitial = getInitialState();
final T mu = formerInitial == null ? state.getMu() : formerInitial.getMu();
final FieldSpacecraftState<T> fixedState = fixState(state.getOrbit(),
state.getAttitude(),
state.getMass(),
mu,
state.getAdditionalStatesValues(),
state.getAdditionalStatesDerivatives());
states = new FieldTimeSpanMap<>(fixedState, state.getDate().getField());
super.resetInitialState(fixedState);
}
/** {@inheritDoc} */
protected void resetIntermediateState(final FieldSpacecraftState<T> state, final boolean forward) {
if (forward) {
states.addValidAfter(state, state.getDate());
} else {
states.addValidBefore(state, state.getDate());
}
stateChanged(state);
}
/** {@inheritDoc} */
protected FieldOrbit<T> propagateOrbit(final FieldAbsoluteDate<T> date, final T[] parameters) {
// propagate orbit
FieldOrbit<T> 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 T getMass(final FieldAbsoluteDate<T> date) {
return states.get(date).getMass();
}
/** {@inheritDoc} */
@Override
protected List<ParameterDriver> getParametersDrivers() {
// Keplerian propagation model does not have parameter drivers.
return Collections.emptyList();
}
}