FieldIntelsatElevenElementsPropagator.java
/* Copyright 2002-2024 Airbus Defence and Space
* Licensed to CS GROUP (CS) under one or more
* contributor license agreements. See the NOTICE file distributed with
* this work for additional information regarding copyright ownership.
* Airbus Defence and Space 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.analytical.intelsat;
import java.util.Collections;
import java.util.List;
import org.hipparchus.CalculusFieldElement;
import org.hipparchus.Field;
import org.hipparchus.analysis.differentiation.FieldUnivariateDerivative2;
import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
import org.hipparchus.util.FastMath;
import org.hipparchus.util.FieldSinCos;
import org.orekit.annotation.DefaultDataContext;
import org.orekit.attitudes.AttitudeProvider;
import org.orekit.attitudes.FieldAttitude;
import org.orekit.attitudes.FrameAlignedProvider;
import org.orekit.data.DataContext;
import org.orekit.errors.OrekitException;
import org.orekit.errors.OrekitMessages;
import org.orekit.frames.Frame;
import org.orekit.frames.FramesFactory;
import org.orekit.orbits.FieldCartesianOrbit;
import org.orekit.orbits.FieldOrbit;
import org.orekit.propagation.FieldSpacecraftState;
import org.orekit.propagation.Propagator;
import org.orekit.propagation.analytical.FieldAbstractAnalyticalPropagator;
import org.orekit.time.FieldAbsoluteDate;
import org.orekit.utils.Constants;
import org.orekit.utils.FieldPVCoordinates;
import org.orekit.utils.IERSConventions;
import org.orekit.utils.ParameterDriver;
import org.orekit.utils.units.Unit;
/**
* This class provides elements to propagate Intelsat's 11 elements.
* <p>
* Intelsat's 11 elements propagation is defined in ITU-R S.1525 standard.
* </p>
*
* @author Bryan Cazabonne
* @since 12.1
*/
public class FieldIntelsatElevenElementsPropagator<T extends CalculusFieldElement<T>> extends FieldAbstractAnalyticalPropagator<T> {
/**
* Intelsat's 11 elements.
*/
private final FieldIntelsatElevenElements<T> elements;
/**
* Inertial frame for the output orbit.
*/
private final Frame inertialFrame;
/**
* ECEF frame related to the Intelsat's 11 elements.
*/
private final Frame ecefFrame;
/**
* Spacecraft mass in kilograms.
*/
private final T mass;
/**
* Compute spacecraft's east longitude.
*/
private FieldUnivariateDerivative2<T> eastLongitudeDegrees;
/**
* Compute spacecraft's geocentric latitude.
*/
private FieldUnivariateDerivative2<T> geocentricLatitudeDegrees;
/**
* Compute spacecraft's orbit radius.
*/
private FieldUnivariateDerivative2<T> orbitRadius;
/**
* Default constructor.
* <p>
* This constructor uses the {@link DataContext#getDefault() default data context}.
* </p>
* <p> The attitude provider is set by default to be aligned with the inertial frame.<br>
* The mass is set by default to the
* {@link org.orekit.propagation.Propagator#DEFAULT_MASS DEFAULT_MASS}.<br>
* The inertial frame is set by default to the
* {@link org.orekit.frames.Predefined#TOD_CONVENTIONS_2010_SIMPLE_EOP TOD frame} in the default data
* context.<br>
* The ECEF frame is set by default to the
* {@link org.orekit.frames.Predefined#ITRF_CIO_CONV_2010_SIMPLE_EOP
* CIO/2010-based ITRF simple EOP} in the default data context.
* </p>
*
* @param elements Intelsat's 11 elements
*/
@DefaultDataContext
public FieldIntelsatElevenElementsPropagator(final FieldIntelsatElevenElements<T> elements) {
this(elements, FramesFactory.getTOD(IERSConventions.IERS_2010, true), FramesFactory.getITRF(IERSConventions.IERS_2010, true));
}
/**
* Constructor.
*
* <p> The attitude provider is set by default to be aligned with the inertial frame.<br>
* The mass is set by default to the
* {@link org.orekit.propagation.Propagator#DEFAULT_MASS DEFAULT_MASS}.<br>
* </p>
*
* @param elements Intelsat's 11 elements
* @param inertialFrame inertial frame for the output orbit
* @param ecefFrame ECEF frame related to the Intelsat's 11 elements
*/
public FieldIntelsatElevenElementsPropagator(final FieldIntelsatElevenElements<T> elements, final Frame inertialFrame, final Frame ecefFrame) {
this(elements, inertialFrame, ecefFrame, FrameAlignedProvider.of(inertialFrame), elements.getEpoch().getField().getZero().add(Propagator.DEFAULT_MASS));
}
/**
* Constructor.
*
* @param elements Intelsat's 11 elements
* @param inertialFrame inertial frame for the output orbit
* @param ecefFrame ECEF frame related to the Intelsat's 11 elements
* @param attitudeProvider attitude provider
* @param mass spacecraft mass
*/
public FieldIntelsatElevenElementsPropagator(final FieldIntelsatElevenElements<T> elements, final Frame inertialFrame, final Frame ecefFrame,
final AttitudeProvider attitudeProvider, final T mass) {
super(elements.getEpoch().getField(), attitudeProvider);
this.elements = elements;
this.inertialFrame = inertialFrame;
this.ecefFrame = ecefFrame;
this.mass = mass;
setStartDate(elements.getEpoch());
final FieldOrbit<T> orbitAtElementsDate = propagateOrbit(elements.getEpoch(), getParameters(elements.getEpoch().getField()));
final FieldAttitude<T> attitude = attitudeProvider.getAttitude(orbitAtElementsDate, elements.getEpoch(), inertialFrame);
super.resetInitialState(new FieldSpacecraftState<>(orbitAtElementsDate, attitude, mass));
}
/**
* Converts the Intelsat's 11 elements into Position/Velocity coordinates in ECEF.
*
* @param date computation epoch
* @return Position/Velocity coordinates in ECEF
*/
public FieldPVCoordinates<T> propagateInEcef(final FieldAbsoluteDate<T> date) {
final Field<T> field = date.getField();
final FieldUnivariateDerivative2<T> tDays = new FieldUnivariateDerivative2<>(date.durationFrom(elements.getEpoch()), field.getOne(), field.getZero()).divide(
Constants.JULIAN_DAY);
final T wDegreesPerDay = elements.getLm1().add(IntelsatElevenElements.DRIFT_RATE_SHIFT_DEG_PER_DAY);
final FieldUnivariateDerivative2<T> wt = FastMath.toRadians(tDays.multiply(wDegreesPerDay));
final FieldSinCos<FieldUnivariateDerivative2<T>> scWt = FastMath.sinCos(wt);
final FieldSinCos<FieldUnivariateDerivative2<T>> sc2Wt = FastMath.sinCos(wt.multiply(2.0));
final FieldUnivariateDerivative2<T> satelliteEastLongitudeDegrees = computeSatelliteEastLongitudeDegrees(tDays, scWt, sc2Wt);
final FieldUnivariateDerivative2<T> satelliteGeocentricLatitudeDegrees = computeSatelliteGeocentricLatitudeDegrees(tDays, scWt);
final FieldUnivariateDerivative2<T> satelliteRadius = computeSatelliteRadiusKilometers(wDegreesPerDay, scWt).multiply(Unit.KILOMETRE.getScale());
this.eastLongitudeDegrees = satelliteEastLongitudeDegrees;
this.geocentricLatitudeDegrees = satelliteGeocentricLatitudeDegrees;
this.orbitRadius = satelliteRadius;
final FieldSinCos<FieldUnivariateDerivative2<T>> scLongitude = FastMath.sinCos(FastMath.toRadians(satelliteEastLongitudeDegrees));
final FieldSinCos<FieldUnivariateDerivative2<T>> scLatitude = FastMath.sinCos(FastMath.toRadians(satelliteGeocentricLatitudeDegrees));
final FieldVector3D<FieldUnivariateDerivative2<T>> positionWithDerivatives = new FieldVector3D<>(satelliteRadius.multiply(scLatitude.cos()).multiply(scLongitude.cos()),
satelliteRadius.multiply(scLatitude.cos()).multiply(scLongitude.sin()),
satelliteRadius.multiply(scLatitude.sin()));
return new FieldPVCoordinates<>(new FieldVector3D<>(positionWithDerivatives.getX().getValue(), //
positionWithDerivatives.getY().getValue(), //
positionWithDerivatives.getZ().getValue()), //
new FieldVector3D<>(positionWithDerivatives.getX().getFirstDerivative(), //
positionWithDerivatives.getY().getFirstDerivative(), //
positionWithDerivatives.getZ().getFirstDerivative()), //
new FieldVector3D<>(positionWithDerivatives.getX().getSecondDerivative(), //
positionWithDerivatives.getY().getSecondDerivative(), //
positionWithDerivatives.getZ().getSecondDerivative()));
}
/**
* {@inheritDoc}.
*/
@Override
public void resetInitialState(final FieldSpacecraftState<T> state) {
throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
}
/**
* {@inheritDoc}.
*/
@Override
protected T getMass(final FieldAbsoluteDate<T> date) {
return mass;
}
/**
* {@inheritDoc}.
*/
@Override
protected void resetIntermediateState(final FieldSpacecraftState<T> state, final boolean forward) {
throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
}
/**
* {@inheritDoc}.
*/
@Override
protected FieldOrbit<T> propagateOrbit(final FieldAbsoluteDate<T> date, final T[] parameters) {
return new FieldCartesianOrbit<>(ecefFrame.getTransformTo(inertialFrame, date).transformPVCoordinates(propagateInEcef(date)), inertialFrame, date,
date.getField().getZero().add(Constants.WGS84_EARTH_MU));
}
/**
* Computes the satellite's east longitude.
*
* @param tDays delta time in days
* @param scW sin/cos of the W angle
* @param sc2W sin/cos of the 2xW angle
* @return the satellite's east longitude in degrees
*/
private FieldUnivariateDerivative2<T> computeSatelliteEastLongitudeDegrees(final FieldUnivariateDerivative2<T> tDays, final FieldSinCos<FieldUnivariateDerivative2<T>> scW,
final FieldSinCos<FieldUnivariateDerivative2<T>> sc2W) {
final FieldUnivariateDerivative2<T> longitude = tDays.multiply(tDays).multiply(elements.getLm2()) //
.add(tDays.multiply(elements.getLm1())) //
.add(elements.getLm0());
final FieldUnivariateDerivative2<T> cosineLongitudeTerm = scW.cos().multiply(tDays.multiply(elements.getLonC1()).add(elements.getLonC()));
final FieldUnivariateDerivative2<T> sineLongitudeTerm = scW.sin().multiply(tDays.multiply(elements.getLonS1()).add(elements.getLonS()));
final FieldUnivariateDerivative2<T> latitudeTerm = sc2W.sin()
.multiply(elements.getLatC()
.multiply(elements.getLatC())
.subtract(elements.getLatS().multiply(elements.getLatS()))
.multiply(0.5)) //
.subtract(sc2W.cos().multiply(elements.getLatC().multiply(elements.getLatS()))) //
.multiply(IntelsatElevenElements.K);
return longitude.add(cosineLongitudeTerm).add(sineLongitudeTerm).add(latitudeTerm);
}
/**
* Computes the satellite's geocentric latitude.
*
* @param tDays delta time in days
* @param scW sin/cos of the W angle
* @return he satellite geocentric latitude in degrees
*/
private FieldUnivariateDerivative2<T> computeSatelliteGeocentricLatitudeDegrees(final FieldUnivariateDerivative2<T> tDays,
final FieldSinCos<FieldUnivariateDerivative2<T>> scW) {
final FieldUnivariateDerivative2<T> cosineTerm = scW.cos().multiply(tDays.multiply(elements.getLatC1()).add(elements.getLatC()));
final FieldUnivariateDerivative2<T> sineTerm = scW.sin().multiply(tDays.multiply(elements.getLatS1()).add(elements.getLatS()));
return cosineTerm.add(sineTerm);
}
/**
* Computes the satellite's orbit radius.
*
* @param wDegreesPerDay W angle in degrees/day
* @param scW sin/cos of the W angle
* @return the satellite's orbit radius in kilometers
*/
private FieldUnivariateDerivative2<T> computeSatelliteRadiusKilometers(final T wDegreesPerDay, final FieldSinCos<FieldUnivariateDerivative2<T>> scW) {
final T coefficient = elements.getLm1()
.multiply(2.0)
.divide(wDegreesPerDay.subtract(elements.getLm1()).multiply(3.0))
.negate()
.add(1.0)
.multiply(IntelsatElevenElements.SYNCHRONOUS_RADIUS_KM);
return scW.sin()
.multiply(elements.getLonC().multiply(IntelsatElevenElements.K))
.add(1.0)
.subtract(scW.cos().multiply(elements.getLonS().multiply(IntelsatElevenElements.K)))
.multiply(coefficient);
}
/**
* Get the computed satellite's east longitude.
*
* @return the satellite's east longitude in degrees
*/
public FieldUnivariateDerivative2<T> getEastLongitudeDegrees() {
return eastLongitudeDegrees;
}
/**
* Get the computed satellite's geocentric latitude.
*
* @return the satellite's geocentric latitude in degrees
*/
public FieldUnivariateDerivative2<T> getGeocentricLatitudeDegrees() {
return geocentricLatitudeDegrees;
}
/**
* Get the computed satellite's orbit.
*
* @return satellite's orbit radius in meters
*/
public FieldUnivariateDerivative2<T> getOrbitRadius() {
return orbitRadius;
}
/**
* {@inheritDoc}.
*/
@Override
public Frame getFrame() {
return inertialFrame;
}
/**
* {@inheritDoc}.
*/
@Override
public List<ParameterDriver> getParametersDrivers() {
return Collections.emptyList();
}
}