IntelsatElevenElementsPropagator.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
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*
* http://www.apache.org/licenses/LICENSE-2.0
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package org.orekit.propagation.analytical.intelsat;
import org.hipparchus.analysis.differentiation.UnivariateDerivative2;
import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
import org.hipparchus.geometry.euclidean.threed.Vector3D;
import org.hipparchus.util.FastMath;
import org.hipparchus.util.FieldSinCos;
import org.orekit.annotation.DefaultDataContext;
import org.orekit.attitudes.Attitude;
import org.orekit.attitudes.AttitudeProvider;
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.CartesianOrbit;
import org.orekit.orbits.Orbit;
import org.orekit.propagation.Propagator;
import org.orekit.propagation.SpacecraftState;
import org.orekit.propagation.analytical.AbstractAnalyticalPropagator;
import org.orekit.time.AbsoluteDate;
import org.orekit.utils.Constants;
import org.orekit.utils.IERSConventions;
import org.orekit.utils.PVCoordinates;
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 IntelsatElevenElementsPropagator extends AbstractAnalyticalPropagator {
/**
* Intelsat's 11 elements.
*/
private final IntelsatElevenElements 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 double mass;
/**
* Compute spacecraft's east longitude.
*/
private UnivariateDerivative2 eastLongitudeDegrees;
/**
* Compute spacecraft's geocentric latitude.
*/
private UnivariateDerivative2 geocentricLatitudeDegrees;
/**
* Compute spacecraft's orbit radius.
*/
private UnivariateDerivative2 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 IntelsatElevenElementsPropagator(final IntelsatElevenElements 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 IntelsatElevenElementsPropagator(final IntelsatElevenElements elements, final Frame inertialFrame, final Frame ecefFrame) {
this(elements, inertialFrame, ecefFrame, FrameAlignedProvider.of(inertialFrame), 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 IntelsatElevenElementsPropagator(final IntelsatElevenElements elements, final Frame inertialFrame, final Frame ecefFrame, final AttitudeProvider attitudeProvider,
final double mass) {
super(attitudeProvider);
this.elements = elements;
this.inertialFrame = inertialFrame;
this.ecefFrame = ecefFrame;
this.mass = mass;
setStartDate(elements.getEpoch());
final Orbit orbitAtElementsDate = propagateOrbit(elements.getEpoch());
final Attitude attitude = attitudeProvider.getAttitude(orbitAtElementsDate, elements.getEpoch(), inertialFrame);
super.resetInitialState(new SpacecraftState(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 PVCoordinates propagateInEcef(final AbsoluteDate date) {
final UnivariateDerivative2 tDays = new UnivariateDerivative2(date.durationFrom(elements.getEpoch()), 1.0, 0.0).divide(Constants.JULIAN_DAY);
final double wDegreesPerDay = elements.getLm1() + IntelsatElevenElements.DRIFT_RATE_SHIFT_DEG_PER_DAY;
final UnivariateDerivative2 wt = FastMath.toRadians(tDays.multiply(wDegreesPerDay));
final FieldSinCos<UnivariateDerivative2> scWt = FastMath.sinCos(wt);
final FieldSinCos<UnivariateDerivative2> sc2Wt = FastMath.sinCos(wt.multiply(2.0));
final UnivariateDerivative2 satelliteEastLongitudeDegrees = computeSatelliteEastLongitudeDegrees(tDays, scWt, sc2Wt);
final UnivariateDerivative2 satelliteGeocentricLatitudeDegrees = computeSatelliteGeocentricLatitudeDegrees(tDays, scWt);
final UnivariateDerivative2 satelliteRadius = computeSatelliteRadiusKilometers(wDegreesPerDay, scWt).multiply(Unit.KILOMETRE.getScale());
this.eastLongitudeDegrees = satelliteEastLongitudeDegrees;
this.geocentricLatitudeDegrees = satelliteGeocentricLatitudeDegrees;
this.orbitRadius = satelliteRadius;
final FieldSinCos<UnivariateDerivative2> scLongitude = FastMath.sinCos(FastMath.toRadians(satelliteEastLongitudeDegrees));
final FieldSinCos<UnivariateDerivative2> scLatitude = FastMath.sinCos(FastMath.toRadians(satelliteGeocentricLatitudeDegrees));
final FieldVector3D<UnivariateDerivative2> positionWithDerivatives = new FieldVector3D<>(satelliteRadius.multiply(scLatitude.cos()).multiply(scLongitude.cos()),
satelliteRadius.multiply(scLatitude.cos()).multiply(scLongitude.sin()),
satelliteRadius.multiply(scLatitude.sin()));
return new PVCoordinates(new Vector3D(positionWithDerivatives.getX().getValue(), //
positionWithDerivatives.getY().getValue(), //
positionWithDerivatives.getZ().getValue()), //
new Vector3D(positionWithDerivatives.getX().getFirstDerivative(), //
positionWithDerivatives.getY().getFirstDerivative(), //
positionWithDerivatives.getZ().getFirstDerivative()), //
new Vector3D(positionWithDerivatives.getX().getSecondDerivative(), //
positionWithDerivatives.getY().getSecondDerivative(), //
positionWithDerivatives.getZ().getSecondDerivative()));
}
/**
* {@inheritDoc}.
*/
@Override
public void resetInitialState(final SpacecraftState state) {
throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
}
/**
* {@inheritDoc}.
*/
@Override
protected double getMass(final AbsoluteDate date) {
return mass;
}
/**
* {@inheritDoc}.
*/
@Override
protected void resetIntermediateState(final SpacecraftState state, final boolean forward) {
throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
}
/**
* {@inheritDoc}.
*/
@Override
protected Orbit propagateOrbit(final AbsoluteDate date) {
return new CartesianOrbit(ecefFrame.getTransformTo(inertialFrame, date).transformPVCoordinates(propagateInEcef(date)), inertialFrame, date, 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 UnivariateDerivative2 computeSatelliteEastLongitudeDegrees(final UnivariateDerivative2 tDays, final FieldSinCos<UnivariateDerivative2> scW,
final FieldSinCos<UnivariateDerivative2> sc2W) {
final UnivariateDerivative2 longitude = tDays.multiply(tDays).multiply(elements.getLm2()) //
.add(tDays.multiply(elements.getLm1())) //
.add(elements.getLm0());
final UnivariateDerivative2 cosineLongitudeTerm = scW.cos().multiply(tDays.multiply(elements.getLonC1()).add(elements.getLonC()));
final UnivariateDerivative2 sineLongitudeTerm = scW.sin().multiply(tDays.multiply(elements.getLonS1()).add(elements.getLonS()));
final UnivariateDerivative2 latitudeTerm = sc2W.sin().multiply(0.5 * (elements.getLatC() * elements.getLatC() - elements.getLatS() * elements.getLatS())) //
.subtract(sc2W.cos().multiply(elements.getLatC() * 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 UnivariateDerivative2 computeSatelliteGeocentricLatitudeDegrees(final UnivariateDerivative2 tDays, final FieldSinCos<UnivariateDerivative2> scW) {
final UnivariateDerivative2 cosineTerm = scW.cos().multiply(tDays.multiply(elements.getLatC1()).add(elements.getLatC()));
final UnivariateDerivative2 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 UnivariateDerivative2 computeSatelliteRadiusKilometers(final double wDegreesPerDay, final FieldSinCos<UnivariateDerivative2> scW) {
final double coefficient = IntelsatElevenElements.SYNCHRONOUS_RADIUS_KM * (1.0 - (2.0 * elements.getLm1()) / (3.0 * (wDegreesPerDay - elements.getLm1())));
return scW.sin()
.multiply(IntelsatElevenElements.K * elements.getLonC())
.add(1.0)
.subtract(scW.cos().multiply(IntelsatElevenElements.K * elements.getLonS()))
.multiply(coefficient);
}
/**
* Get the computed satellite's east longitude.
*
* @return the satellite's east longitude in degrees
*/
public UnivariateDerivative2 getEastLongitudeDegrees() {
return eastLongitudeDegrees;
}
/**
* Get the computed satellite's geocentric latitude.
*
* @return the satellite's geocentric latitude in degrees
*/
public UnivariateDerivative2 getGeocentricLatitudeDegrees() {
return geocentricLatitudeDegrees;
}
/**
* Get the computed satellite's orbit.
*
* @return satellite's orbit radius in meters
*/
public UnivariateDerivative2 getOrbitRadius() {
return orbitRadius;
}
/**
* {@inheritDoc}.
*/
@Override
public Frame getFrame() {
return inertialFrame;
}
/**
* Get the Intelsat's 11 elements used by the propagator.
*
* @return the Intelsat's 11 elements used by the propagator
*/
public IntelsatElevenElements getIntelsatElevenElements() {
return elements;
}
}