/* Copyright 2002-2026 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>
   *
   * @param <T> type of the field elements
   * @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).withMass(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
     public 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();
     }
 
     /**
      * Get the Intelsat's 11 elements used by the propagator.
      *
      * @return the Intelsat's 11 elements used by the propagator
      */
     public FieldIntelsatElevenElements<T> getIntelsatElevenElements() {
         return elements;
     }
 }