/* Copyright 2002-2025 CS GROUP
    * 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.
    * 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,
   * 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.gnss;
  
  import org.hipparchus.analysis.differentiation.UnivariateDerivative2;
  import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
  import org.hipparchus.geometry.euclidean.threed.Vector3D;
  import org.orekit.attitudes.Attitude;
  import org.orekit.attitudes.AttitudeProvider;
  import org.orekit.errors.OrekitException;
  import org.orekit.errors.OrekitMessages;
  import org.orekit.frames.Frame;
  import org.orekit.orbits.CartesianOrbit;
  import org.orekit.orbits.Orbit;
  import org.orekit.propagation.SpacecraftState;
  import org.orekit.propagation.analytical.AbstractAnalyticalPropagator;
  import org.orekit.propagation.analytical.gnss.data.SBASOrbitalElements;
  import org.orekit.time.AbsoluteDate;
  import org.orekit.utils.PVCoordinates;
  
  /**
   * This class aims at propagating a SBAS orbit from {@link SBASOrbitalElements}.
   *
   * @see "Tyler Reid, Todd Walker, Per Enge, L1/L5 SBAS MOPS Ephemeris Message to
   *       Support Multiple Orbit Classes, ION ITM, 2013"
   *
   * @author Bryan Cazabonne
   * @since 10.1
   *
   */
  public class SBASPropagator extends AbstractAnalyticalPropagator {
  
      /** The SBAS orbital elements used. */
      private final SBASOrbitalElements sbasOrbit;
  
      /** The spacecraft mass (kg). */
      private final double mass;
  
      /** The Earth gravity coefficient used for SBAS propagation. */
      private final double mu;
  
      /** The ECI frame used for SBAS propagation. */
      private final Frame eci;
  
      /** The ECEF frame used for SBAS propagation. */
      private final Frame ecef;
  
      /**
       * Private constructor.
       * @param sbasOrbit Glonass orbital elements
       * @param eci Earth Centered Inertial frame
       * @param ecef Earth Centered Earth Fixed frame
       * @param provider Attitude provider
       * @param mass Satellite mass (kg)
       * @param mu Earth's gravity coefficient used for SBAS propagation
       */
      SBASPropagator(final SBASOrbitalElements sbasOrbit, final Frame eci,
                     final Frame ecef, final AttitudeProvider provider,
                     final double mass, final double mu) {
          super(provider);
          // Stores the SBAS orbital elements
          this.sbasOrbit = sbasOrbit;
          // Sets the start date as the date of the orbital elements
          setStartDate(sbasOrbit.getDate());
          // Sets the mu
          this.mu = mu;
          // Sets the mass
          this.mass = mass;
          // Sets the Earth Centered Inertial frame
          this.eci  = eci;
          // Sets the Earth Centered Earth Fixed frame
          this.ecef = ecef;
          // Sets initial state
          final Orbit orbit = propagateOrbit(getStartDate());
          final Attitude attitude = provider.getAttitude(orbit, orbit.getDate(), orbit.getFrame());
          super.resetInitialState(new SpacecraftState(orbit, attitude).withMass(mass));
      }
  
      /**
       * Gets the PVCoordinates of the GNSS SV in {@link #getECEF() ECEF frame}.
       *
       * <p>The algorithm uses automatic differentiation to compute velocity and
       * acceleration.</p>
       *
       * @param date the computation date
      * @return the GNSS SV PVCoordinates in {@link #getECEF() ECEF frame}
      */
     public PVCoordinates propagateInEcef(final AbsoluteDate date) {
         // Duration from SBAS ephemeris Reference date
         final UnivariateDerivative2 dt = new UnivariateDerivative2(getDT(date), 1.0, 0.0);
         // Satellite coordinates
         final UnivariateDerivative2 x = dt.multiply(dt.multiply(0.5 * sbasOrbit.getXDotDot()).add(sbasOrbit.getXDot())).add(sbasOrbit.getX());
         final UnivariateDerivative2 y = dt.multiply(dt.multiply(0.5 * sbasOrbit.getYDotDot()).add(sbasOrbit.getYDot())).add(sbasOrbit.getY());
         final UnivariateDerivative2 z = dt.multiply(dt.multiply(0.5 * sbasOrbit.getZDotDot()).add(sbasOrbit.getZDot())).add(sbasOrbit.getZ());
         // Returns the Earth-fixed coordinates
         final FieldVector3D<UnivariateDerivative2> positionwithDerivatives =
                         new FieldVector3D<>(x, y, z);
         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} */
     public Orbit propagateOrbit(final AbsoluteDate date) {
         // Gets the PVCoordinates in ECEF frame
         final PVCoordinates pvaInECEF = propagateInEcef(date);
         // Transforms the PVCoordinates to ECI frame
         final PVCoordinates pvaInECI = ecef.getTransformTo(eci, date).transformPVCoordinates(pvaInECEF);
         // Returns the Cartesian orbit
         return new CartesianOrbit(pvaInECI, eci, date, mu);
     }
 
     /**
      * Get the Earth gravity coefficient used for SBAS propagation.
      * @return the Earth gravity coefficient.
      */
     public double getMU() {
         return mu;
     }
 
     /**
      * Gets the Earth Centered Inertial frame used to propagate the orbit.
      *
      * @return the ECI frame
      */
     public Frame getECI() {
         return eci;
     }
 
     /**
      * Gets the Earth Centered Earth Fixed frame used to propagate GNSS orbits.
      *
      * @return the ECEF frame
      */
     public Frame getECEF() {
         return ecef;
     }
 
     /**
      * Get the underlying SBAS orbital elements.
      *
      * @return the underlying SBAS orbital elements
      */
     public SBASOrbitalElements getSBASOrbitalElements() {
         return sbasOrbit;
     }
 
     /** {@inheritDoc} */
     public Frame getFrame() {
         return eci;
     }
 
     /** {@inheritDoc} */
     public void resetInitialState(final SpacecraftState state) {
         throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
     }
 
     /** {@inheritDoc} */
     protected double getMass(final AbsoluteDate date) {
         return mass;
     }
 
     /** {@inheritDoc} */
     protected void resetIntermediateState(final SpacecraftState state, final boolean forward) {
         throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
     }
 
     /** Get the duration from SBAS Reference epoch.
      * @param date the considered date
      * @return the duration from SBAS orbit Reference epoch (s)
      */
     private double getDT(final AbsoluteDate date) {
         // Time from ephemeris reference epoch
         return date.durationFrom(sbasOrbit.getDate());
     }
 
 }