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  package org.orekit.forces.radiation;
  
  import java.util.ArrayList;
  import java.util.Collections;
  import java.util.List;
  
  import org.hipparchus.CalculusFieldElement;
  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.hipparchus.util.SinCos;
  import org.orekit.annotation.DefaultDataContext;
  import org.orekit.bodies.OneAxisEllipsoid;
  import org.orekit.frames.FramesFactory;
  import org.orekit.propagation.FieldSpacecraftState;
  import org.orekit.propagation.SpacecraftState;
  import org.orekit.utils.ExtendedPositionProvider;
  import org.orekit.utils.ParameterDriver;
  
  /**
   * The Empirical CODE Orbit Model 2 (ECOM2) of the Center for Orbit Determination in Europe (CODE).
   * <p>
   * The drag acceleration is computed as follows :
   * γ = γ<sub>0</sub> + D(u)e<sub>D</sub> + Y(u)e<sub>Y</sub> + B(u)e<sub>B</sub>
   * </p> <p>
   * In the above equation, γ<sub>0</sub> is a selectable a priori model. Since 2013, no
   * a priori model is used for CODE IGS contribution (i.e. γ<sub>0</sub> = 0). Moreover,
   * u denotes the satellite's argument of latitude.
   * </p> <p>
   * D(u), Y(u) and B(u) are three functions of the ECOM2 model that can be represented
   * as Fourier series. The coefficients of the Fourier series are estimated during the
   * estimation process. he ECOM2 model has user-defines upper limits <i>nD</i> and
   * <i>nB</i> for the Fourier series (i.e. <i>nD</i> for D(u) and <i>nB</i> for
   * B(u). Y(u) is defined as a constant value).
   * </p> <p>
   * It exists several configurations to initialize <i>nD</i> and <i>nB</i> values. However,
   * Arnold et al recommend to use <b>D2B1</b> (i.e. <i>nD</i> = 1 and <i>nB</i> = 1) and
   * <b>D4B1</b> (i.e. <i>nD</i> = 2 an <i>nB</i> = 1) configurations. At the opposite, in Arnold paper, it
   * is recommend to not use <b>D2B0</b> (i.e. <i>nD</i> = 1 and <i>nB</i> = 0) configuration.
   * </p> <p>
   * Since Orekit 11.0, it is possible to take into account
   * the eclipses generated by Moon in the solar radiation
   * pressure force model using the
   * {@link #addOccultingBody(ExtendedPositionProvider, double)}
   * method.<br>
   * <code> ECOM2 srp =</code>
   * <code>       new ECOM2(1, 1, 0.0, CelestialBodyFactory.getSun(), Constants.EIGEN5C_EARTH_EQUATORIAL_RADIUS);</code><br>
   * <code> srp.addOccultingBody(CelestialBodyFactory.getMoon(), Constants.MOON_EQUATORIAL_RADIUS);</code><br>
   *
   * @see "Arnold, Daniel, et al, CODE’s new solar radiation pressure model for GNSS orbit determination,
   *       Journal of geodesy 89.8 (2015): 775-791."
   *
   * @see "Tzu-Pang tseng and Michael Moore, Impact of solar radiation pressure mis-modeling on
   *       GNSS satellite orbit determination, IGS Worshop, Wuhan, China, 2018."
   *
   * @author David Soulard
   * @since 10.2
   */
  public class ECOM2 extends AbstractRadiationForceModel {
  
      /** Parameter name for ECOM model coefficients enabling Jacobian processing. */
      public static final String ECOM_COEFFICIENT = "ECOM coefficient";
  
      /** Minimum value for ECOM2 estimated parameters. */
      private static final double MIN_VALUE = Double.NEGATIVE_INFINITY;
  
      /** Maximum value for ECOM2 estimated parameters. */
      private static final double MAX_VALUE = Double.POSITIVE_INFINITY;
  
      /** Parameters scaling factor.
       * <p>
       * We use a power of 2 to avoid numeric noise introduction
       * in the multiplications/divisions sequences.
       * </p>
       */
      private final double SCALE = FastMath.scalb(1.0, -22);
  
      /** Highest order for parameter along eD axis (satellite --> sun direction). */
      private final int nD;
  
      /** Highest order for parameter along eB axis. */
      private final int nB;
 
     /** Estimated acceleration coefficients.
      * <p>
      * The 2 * nD first driver are Fourier driver along eD, axis,
      * then along eY, then 2*nB following are along eB axis.
      * </p>
      */
     private final List<ParameterDriver> coefficients;
 
     /** Sun model. */
     private final ExtendedPositionProvider sun;
 
     /**
      * Constructor.
      * @param nD truncation rank of Fourier series in D term.
      * @param nB truncation rank of Fourier series in B term.
      * @param value parameters initial value
      * @param sun provide for Sun parameter
      * @param equatorialRadius spherical shape model (for umbra/penumbra computation)
      */
     @DefaultDataContext
     public ECOM2(final int nD, final int nB, final double value,
                  final ExtendedPositionProvider sun, final double equatorialRadius) {
         super(sun, new OneAxisEllipsoid(equatorialRadius, 0.0, FramesFactory.getGCRF()),
                 getDefaultEclipseDetectionSettings());
         this.nB = nB;
         this.nD = nD;
         this.coefficients = new ArrayList<>(2 * (nD + nB) + 3);
 
         // Add parameter along eB axis in alphabetical order
         coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " B0", value, SCALE, MIN_VALUE, MAX_VALUE));
         for (int i = 1; i < nB + 1; i++) {
             coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Bcos" + Integer.toString(i - 1), value, SCALE, MIN_VALUE, MAX_VALUE));
         }
         for (int i = nB + 1; i < 2 * nB + 1; i++) {
             coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Bsin" + Integer.toString(i - (nB + 1)), value, SCALE, MIN_VALUE, MAX_VALUE));
         }
         // Add driver along eD axis in alphabetical order
         coefficients.add(2 * nB + 1, new ParameterDriver(ECOM_COEFFICIENT + " D0", value, SCALE, MIN_VALUE, MAX_VALUE));
         for (int i = 2 * nB + 2; i < 2 * nB + 2 + nD; i++) {
             coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Dcos" + Integer.toString(i - (2 * nB + 2)), value, SCALE, MIN_VALUE, MAX_VALUE));
         }
         for (int i = 2 * nB + 2 + nD; i < 2 * (nB + nD) + 2; i++) {
             coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Dsin" + Integer.toString(i - (2 * nB + nD + 2)), value, SCALE, MIN_VALUE, MAX_VALUE));
         }
         // Add  Y0
         coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Y0", value, SCALE, MIN_VALUE, MAX_VALUE));
 
         // For ECOM2 model, all parameters are estimated
         coefficients.forEach(parameter -> parameter.setSelected(true));
         this.sun = sun;
     }
 
     /** {@inheritDoc} */
     @Override
     public Vector3D acceleration(final SpacecraftState s, final double[] parameters) {
 
         // Spacecraft and Sun position vectors (expressed in the spacecraft's frame)
         final Vector3D satPos = s.getPosition();
         final Vector3D sunPos = sun.getPosition(s.getDate(), s.getFrame());
 
         // Build the coordinate system
         final Vector3D Z = s.getPVCoordinates().getMomentum();
         final Vector3D Y = Z.crossProduct(sunPos).normalize();
         final Vector3D X = Y.crossProduct(Z).normalize();
 
         // Build eD, eY, eB vectors
         final Vector3D eD = sunPos.subtract(satPos).normalize();
         final Vector3D eY = eD.crossProduct(satPos).normalize();
         final Vector3D eB = eD.crossProduct(eY);
 
         // Angular argument difference u_s - u
         final double delta_u = FastMath.atan2(satPos.dotProduct(Y), satPos.dotProduct(X));
 
         // Compute B(u)
         double b_u = parameters[0];
         for (int i = 1; i < nB + 1; i++) {
             final SinCos sc = FastMath.sinCos((2 * i - 1) * delta_u);
             b_u += parameters[i] * sc.cos() + parameters[i + nB] * sc.sin();
         }
         // Compute D(u)
         double d_u = parameters[2 * nB + 1];
         for (int i = 1; i < nD + 1; i++) {
             final SinCos sc = FastMath.sinCos(2 * i * delta_u);
             d_u += parameters[2 * nB + 1 + i] * sc.cos() + parameters[2 * nB + 1 + i + nD] * sc.sin();
         }
         // Return acceleration
         return new Vector3D(d_u, eD, parameters[2 * (nD + nB) + 2], eY, b_u, eB);
     }
 
     /** {@inheritDoc} */
     @Override
     public <T extends CalculusFieldElement<T>> FieldVector3D<T> acceleration(final FieldSpacecraftState<T> s, final T[] parameters) {
 
         // Spacecraft and Sun position vectors (expressed in the spacecraft's frame)
         final FieldVector3D<T> satPos = s.getPosition();
         final FieldVector3D<T> sunPos = sun.getPosition(s.getDate(), s.getFrame());
 
         // Build the coordinate system
         final FieldVector3D<T> Z = s.getPVCoordinates().getMomentum();
         final FieldVector3D<T> Y = Z.crossProduct(sunPos).normalize();
         final FieldVector3D<T> X = Y.crossProduct(Z).normalize();
 
         // Build eD, eY, eB vectors
         final FieldVector3D<T> eD = sunPos.subtract(satPos).normalize();
         final FieldVector3D<T> eY = eD.crossProduct(satPos).normalize();
         final FieldVector3D<T> eB = eD.crossProduct(eY);
 
         // Angular argument difference u_s - u
         final T  delta_u = FastMath.atan2(satPos.dotProduct(Y), satPos.dotProduct(X));
 
         // Compute B(u)
         T b_u =  parameters[0];
         for (int i = 1; i < nB + 1; i++) {
             final FieldSinCos<T> sc = FastMath.sinCos(delta_u.multiply(2 * i - 1));
             b_u = b_u.add(sc.cos().multiply(parameters[i])).add(sc.sin().multiply(parameters[i + nB]));
         }
         // Compute D(u)
         T d_u = parameters[2 * nB + 1];
 
         for (int i = 1; i < nD + 1; i++) {
             final FieldSinCos<T> sc = FastMath.sinCos(delta_u.multiply(2 * i));
             d_u =  d_u.add(sc.cos().multiply(parameters[2 * nB + 1 + i])).add(sc.sin().multiply(parameters[2 * nB + 1 + i + nD]));
         }
         // Return the acceleration
         return new FieldVector3D<>(d_u, eD, parameters[2 * (nD + nB) + 2], eY, b_u, eB);
     }
 
     /** {@inheritDoc} */
     @Override
     public List<ParameterDriver> getParametersDrivers() {
         return Collections.unmodifiableList(coefficients);
     }
 
 }