/* Copyright 2022-2026 Luc Maisonobe
    * 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.CalculusFieldElement;
  import org.hipparchus.Field;
  import org.hipparchus.analysis.differentiation.FieldGradient;
  import org.hipparchus.analysis.differentiation.FieldUnivariateDerivative2;
  import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
  import org.hipparchus.geometry.euclidean.threed.Vector3D;
  import org.hipparchus.linear.FieldMatrix;
  import org.hipparchus.linear.FieldQRDecomposition;
  import org.hipparchus.linear.FieldVector;
  import org.hipparchus.linear.MatrixUtils;
  import org.hipparchus.util.FastMath;
  import org.hipparchus.util.FieldSinCos;
  import org.hipparchus.util.MathArrays;
  import org.orekit.attitudes.AttitudeProvider;
  import org.orekit.attitudes.FieldAttitude;
  import org.orekit.frames.Frame;
  import org.orekit.orbits.FieldCartesianOrbit;
  import org.orekit.orbits.FieldKeplerianAnomalyUtility;
  import org.orekit.orbits.FieldKeplerianOrbit;
  import org.orekit.orbits.FieldOrbit;
  import org.orekit.orbits.PositionAngleType;
  import org.orekit.propagation.FieldSpacecraftState;
  import org.orekit.propagation.analytical.FieldAbstractAnalyticalPropagator;
  import org.orekit.propagation.analytical.gnss.data.FieldGnssOrbitalElements;
  import org.orekit.propagation.analytical.gnss.data.GNSSOrbitalElements;
  import org.orekit.time.FieldAbsoluteDate;
  import org.orekit.utils.FieldPVCoordinates;
  import org.orekit.utils.ParameterDriver;
  
  import java.util.List;
  
  /** Common handling of {@link FieldAbstractAnalyticalPropagator} methods for GNSS propagators.
   * <p>
   * This class allows to provide easily a subset of {@link FieldAbstractAnalyticalPropagator} methods
   * for specific GNSS propagators.
   * </p>
   * @author Pascal Parraud
   * @author Luc Maisonobe
   * @param <T> type of the field elements
   * @since 13.0
   */
  public class FieldGnssPropagator<T extends CalculusFieldElement<T>> extends FieldAbstractAnalyticalPropagator<T> {
  
      /** Maximum number of iterations for internal loops. */
      private static final int MAX_ITER = 100;
  
      /** Tolerance on position for rebuilding orbital elements from initial state. */
      private static final double TOL_P = 1.0e-6;
  
      /** Tolerance on velocity for rebuilding orbital elements from initial state. */
      private static final double TOL_V = 1.0e-9;
  
      /** Number of free parameters for orbital elements. */
      private static final int FREE_PARAMETERS = 6;
  
      /** Convergence parameter. */
      private static final double EPS = 1.0e-12;
  
      /** The GNSS propagation model used. */
      private FieldGnssOrbitalElements<T, ?> orbitalElements;
  
      /** The ECI frame used for GNSS propagation. */
      private final Frame eci;
  
      /** The ECEF frame used for GNSS propagation. */
      private final Frame ecef;
  
      /**
       * Build a new instance.
       * @param orbitalElements GNSS orbital elements
       * @param eci Earth Centered Inertial frame
       * @param ecef Earth Centered Earth Fixed frame
       * @param provider Attitude provider
       * @param mass Satellite mass (kg)
       */
      public FieldGnssPropagator(final FieldGnssOrbitalElements<T, ?> orbitalElements,
                                 final Frame eci, final Frame ecef,
                                 final AttitudeProvider provider, final T mass) {
          super(orbitalElements.getDate().getField(), provider);
          // Stores the GNSS orbital elements
          this.orbitalElements = orbitalElements;
        // Sets the Earth Centered Inertial frame
         this.eci  = eci;
         // Sets the Earth Centered Earth Fixed frame
         this.ecef = ecef;
         // Sets initial state
         final FieldOrbit<T> orbit = propagateOrbit(orbitalElements.getDate(), defaultParameters());
         final FieldAttitude<T> attitude = provider.getAttitude(orbit, orbit.getDate(), orbit.getFrame());
 
         // calling the method from base class because the one overridden below recomputes the orbital elements
         super.resetInitialState(new FieldSpacecraftState<>(orbit, attitude).withMass(mass));
 
     }
 
     /**
      * Build a new instance from an initial state.
      * <p>
      * The Keplerian elements already present in the {@code nonKeplerianElements} argument
      * will be ignored as it is the {@code initialState} argument that will be used to
      * build the complete orbital elements of the propagator
      * </p>
      * @param initialState         initial state
      * @param nonKeplerianElements non-Keplerian orbital elements (the Keplerian orbital elements will be ignored)
      * @param ecef                 Earth Centered Earth Fixed frame
      * @param provider             attitude provider
      * @param mass                 spacecraft mass
      */
     public FieldGnssPropagator(final FieldSpacecraftState<T> initialState,
                                final FieldGnssOrbitalElements<T, ?> nonKeplerianElements,
                                final Frame ecef, final AttitudeProvider provider, final T mass) {
         this(buildOrbitalElements(initialState, nonKeplerianElements, ecef, provider, mass),
              initialState.getFrame(), ecef, provider, initialState.getMass());
     }
 
     /** Build default parameters.
      * @return default parameters
      */
     private T[] defaultParameters() {
         final T[] parameters = MathArrays.buildArray(orbitalElements.getDate().getField(), GNSSOrbitalElements.SIZE);
         parameters[GNSSOrbitalElements.TIME_INDEX]         = getMU().newInstance(orbitalElements.getTime());
         parameters[GNSSOrbitalElements.A_DOT_INDEX]        = getMU().newInstance(orbitalElements.getADot());
         parameters[GNSSOrbitalElements.DELTA_N0_INDEX]     = getMU().newInstance(orbitalElements.getDeltaN0());
         parameters[GNSSOrbitalElements.DELTA_N0_DOT_INDEX] = getMU().newInstance(orbitalElements.getDeltaN0Dot());
         parameters[GNSSOrbitalElements.I_DOT_INDEX]        = getMU().newInstance(orbitalElements.getIDot());
         parameters[GNSSOrbitalElements.OMEGA_DOT_INDEX]    = getMU().newInstance(orbitalElements.getOmegaDot());
         parameters[GNSSOrbitalElements.CUC_INDEX]          = getMU().newInstance(orbitalElements.getCuc());
         parameters[GNSSOrbitalElements.CUS_INDEX]          = getMU().newInstance(orbitalElements.getCus());
         parameters[GNSSOrbitalElements.CRC_INDEX]          = getMU().newInstance(orbitalElements.getCrc());
         parameters[GNSSOrbitalElements.CRS_INDEX]          = getMU().newInstance(orbitalElements.getCrs());
         parameters[GNSSOrbitalElements.CIC_INDEX]          = getMU().newInstance(orbitalElements.getCic());
         parameters[GNSSOrbitalElements.CIS_INDEX]          = getMU().newInstance(orbitalElements.getCis());
         return parameters;
     }
 
     /** {@inheritDoc} */
     @Override
     public List<ParameterDriver> getParametersDrivers() {
         return orbitalElements.getParametersDrivers();
     }
 
     /**
      * 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 according to the
      * Interface Control Document.
      *
      * @return the ECEF frame
      */
     public Frame getECEF() {
         return ecef;
     }
 
     /**
      * Gets the Earth gravity coefficient used for GNSS propagation.
      *
      * @return the Earth gravity coefficient.
      */
     public T getMU() {
         return orbitalElements.getMu();
     }
 
     /** {@inheritDoc} */
     @Override
     public FieldOrbit<T> propagateOrbit(final FieldAbsoluteDate<T> date,
                                         final T[] parameters) {
         // Gets the PVCoordinates in ECEF frame
         final FieldPVCoordinates<T> pvaInECEF = propagateInEcef(date, parameters);
         // Transforms the PVCoordinates to ECI frame
         final FieldPVCoordinates<T> pvaInECI = ecef.getTransformTo(eci, date).transformPVCoordinates(pvaInECEF);
         // Returns the Cartesian orbit
         return new FieldCartesianOrbit<>(pvaInECI, eci, date, getMU());
     }
 
     /**
      * 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
      * @param parameters propagation parameters
      * @return the GNSS SV PVCoordinates in {@link #getECEF() ECEF frame}
      */
     public FieldPVCoordinates<T> propagateInEcef(final FieldAbsoluteDate<T> date, final T[] parameters) {
         // Duration from GNSS ephemeris Reference date
         final FieldUnivariateDerivative2<T> tk = new FieldUnivariateDerivative2<>(getTk(date),
                                                                                   date.getField().getOne(),
                                                                                   date.getField().getZero());
 
         // Semi-major axis
         final FieldUnivariateDerivative2<T> ak = tk.multiply(parameters[GNSSOrbitalElements.A_DOT_INDEX]).
                                                  add(orbitalElements.getSma());
         // Mean motion
         final FieldUnivariateDerivative2<T> nA = tk.multiply(parameters[GNSSOrbitalElements.DELTA_N0_DOT_INDEX].multiply(0.5)).
                                                  add(parameters[GNSSOrbitalElements.DELTA_N0_INDEX]).
                                                  add(orbitalElements.getMeanMotion0());
         // Mean anomaly
         final FieldUnivariateDerivative2<T> mk = tk.multiply(nA).add(orbitalElements.getM0());
         // Eccentric Anomaly
         final FieldUnivariateDerivative2<T> e  = tk.newInstance(orbitalElements.getE());
         final FieldUnivariateDerivative2<T> ek = FieldKeplerianAnomalyUtility.ellipticMeanToEccentric(e, mk);
         // True Anomaly
         final FieldUnivariateDerivative2<T> vk = FieldKeplerianAnomalyUtility.ellipticEccentricToTrue(e, ek);
         // Argument of Latitude
         final FieldUnivariateDerivative2<T> phik    = vk.add(orbitalElements.getPa());
         final FieldSinCos<FieldUnivariateDerivative2<T>> cs2phi = FastMath.sinCos(phik.multiply(2));
         // Argument of Latitude Correction
         final FieldUnivariateDerivative2<T> dphik = cs2phi.cos().multiply(parameters[GNSSOrbitalElements.CUC_INDEX]).
                                                 add(cs2phi.sin().multiply(parameters[GNSSOrbitalElements.CUS_INDEX]));
         // Radius Correction
         final FieldUnivariateDerivative2<T> drk = cs2phi.cos().multiply(parameters[GNSSOrbitalElements.CRC_INDEX]).
                                               add(cs2phi.sin().multiply(parameters[GNSSOrbitalElements.CRS_INDEX]));
         // Inclination Correction
         final FieldUnivariateDerivative2<T> dik = cs2phi.cos().multiply(parameters[GNSSOrbitalElements.CIC_INDEX]).
                                               add(cs2phi.sin().multiply(parameters[GNSSOrbitalElements.CIS_INDEX]));
         // Corrected Argument of Latitude
         final FieldSinCos<FieldUnivariateDerivative2<T>> csuk = FastMath.sinCos(phik.add(dphik));
         // Corrected Radius
         final FieldUnivariateDerivative2<T> rk = ek.cos().multiply(e.negate()).add(1).multiply(ak).add(drk);
         // Corrected Inclination
         final FieldUnivariateDerivative2<T> ik  = tk.multiply(parameters[GNSSOrbitalElements.I_DOT_INDEX]).
                                                   add(orbitalElements.getI0()).add(dik);
         final FieldSinCos<FieldUnivariateDerivative2<T>> csik = FastMath.sinCos(ik);
         // Positions in orbital plane
         final FieldUnivariateDerivative2<T> xk = csuk.cos().multiply(rk);
         final FieldUnivariateDerivative2<T> yk = csuk.sin().multiply(rk);
         // Corrected longitude of ascending node
         final FieldSinCos<FieldUnivariateDerivative2<T>> csomk =
             FastMath.sinCos(tk.multiply(parameters[GNSSOrbitalElements.OMEGA_DOT_INDEX].
                             subtract(orbitalElements.getAngularVelocity())).
                             add(orbitalElements.getOmega0()).
                             subtract(parameters[GNSSOrbitalElements.TIME_INDEX].multiply(orbitalElements.getAngularVelocity())));
         // returns the Earth-fixed coordinates
         final FieldVector3D<FieldUnivariateDerivative2<T>> positionWithDerivatives =
                         new FieldVector3D<>(xk.multiply(csomk.cos()).subtract(yk.multiply(csomk.sin()).multiply(csik.cos())),
                                             xk.multiply(csomk.sin()).add(yk.multiply(csomk.cos()).multiply(csik.cos())),
                                             yk.multiply(csik.sin()));
         return new FieldPVCoordinates<>(positionWithDerivatives);
 
     }
 
     /**
      * Gets the duration from GNSS Reference epoch.
      * <p>This takes the GNSS week roll-over into account.</p>
      * @param date the considered date
      * @return the duration from GNSS orbit Reference epoch (s)
      */
     private T getTk(final FieldAbsoluteDate<T> date) {
         // Time from ephemeris reference epoch
         T tk = date.durationFrom(orbitalElements.getDate());
         // Adjusts the time to take roll over week into account
         while (tk.getReal() > 0.5 * orbitalElements.getCycleDuration()) {
             tk = tk.subtract(orbitalElements.getCycleDuration());
         }
         while (tk.getReal() < -0.5 * orbitalElements.getCycleDuration()) {
             tk = tk.add(orbitalElements.getCycleDuration());
         }
         // Returns the time from ephemeris reference epoch
         return tk;
     }
 
     /** {@inheritDoc} */
     @Override
     public Frame getFrame() {
         return eci;
     }
 
     /** {@inheritDoc} */
     @Override
     protected T getMass(final FieldAbsoluteDate<T> date) {
         return getInitialState().getMass();
     }
 
     /** {@inheritDoc} */
     @Override
     public void resetInitialState(final FieldSpacecraftState<T> state) {
         orbitalElements = buildOrbitalElements(state, orbitalElements, ecef, getAttitudeProvider(), state.getMass());
         final FieldOrbit<T> orbit = propagateOrbit(orbitalElements.getDate(), defaultParameters());
         final FieldAttitude<T> attitude = getAttitudeProvider().getAttitude(orbit, orbit.getDate(), orbit.getFrame());
         super.resetInitialState(new FieldSpacecraftState<>(orbit, attitude).withMass(state.getMass()));
     }
 
     /** {@inheritDoc} */
     @Override
     protected void resetIntermediateState(final FieldSpacecraftState<T> state, final boolean forward) {
         resetInitialState(state);
     }
 
     /**
      * Build orbital elements from initial state.
      * <p>
      * This method is roughly the inverse of {@link #propagateInEcef(FieldAbsoluteDate, CalculusFieldElement[])},
      * except it starts from a state in inertial frame
      * </p>
      *
      * @param <T> type of the field elements
      * @param initialState    initial state
      * @param nonKeplerianElements non-Keplerian orbital elements (the Keplerian orbital elements will be overridden)
      * @param ecef            Earth Centered Earth Fixed frame
      * @param provider        attitude provider
      * @param mass            satellite mass (kg)
      * @return orbital elements that generate the {@code initialState} when used with a propagator
      */
     private static <T extends CalculusFieldElement<T>> FieldGnssOrbitalElements<T, ?>
         buildOrbitalElements(final FieldSpacecraftState<T> initialState,
                              final FieldGnssOrbitalElements<T, ?> nonKeplerianElements,
                              final Frame ecef, final AttitudeProvider provider,
                              final T mass) {
 
         final Field<T> field = initialState.getDate().getField();
 
         // get approximate initial orbit
         final Frame frozenEcef = ecef.getFrozenFrame(initialState.getFrame(),
                                                      initialState.getDate().toAbsoluteDate(),
                                                      "frozen");
         final FieldKeplerianOrbit<T> orbit = approximateInitialOrbit(initialState, nonKeplerianElements, frozenEcef);
 
         // refine orbit using simple differential correction to reach target PV
         final FieldPVCoordinates<T> targetPV = initialState.getPVCoordinates();
         final FieldGnssOrbitalElements<FieldGradient<T>, ?> gElements = convert(nonKeplerianElements, orbit);
         for (int i = 0; i < MAX_ITER; ++i) {
 
             // get position-velocity derivatives with respect to initial orbit
             final FieldGnssPropagator<FieldGradient<T>> gPropagator =
                 new FieldGnssPropagator<>(gElements, initialState.getFrame(), ecef, provider,
                                           gElements.getMu().newInstance(mass));
             final FieldPVCoordinates<FieldGradient<T>> gPV = gPropagator.getInitialState().getPVCoordinates();
 
             // compute Jacobian matrix
             final FieldMatrix<T> jacobian = MatrixUtils.createFieldMatrix(field, FREE_PARAMETERS, FREE_PARAMETERS);
             jacobian.setRow(0, gPV.getPosition().getX().getGradient());
             jacobian.setRow(1, gPV.getPosition().getY().getGradient());
             jacobian.setRow(2, gPV.getPosition().getZ().getGradient());
             jacobian.setRow(3, gPV.getVelocity().getX().getGradient());
             jacobian.setRow(4, gPV.getVelocity().getY().getGradient());
             jacobian.setRow(5, gPV.getVelocity().getZ().getGradient());
 
             // linear correction to get closer to target PV
             final FieldVector<T> residuals = MatrixUtils.createFieldVector(field, FREE_PARAMETERS);
             residuals.setEntry(0, targetPV.getPosition().getX().subtract(gPV.getPosition().getX().getValue()));
             residuals.setEntry(1, targetPV.getPosition().getY().subtract(gPV.getPosition().getY().getValue()));
             residuals.setEntry(2, targetPV.getPosition().getZ().subtract(gPV.getPosition().getZ().getValue()));
             residuals.setEntry(3, targetPV.getVelocity().getX().subtract(gPV.getVelocity().getX().getValue()));
             residuals.setEntry(4, targetPV.getVelocity().getY().subtract(gPV.getVelocity().getY().getValue()));
             residuals.setEntry(5, targetPV.getVelocity().getZ().subtract(gPV.getVelocity().getZ().getValue()));
             final FieldVector<T> correction = new FieldQRDecomposition<>(jacobian, field.getZero().newInstance(EPS)).
                                               getSolver().
                                               solve(residuals);
 
             // update initial orbit
             gElements.setSma(gElements.getSma().add(correction.getEntry(0)));
             gElements.setE(gElements.getE().add(correction.getEntry(1)));
             gElements.setI0(gElements.getI0().add(correction.getEntry(2)));
             gElements.setPa(gElements.getPa().add(correction.getEntry(3)));
             gElements.setOmega0(gElements.getOmega0().add(correction.getEntry(4)));
             gElements.setM0(gElements.getM0().add(correction.getEntry(5)));
 
             final double deltaP = FastMath.sqrt(residuals.getEntry(0).getReal() * residuals.getEntry(0).getReal() +
                                                 residuals.getEntry(1).getReal() * residuals.getEntry(1).getReal() +
                                                 residuals.getEntry(2).getReal() * residuals.getEntry(2).getReal());
             final double deltaV = FastMath.sqrt(residuals.getEntry(3).getReal() * residuals.getEntry(3).getReal() +
                                                 residuals.getEntry(4).getReal() * residuals.getEntry(4).getReal() +
                                                 residuals.getEntry(5).getReal() * residuals.getEntry(5).getReal());
 
             if (deltaP < TOL_P && deltaV < TOL_V) {
                 break;
             }
 
         }
 
         return gElements.changeField(FieldGradient::getValue);
 
     }
 
     /** Compute approximate initial orbit.
      * @param <T> type of the field elements
      * @param initialState         initial state
      * @param nonKeplerianElements non-Keplerian orbital elements (the Keplerian orbital elements will be ignored)
      * @param frozenEcef           inertial frame aligned with Earth Centered Earth Fixed frame at orbit date
      * @return approximate initial orbit that generate a state close to {@code initialState}
      */
     private static <T extends CalculusFieldElement<T>> FieldKeplerianOrbit<T>
         approximateInitialOrbit(final FieldSpacecraftState<T> initialState,
                                 final FieldGnssOrbitalElements<T, ?> nonKeplerianElements,
                                 final Frame frozenEcef) {
 
         // rotate the state to a frame that is inertial but aligned with Earth frame,
         // as analytical model is expressed in Earth frame
         final FieldPVCoordinates<T> pv = initialState.getPVCoordinates(frozenEcef);
         final FieldVector3D<T> p  = pv.getPosition();
         final FieldVector3D<T>      v  = pv.getVelocity();
 
         // compute Keplerian orbital parameters
         final T   rk  = p.getNorm();
 
         // compute orbital plane orientation
         final FieldVector3D<T> normal = pv.getMomentum().normalize();
         final T   cosIk  = normal.getZ();
         final T   ik     = FieldVector3D.angle(normal, Vector3D.PLUS_K);
 
         // compute position in orbital plane
         final T   q   = FastMath.hypot(normal.getX(), normal.getY());
         final T   cos = normal.getY().negate().divide(q);
         final T   sin =  normal.getX().divide(q);
         final T   xk  =  p.getX().multiply(cos).add(p.getY().multiply(sin));
         final T   yk  = p.getY().multiply(cos).subtract(p.getX().multiply(sin)).divide(cosIk);
 
         // corrected latitude argument
         final T   uk  = FastMath.atan2(yk, xk);
 
         // recover latitude argument before correction, using a fixed-point method
         T phi = uk;
         for (int i = 0; i < MAX_ITER; ++i) {
             final T previous = phi;
             final FieldSinCos<T> cs2Phi = FastMath.sinCos(phi.multiply(2));
             phi = uk.subtract(cs2Phi.cos().multiply(nonKeplerianElements.getCuc()).add(cs2Phi.sin().multiply(nonKeplerianElements.getCus())));
             if (FastMath.abs(phi.subtract(previous).getReal()) <= EPS) {
                 break;
             }
         }
         final FieldSinCos<T> cs2phi = FastMath.sinCos(phi.multiply(2));
 
         // recover plane orientation before correction
         // here, we know that tk = 0 since our orbital elements will be at initial state date
         final T i0  = ik.subtract(cs2phi.cos().multiply(nonKeplerianElements.getCic()).add(cs2phi.sin().multiply(nonKeplerianElements.getCis())));
         final T om0 = FastMath.atan2(sin, cos).
                       add(nonKeplerianElements.getAngularVelocity() * nonKeplerianElements.getTime());
 
         // recover eccentricity and anomaly
         final T mu = initialState.getOrbit().getMu();
         final T rV2OMu           = rk.multiply(v.getNorm2Sq()).divide(mu);
         final T sma              = rk.divide(rV2OMu.negate().add(2));
         final T eCosE            = rV2OMu.subtract(1);
         final T eSinE            = FieldVector3D.dotProduct(p, v).divide(FastMath.sqrt(mu.multiply(sma)));
         final T e                = FastMath.hypot(eCosE, eSinE);
         final T eccentricAnomaly = FastMath.atan2(eSinE, eCosE);
         final T aop              = phi.subtract(eccentricAnomaly);
         final T meanAnomaly      = FieldKeplerianAnomalyUtility.ellipticEccentricToMean(e, eccentricAnomaly);
 
         return new FieldKeplerianOrbit<>(sma, e, i0, aop, om0, meanAnomaly, PositionAngleType.MEAN,
                                          PositionAngleType.MEAN, frozenEcef,
                                          initialState.getDate(), mu);
 
     }
 
     /** Convert orbital elements to gradient.
      * @param <T> type of the field elements
      * @param elements   primitive double elements
      * @param orbit      Keplerian orbit
      * @return converted elements, set up as gradient relative to Keplerian orbit
      */
     private static <T extends CalculusFieldElement<T>> FieldGnssOrbitalElements<FieldGradient<T>, ?>
         convert(final FieldGnssOrbitalElements<T, ?> elements, final FieldKeplerianOrbit<T> orbit) {
 
         final FieldGnssOrbitalElements<FieldGradient<T>, ?> gElements =
             elements.changeField(t -> FieldGradient.constant(FREE_PARAMETERS, t));
 
         // Keplerian parameters
         gElements.setSma(FieldGradient.variable(FREE_PARAMETERS, 0, orbit.getA()));
         gElements.setE(FieldGradient.variable(FREE_PARAMETERS, 1, orbit.getE()));
         gElements.setI0(FieldGradient.variable(FREE_PARAMETERS, 2, orbit.getI()));
         gElements.setPa(FieldGradient.variable(FREE_PARAMETERS, 3, orbit.getPerigeeArgument()));
         gElements.setOmega0(FieldGradient.variable(FREE_PARAMETERS, 4, orbit.getRightAscensionOfAscendingNode()));
         gElements.setM0(FieldGradient.variable(FREE_PARAMETERS, 5, orbit.getMeanAnomaly()));
 
         return gElements;
 
     }
 
 }