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  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.hipparchus.util.FastMath;
  import org.hipparchus.util.FieldSinCos;
  import org.hipparchus.util.MathArrays;
  import org.hipparchus.util.MathUtils;
  import org.hipparchus.util.Precision;
  import org.orekit.attitudes.Attitude;
  import org.orekit.attitudes.AttitudeProvider;
  import org.orekit.data.DataContext;
  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.GLONASSAlmanac;
  import org.orekit.propagation.analytical.gnss.data.GLONASSNavigationMessage;
  import org.orekit.propagation.analytical.gnss.data.GLONASSOrbitalElements;
  import org.orekit.propagation.analytical.gnss.data.GNSSConstants;
  import org.orekit.time.AbsoluteDate;
  import org.orekit.time.GLONASSDate;
  import org.orekit.time.TimeScale;
  import org.orekit.utils.PVCoordinates;
  
  /**
   * This class aims at propagating a GLONASS orbit from {@link GLONASSOrbitalElements}.
   * <p>
   * <b>Caution:</b> The Glonass analytical propagator can only be used with {@link GLONASSAlmanac}.
   * Using this propagator with a {@link GLONASSNavigationMessage} is prone to error.
   * </p>
   *
   * @see <a href="http://russianspacesystems.ru/wp-content/uploads/2016/08/ICD-GLONASS-CDMA-General.-Edition-1.0-2016.pdf">
   *       GLONASS Interface Control Document</a>
   *
   * @author Bryan Cazabonne
   * @since 10.0
   *
   */
  public class GLONASSAnalyticalPropagator extends AbstractAnalyticalPropagator {
  
      // Constants
      /** Constant 7.0 / 3.0. */
      private static final double SEVEN_THIRD = 7.0 / 3.0;
  
      /** Constant 7.0 / 6.0. */
      private static final double SEVEN_SIXTH = 7.0 / 6.0;
  
      /** Constant 7.0 / 24.0. */
      private static final double SEVEN_24TH = 7.0 / 24.0;
  
      /** Constant 49.0 / 72.0. */
      private static final double FN_72TH = 49.0 / 72.0;
  
      /** Value of the earth's rotation rate in rad/s. */
      private static final double GLONASS_AV = 7.2921150e-5;
  
      /** Mean value of inclination for Glonass orbit is equal to 63°. */
      private static final double GLONASS_MEAN_INCLINATION = 64.8;
  
      /** Mean value of Draconian period for Glonass orbit is equal to 40544s : 11 hours 15 minutes 44 seconds. */
      private static final double GLONASS_MEAN_DRACONIAN_PERIOD = 40544;
  
      /** Second degree zonal coefficient of normal potential. */
      private static final double GLONASS_J20 = 1.08262575e-3;
  
      /** Equatorial radius of Earth (m). */
      private static final double GLONASS_EARTH_EQUATORIAL_RADIUS = 6378136;
  
      // Data used to solve Kepler's equation
      /** First coefficient to compute Kepler equation solver starter. */
      private static final double A;
  
      /** Second coefficient to compute Kepler equation solver starter. */
      private static final double B;
  
      static {
          final double k1 = 3 * FastMath.PI + 2;
          final double k2 = FastMath.PI - 1;
         final double k3 = 6 * FastMath.PI - 1;
         A  = 3 * k2 * k2 / k1;
         B  = k3 * k3 / (6 * k1);
     }
 
     /** The GLONASS orbital elements used. */
     private final GLONASSOrbitalElements glonassOrbit;
 
     /** The spacecraft mass (kg). */
     private final double mass;
 
     /** The ECI frame used for GLONASS propagation. */
     private final Frame eci;
 
     /** The ECEF frame used for GLONASS propagation. */
     private final Frame ecef;
 
     /** Data context for propagation. */
     private final DataContext dataContext;
 
     /**
      * Private constructor.
      * @param glonassOrbit 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 context Data context
      */
     GLONASSAnalyticalPropagator(final GLONASSOrbitalElements glonassOrbit, final Frame eci,
                                 final Frame ecef, final AttitudeProvider provider,
                                 final double mass, final DataContext context) {
         super(provider);
         this.dataContext = context;
         // Stores the GLONASS orbital elements
         this.glonassOrbit = glonassOrbit;
         // Sets the start date as the date of the orbital elements
         setStartDate(glonassOrbit.getDate());
         // 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 GLONASS SV in {@link #getECEF() ECEF frame}.
      *
      * <p>The algorithm is defined at Appendix M.1 from GLONASS Interface Control Document,
      * with automatic differentiation added to compute velocity and
      * acceleration.</p>
      *
      * @param date the computation date
      * @return the GLONASS SV PVCoordinates in {@link #getECEF() ECEF frame}
      */
     public PVCoordinates propagateInEcef(final AbsoluteDate date) {
 
         // Interval of prediction dTpr
         final UnivariateDerivative2 dTpr = getdTpr(date);
 
         // Zero
         final UnivariateDerivative2 zero = dTpr.getField().getZero();
 
         // The number of whole orbits "w" on a prediction interval
         final UnivariateDerivative2 w = FastMath.floor(dTpr.divide(GLONASS_MEAN_DRACONIAN_PERIOD + glonassOrbit.getDeltaT()));
 
         // Current inclination
         final UnivariateDerivative2 i = zero.newInstance(GLONASS_MEAN_INCLINATION / 180 * GNSSConstants.GLONASS_PI + glonassOrbit.getDeltaI());
 
         // Eccentricity
         final UnivariateDerivative2 e = zero.newInstance(glonassOrbit.getE());
 
         // Mean draconique period in orbite w+1 and mean motion
         final UnivariateDerivative2 tDR = w.multiply(2.0).add(1.0).multiply(glonassOrbit.getDeltaTDot()).
                                           add(glonassOrbit.getDeltaT()).
                                           add(GLONASS_MEAN_DRACONIAN_PERIOD);
         final UnivariateDerivative2 n = tDR.divide(2.0 * GNSSConstants.GLONASS_PI).reciprocal();
 
         // Semi-major axis : computed by successive approximation
         final UnivariateDerivative2 sma = computeSma(tDR, i, e);
 
         // (ae / p)^2 term
         final UnivariateDerivative2 p     = sma.multiply(e.multiply(e).negate().add(1.0));
         final UnivariateDerivative2 aeop  = p.divide(GLONASS_EARTH_EQUATORIAL_RADIUS).reciprocal();
         final UnivariateDerivative2 aeop2 = aeop.multiply(aeop);
 
         // Current longitude of the ascending node
         final UnivariateDerivative2 lambda = computeLambda(dTpr, n, aeop2, i);
 
         // Current argument of perigee
         final UnivariateDerivative2 pa = computePA(dTpr, n, aeop2, i);
 
         // Mean longitude at the instant the spacecraft passes the current ascending node
         final UnivariateDerivative2 tanPAo2 = FastMath.tan(pa.divide(2.0));
         final UnivariateDerivative2 coef    = tanPAo2.multiply(FastMath.sqrt(e.negate().add(1.0).divide(e.add(1.0))));
         final UnivariateDerivative2 e0      = FastMath.atan(coef).multiply(2.0).negate();
         final UnivariateDerivative2 m1      = pa.add(e0).subtract(FastMath.sin(e0).multiply(e));
 
         // Current mean longitude
         final UnivariateDerivative2 correction = dTpr.
                                                  subtract(w.multiply(GLONASS_MEAN_DRACONIAN_PERIOD + glonassOrbit.getDeltaT())).
                                                  subtract(w.square().multiply(glonassOrbit.getDeltaTDot()));
         final UnivariateDerivative2 m = m1.add(n.multiply(correction));
 
         // Take into consideration the periodic perturbations
         final FieldSinCos<UnivariateDerivative2> scPa = FastMath.sinCos(pa);
         final UnivariateDerivative2 h = e.multiply(scPa.sin());
         final UnivariateDerivative2 l = e.multiply(scPa.cos());
         // δa1
         final UnivariateDerivative2[] d1 = getParameterDifferentials(sma, i, h, l, m1);
         // δa2
         final UnivariateDerivative2[] d2 = getParameterDifferentials(sma, i, h, l, m);
         // Apply corrections
         final UnivariateDerivative2 smaCorr    = sma.add(d2[0]).subtract(d1[0]);
         final UnivariateDerivative2 hCorr      = h.add(d2[1]).subtract(d1[1]);
         final UnivariateDerivative2 lCorr      = l.add(d2[2]).subtract(d1[2]);
         final UnivariateDerivative2 lambdaCorr = lambda.add(d2[3]).subtract(d1[3]);
         final UnivariateDerivative2 iCorr      = i.add(d2[4]).subtract(d1[4]);
         final UnivariateDerivative2 mCorr      = m.add(d2[5]).subtract(d1[5]);
         final UnivariateDerivative2 eCorr      = FastMath.sqrt(hCorr.multiply(hCorr).add(lCorr.multiply(lCorr)));
         final UnivariateDerivative2 paCorr;
         if (eCorr.getValue() == 0.) {
             paCorr = zero;
         } else {
             if (lCorr.getValue() == eCorr.getValue()) {
                 paCorr = zero.newInstance(0.5 * GNSSConstants.GLONASS_PI);
             } else if (lCorr.getValue() == -eCorr.getValue()) {
                 paCorr = zero.newInstance(-0.5 * GNSSConstants.GLONASS_PI);
             } else {
                 paCorr = FastMath.atan2(hCorr, lCorr);
             }
         }
 
         // Eccentric Anomaly
         final UnivariateDerivative2 mk = mCorr.subtract(paCorr);
         final UnivariateDerivative2 ek = getEccentricAnomaly(mk, eCorr);
 
         // True Anomaly
         final UnivariateDerivative2 vk =  getTrueAnomaly(ek, eCorr);
 
         // Argument of Latitude
         final UnivariateDerivative2 phik = vk.add(paCorr);
 
         // Corrected Radius
         final UnivariateDerivative2 pCorr = smaCorr.multiply(eCorr.multiply(eCorr).negate().add(1.0));
         final UnivariateDerivative2 rk    = pCorr.divide(eCorr.multiply(FastMath.cos(vk)).add(1.0));
 
         // Positions in orbital plane
         final FieldSinCos<UnivariateDerivative2> scPhik = FastMath.sinCos(phik);
         final UnivariateDerivative2 xk = scPhik.cos().multiply(rk);
         final UnivariateDerivative2 yk = scPhik.sin().multiply(rk);
 
         // Coordinates of position
         final FieldSinCos<UnivariateDerivative2> scL = FastMath.sinCos(lambdaCorr);
         final FieldSinCos<UnivariateDerivative2> scI = FastMath.sinCos(iCorr);
         final FieldVector3D<UnivariateDerivative2> positionwithDerivatives =
                         new FieldVector3D<>(xk.multiply(scL.cos()).subtract(yk.multiply(scL.sin()).multiply(scI.cos())),
                                             xk.multiply(scL.sin()).add(yk.multiply(scL.cos()).multiply(scI.cos())),
                                             yk.multiply(scI.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()));
     }
 
     /**
      * Gets eccentric anomaly from mean anomaly.
      * <p>The algorithm used to solve the Kepler equation has been published in:
      * "Procedures for  solving Kepler's Equation", A. W. Odell and R. H. Gooding,
      * Celestial Mechanics 38 (1986) 307-334</p>
      * <p>It has been copied from the OREKIT library (KeplerianOrbit class).</p>
      *
      * @param mk the mean anomaly (rad)
      * @param e the eccentricity
      * @return the eccentric anomaly (rad)
      */
     private UnivariateDerivative2 getEccentricAnomaly(final UnivariateDerivative2 mk, final UnivariateDerivative2 e) {
 
         // reduce M to [-PI PI] interval
         final UnivariateDerivative2 reducedM = new UnivariateDerivative2(MathUtils.normalizeAngle(mk.getValue(), 0.0),
                                                                          mk.getFirstDerivative(),
                                                                          mk.getSecondDerivative());
 
         // compute start value according to A. W. Odell and R. H. Gooding S12 starter
         UnivariateDerivative2 ek;
         if (FastMath.abs(reducedM.getValue()) < 1.0 / 6.0) {
             if (FastMath.abs(reducedM.getValue()) < Precision.SAFE_MIN) {
                 // this is an Orekit change to the S12 starter.
                 // If reducedM is 0.0, the derivative of cbrt is infinite which induces NaN appearing later in
                 // the computation. As in this case E and M are almost equal, we initialize ek with reducedM
                 ek = reducedM;
             } else {
                 // this is the standard S12 starter
                 ek = reducedM.add(reducedM.multiply(6).cbrt().subtract(reducedM).multiply(e));
             }
         } else {
             if (reducedM.getValue() < 0) {
                 final UnivariateDerivative2 w = reducedM.add(FastMath.PI);
                 ek = reducedM.add(w.multiply(-A).divide(w.subtract(B)).subtract(FastMath.PI).subtract(reducedM).multiply(e));
             } else {
                 final UnivariateDerivative2 minusW = reducedM.subtract(FastMath.PI);
                 ek = reducedM.add(minusW.multiply(A).divide(minusW.add(B)).add(FastMath.PI).subtract(reducedM).multiply(e));
             }
         }
 
         final UnivariateDerivative2 e1 = e.negate().add(1.0);
         final boolean noCancellationRisk = (e1.getValue() + ek.getValue() * ek.getValue() / 6) >= 0.1;
 
         // perform two iterations, each consisting of one Halley step and one Newton-Raphson step
         for (int j = 0; j < 2; ++j) {
             final UnivariateDerivative2 f;
             UnivariateDerivative2 fd;
             final UnivariateDerivative2 fdd  = ek.sin().multiply(e);
             final UnivariateDerivative2 fddd = ek.cos().multiply(e);
             if (noCancellationRisk) {
                 f  = ek.subtract(fdd).subtract(reducedM);
                 fd = fddd.subtract(1).negate();
             } else {
                 f  = eMeSinE(ek, e).subtract(reducedM);
                 final UnivariateDerivative2 s = ek.multiply(0.5).sin();
                 fd = s.multiply(s).multiply(e.multiply(2.0)).add(e1);
             }
             final UnivariateDerivative2 dee = f.multiply(fd).divide(f.multiply(0.5).multiply(fdd).subtract(fd.multiply(fd)));
 
             // update eccentric anomaly, using expressions that limit underflow problems
             final UnivariateDerivative2 w = fd.add(dee.multiply(0.5).multiply(fdd.add(dee.multiply(fdd).divide(3))));
             fd = fd.add(dee.multiply(fdd.add(dee.multiply(0.5).multiply(fdd))));
             ek = ek.subtract(f.subtract(dee.multiply(fd.subtract(w))).divide(fd));
         }
 
         // expand the result back to original range
         ek = ek.add(mk.getValue() - reducedM.getValue());
 
         // Returns the eccentric anomaly
         return ek;
     }
 
     /**
      * Accurate computation of E - e sin(E).
      *
      * @param E eccentric anomaly
      * @param ecc the eccentricity
      * @return E - e sin(E)
      */
     private UnivariateDerivative2 eMeSinE(final UnivariateDerivative2 E, final UnivariateDerivative2 ecc) {
         UnivariateDerivative2 x = E.sin().multiply(ecc.negate().add(1.0));
         final UnivariateDerivative2 mE2 = E.negate().multiply(E);
         UnivariateDerivative2 term = E;
         UnivariateDerivative2 d    = E.getField().getZero();
         // the inequality test below IS intentional and should NOT be replaced by a check with a small tolerance
         for (UnivariateDerivative2 x0 = d.add(Double.NaN); !Double.valueOf(x.getValue()).equals(x0.getValue());) {
             d = d.add(2);
             term = term.multiply(mE2.divide(d.multiply(d.add(1))));
             x0 = x;
             x = x.subtract(term);
         }
         return x;
     }
 
     /** Gets true anomaly from eccentric anomaly.
     *
     * @param ek the eccentric anomaly (rad)
     * @param ecc the eccentricity
     * @return the true anomaly (rad)
     */
     private UnivariateDerivative2 getTrueAnomaly(final UnivariateDerivative2 ek, final UnivariateDerivative2 ecc) {
         final UnivariateDerivative2 svk = ek.sin().multiply(FastMath.sqrt( ecc.square().negate().add(1.0)));
         final UnivariateDerivative2 cvk = ek.cos().subtract(ecc);
         return svk.atan2(cvk);
     }
 
     /**
      * Get the interval of prediction.
      *
      * @param date the considered date
      * @return the duration from GLONASS orbit Reference epoch (s)
      */
     private UnivariateDerivative2 getdTpr(final AbsoluteDate date) {
         final TimeScale glonass = dataContext.getTimeScales().getGLONASS();
         final GLONASSDate tEnd = new GLONASSDate(date, glonass);
         final GLONASSDate tSta = new GLONASSDate(glonassOrbit.getDate(), glonass);
         final int n  = tEnd.getDayNumber();
         final int na = tSta.getDayNumber();
         final int deltaN;
         if (na == 27) {
             deltaN = n - na - FastMath.round((float) (n - na) / 1460) * 1460;
         } else {
             deltaN = n - na - FastMath.round((float) (n - na) / 1461) * 1461;
         }
         final UnivariateDerivative2 ti = new UnivariateDerivative2(tEnd.getSecInDay(), 1.0, 0.0);
 
         return ti.subtract(glonassOrbit.getTime()).add(86400 * deltaN);
     }
 
     /**
      * Computes the semi-major axis of orbit using technique of successive approximations.
      * @param tDR mean draconique period (s)
      * @param i current inclination (rad)
      * @param e eccentricity
      * @return the semi-major axis (m).
      */
     private UnivariateDerivative2 computeSma(final UnivariateDerivative2 tDR,
                                              final UnivariateDerivative2 i,
                                              final UnivariateDerivative2 e) {
 
         // Zero
         final UnivariateDerivative2 zero = tDR.getField().getZero();
 
         // If one of the input parameter is equal to Double.NaN, an infinite loop can occur.
         // In that case, we do not compute the value of the semi major axis.
         // We decided to return a Double.NaN value instead.
         if (Double.isNaN(tDR.getValue()) || Double.isNaN(i.getValue()) || Double.isNaN(e.getValue())) {
             return zero.add(Double.NaN);
         }
 
         // Common parameters
         final UnivariateDerivative2 sinI         = FastMath.sin(i);
         final UnivariateDerivative2 sin2I        = sinI.multiply(sinI);
         final UnivariateDerivative2 ome2         = e.multiply(e).negate().add(1.0);
         final UnivariateDerivative2 ome2Pow3o2   = FastMath.sqrt(ome2).multiply(ome2);
         final UnivariateDerivative2 pa           = zero.newInstance(glonassOrbit.getPa());
         final UnivariateDerivative2 cosPA        = FastMath.cos(pa);
         final UnivariateDerivative2 opecosPA     = e.multiply(cosPA).add(1.0);
         final UnivariateDerivative2 opecosPAPow2 = opecosPA.multiply(opecosPA);
         final UnivariateDerivative2 opecosPAPow3 = opecosPAPow2.multiply(opecosPA);
 
         // Initial approximation
         UnivariateDerivative2 tOCK = tDR;
 
         // Successive approximations
         // The process of approximation ends when fulfilling the following condition: |a(n+1) - a(n)| < 1cm
         UnivariateDerivative2 an   = zero;
         UnivariateDerivative2 anp1 = zero;
         boolean isLastStep = false;
         while (!isLastStep) {
 
             // a(n+1) computation
             final UnivariateDerivative2 tOCKo2p     = tOCK.divide(2.0 * GNSSConstants.GLONASS_PI);
             final UnivariateDerivative2 tOCKo2pPow2 = tOCKo2p.multiply(tOCKo2p);
             anp1 = FastMath.cbrt(tOCKo2pPow2.multiply(GNSSConstants.GLONASS_MU));
 
             // p(n+1) computation
             final UnivariateDerivative2 p = anp1.multiply(ome2);
 
             // Tock(n+1) computation
             final UnivariateDerivative2 aeop  = p.divide(GLONASS_EARTH_EQUATORIAL_RADIUS).reciprocal();
             final UnivariateDerivative2 aeop2 = aeop.multiply(aeop);
             final UnivariateDerivative2 term1 = aeop2.multiply(GLONASS_J20).multiply(1.5);
             final UnivariateDerivative2 term2 = sin2I.multiply(2.5).negate().add(2.0);
             final UnivariateDerivative2 term3 = ome2Pow3o2.divide(opecosPAPow2);
             final UnivariateDerivative2 term4 = opecosPAPow3.divide(ome2);
             tOCK = tDR.divide(term1.multiply(term2.multiply(term3).add(term4)).negate().add(1.0));
 
             // Check convergence
             if (FastMath.abs(anp1.subtract(an).getReal()) <= 0.01) {
                 isLastStep = true;
             }
 
             an = anp1;
         }
 
         return an;
 
     }
 
     /**
      * Computes the current longitude of the ascending node.
      * @param dTpr interval of prediction (s)
      * @param n mean motion (rad/s)
      * @param aeop2 square of the ratio between the radius of the ellipsoid and p, with p = sma * (1 - ecc²)
      * @param i inclination (rad)
      * @return the current longitude of the ascending node (rad)
      */
     private UnivariateDerivative2 computeLambda(final UnivariateDerivative2 dTpr,
                                                 final UnivariateDerivative2 n,
                                                 final UnivariateDerivative2 aeop2,
                                                 final UnivariateDerivative2 i) {
         final UnivariateDerivative2 cosI = FastMath.cos(i);
         final UnivariateDerivative2 precession = aeop2.multiply(n).multiply(cosI).multiply(1.5 * GLONASS_J20);
         return dTpr.multiply(precession.add(GLONASS_AV)).negate().add(glonassOrbit.getLambda());
     }
 
     /**
      * Computes the current argument of perigee.
      * @param dTpr interval of prediction (s)
      * @param n mean motion (rad/s)
      * @param aeop2 square of the ratio between the radius of the ellipsoid and p, with p = sma * (1 - ecc²)
      * @param i inclination (rad)
      * @return the current argument of perigee (rad)
      */
     private UnivariateDerivative2 computePA(final UnivariateDerivative2 dTpr,
                                             final UnivariateDerivative2 n,
                                             final UnivariateDerivative2 aeop2,
                                             final UnivariateDerivative2 i) {
         final UnivariateDerivative2 cosI  = FastMath.cos(i);
         final UnivariateDerivative2 cos2I = cosI.multiply(cosI);
         final UnivariateDerivative2 precession = aeop2.multiply(n).multiply(cos2I.multiply(5.0).negate().add(1.0)).multiply(0.75 * GLONASS_J20);
         return dTpr.multiply(precession).negate().add(glonassOrbit.getPa());
     }
 
     /**
      * Computes the differentials δa<sub>i</sub>.
      * <p>
      * The value of i depends of the type of longitude (i = 2 for the current mean longitude;
      * i = 1 for the mean longitude at the instant the spacecraft passes the current ascending node)
      * </p>
      * @param a semi-major axis (m)
      * @param i inclination (rad)
      * @param h x component of the eccentricity (rad)
      * @param l y component of the eccentricity (rad)
      * @param m longitude (current or at the ascending node instant)
      * @return the differentials of the orbital parameters
      */
     private UnivariateDerivative2[] getParameterDifferentials(final UnivariateDerivative2 a, final UnivariateDerivative2 i,
                                                               final UnivariateDerivative2 h, final UnivariateDerivative2 l,
                                                               final UnivariateDerivative2 m) {
 
         // B constant
         final UnivariateDerivative2 aeoa  = a.divide(GLONASS_EARTH_EQUATORIAL_RADIUS).reciprocal();
         final UnivariateDerivative2 aeoa2 = aeoa.multiply(aeoa);
         final UnivariateDerivative2 b     = aeoa2.multiply(1.5 * GLONASS_J20);
 
         // Commons Parameters
         final FieldSinCos<UnivariateDerivative2> scI   = FastMath.sinCos(i);
         final FieldSinCos<UnivariateDerivative2> scLk  = FastMath.sinCos(m);
         final FieldSinCos<UnivariateDerivative2> sc2Lk = FieldSinCos.sum(scLk, scLk);
         final FieldSinCos<UnivariateDerivative2> sc3Lk = FieldSinCos.sum(scLk, sc2Lk);
         final FieldSinCos<UnivariateDerivative2> sc4Lk = FieldSinCos.sum(sc2Lk, sc2Lk);
         final UnivariateDerivative2 cosI   = scI.cos();
         final UnivariateDerivative2 sinI   = scI.sin();
         final UnivariateDerivative2 cosI2  = cosI.multiply(cosI);
         final UnivariateDerivative2 sinI2  = sinI.multiply(sinI);
         final UnivariateDerivative2 cosLk  = scLk.cos();
         final UnivariateDerivative2 sinLk  = scLk.sin();
         final UnivariateDerivative2 cos2Lk = sc2Lk.cos();
         final UnivariateDerivative2 sin2Lk = sc2Lk.sin();
         final UnivariateDerivative2 cos3Lk = sc3Lk.cos();
         final UnivariateDerivative2 sin3Lk = sc3Lk.sin();
         final UnivariateDerivative2 cos4Lk = sc4Lk.cos();
         final UnivariateDerivative2 sin4Lk = sc4Lk.sin();
 
         // h*cos(nLk), l*cos(nLk), h*sin(nLk) and l*sin(nLk)
         // n = 1
         final UnivariateDerivative2 hCosLk = h.multiply(cosLk);
         final UnivariateDerivative2 hSinLk = h.multiply(sinLk);
         final UnivariateDerivative2 lCosLk = l.multiply(cosLk);
         final UnivariateDerivative2 lSinLk = l.multiply(sinLk);
         // n = 2
         final UnivariateDerivative2 hCos2Lk = h.multiply(cos2Lk);
         final UnivariateDerivative2 hSin2Lk = h.multiply(sin2Lk);
         final UnivariateDerivative2 lCos2Lk = l.multiply(cos2Lk);
         final UnivariateDerivative2 lSin2Lk = l.multiply(sin2Lk);
         // n = 3
         final UnivariateDerivative2 hCos3Lk = h.multiply(cos3Lk);
         final UnivariateDerivative2 hSin3Lk = h.multiply(sin3Lk);
         final UnivariateDerivative2 lCos3Lk = l.multiply(cos3Lk);
         final UnivariateDerivative2 lSin3Lk = l.multiply(sin3Lk);
         // n = 4
         final UnivariateDerivative2 hCos4Lk = h.multiply(cos4Lk);
         final UnivariateDerivative2 hSin4Lk = h.multiply(sin4Lk);
         final UnivariateDerivative2 lCos4Lk = l.multiply(cos4Lk);
         final UnivariateDerivative2 lSin4Lk = l.multiply(sin4Lk);
 
         // 1 - (3 / 2)*sin²i
         final UnivariateDerivative2 om3o2xSinI2 = sinI2.multiply(1.5).negate().add(1.0);
 
         // Compute Differentials
         // δa
         final UnivariateDerivative2 dakT1 = b.multiply(2.0).multiply(om3o2xSinI2).multiply(lCosLk.add(hSinLk));
         final UnivariateDerivative2 dakT2 = b.multiply(sinI2).multiply(hSinLk.multiply(0.5).subtract(lCosLk.multiply(0.5)).
                                                                      add(cos2Lk).add(lCos3Lk.multiply(3.5)).add(hSin3Lk.multiply(3.5)));
         final UnivariateDerivative2 dak = dakT1.add(dakT2);
 
         // δh
         final UnivariateDerivative2 dhkT1 = b.multiply(om3o2xSinI2).multiply(sinLk.add(lSin2Lk.multiply(1.5)).subtract(hCos2Lk.multiply(1.5)));
         final UnivariateDerivative2 dhkT2 = b.multiply(sinI2).multiply(0.25).multiply(sinLk.subtract(sin3Lk.multiply(SEVEN_THIRD)).add(lSin2Lk.multiply(5.0)).
                                                                                     subtract(lSin4Lk.multiply(8.5)).add(hCos4Lk.multiply(8.5)).add(hCos2Lk));
         final UnivariateDerivative2 dhkT3 = lSin2Lk.multiply(cosI2).multiply(b).multiply(0.5).negate();
         final UnivariateDerivative2 dhk = dhkT1.subtract(dhkT2).add(dhkT3);
 
         // δl
         final UnivariateDerivative2 dlkT1 = b.multiply(om3o2xSinI2).multiply(cosLk.add(lCos2Lk.multiply(1.5)).add(hSin2Lk.multiply(1.5)));
         final UnivariateDerivative2 dlkT2 = b.multiply(sinI2).multiply(0.25).multiply(cosLk.negate().subtract(cos3Lk.multiply(SEVEN_THIRD)).subtract(hSin2Lk.multiply(5.0)).
                                                                                     subtract(lCos4Lk.multiply(8.5)).subtract(hSin4Lk.multiply(8.5)).add(lCos2Lk));
         final UnivariateDerivative2 dlkT3 = hSin2Lk.multiply(cosI2).multiply(b).multiply(0.5);
         final UnivariateDerivative2 dlk = dlkT1.subtract(dlkT2).add(dlkT3);
 
         // δλ
         final UnivariateDerivative2 dokT1 = b.negate().multiply(cosI);
         final UnivariateDerivative2 dokT2 = lSinLk.multiply(3.5).subtract(hCosLk.multiply(2.5)).subtract(sin2Lk.multiply(0.5)).
                                           subtract(lSin3Lk.multiply(SEVEN_SIXTH)).add(hCos3Lk.multiply(SEVEN_SIXTH));
         final UnivariateDerivative2 dok = dokT1.multiply(dokT2);
 
         // δi
         final UnivariateDerivative2 dik = b.multiply(sinI).multiply(cosI).multiply(0.5).
                         multiply(lCosLk.negate().add(hSinLk).add(cos2Lk).add(lCos3Lk.multiply(SEVEN_THIRD)).add(hSin3Lk.multiply(SEVEN_THIRD)));
 
         // δL
         final UnivariateDerivative2 dLkT1 = b.multiply(2.0).multiply(om3o2xSinI2).multiply(lSinLk.multiply(1.75).subtract(hCosLk.multiply(1.75)));
         final UnivariateDerivative2 dLkT2 = b.multiply(sinI2).multiply(3.0).multiply(hCosLk.multiply(SEVEN_24TH).negate().subtract(lSinLk.multiply(SEVEN_24TH)).
                                                                                    subtract(hCos3Lk.multiply(FN_72TH)).add(lSin3Lk.multiply(FN_72TH)).add(sin2Lk.multiply(0.25)));
         final UnivariateDerivative2 dLkT3 = b.multiply(cosI2).multiply(lSinLk.multiply(3.5).subtract(hCosLk.multiply(2.5)).subtract(sin2Lk.multiply(0.5)).
                                                                      subtract(lSin3Lk.multiply(SEVEN_SIXTH)).add(hCos3Lk.multiply(SEVEN_SIXTH)));
         final UnivariateDerivative2 dLk = dLkT1.add(dLkT2).add(dLkT3);
 
         // Final array
         final UnivariateDerivative2[] differentials = MathArrays.buildArray(a.getField(), 6);
         differentials[0] = dak.multiply(a);
         differentials[1] = dhk;
         differentials[2] = dlk;
         differentials[3] = dok;
         differentials[4] = dik;
         differentials[5] = dLk;
 
         return differentials;
     }
 
     /** {@inheritDoc} */
     protected double getMass(final AbsoluteDate date) {
         return mass;
     }
 
     /**
      * Get the Earth gravity coefficient used for GLONASS propagation.
      * @return the Earth gravity coefficient.
      */
     public static double getMU() {
         return GNSSConstants.GLONASS_MU;
     }
 
     /**
      * Gets the underlying GLONASS orbital elements.
      *
      * @return the underlying GLONASS orbital elements
      */
     public GLONASSOrbitalElements getGLONASSOrbitalElements() {
         return glonassOrbit;
     }
 
     /**
      * 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 GLONASS orbits.
      * @return the ECEF frame
      */
     public Frame getECEF() {
         return ecef;
     }
 
     /** {@inheritDoc} */
     public Frame getFrame() {
         return eci;
     }
 
     /** {@inheritDoc} */
     public void resetInitialState(final SpacecraftState state) {
         throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
     }
 
     /** {@inheritDoc} */
     protected void resetIntermediateState(final SpacecraftState state, final boolean forward) {
         throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
     }
 
     /** {@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, GNSSConstants.GLONASS_MU);
     }
 
 }