TLEPropagator.java

  1. /* Copyright 2002-2020 CS GROUP
  2.  * Licensed to CS GROUP (CS) under one or more
  3.  * contributor license agreements.  See the NOTICE file distributed with
  4.  * this work for additional information regarding copyright ownership.
  5.  * CS licenses this file to You under the Apache License, Version 2.0
  6.  * (the "License"); you may not use this file except in compliance with
  7.  * the License.  You may obtain a copy of the License at
  8.  *
  9.  *   http://www.apache.org/licenses/LICENSE-2.0
  10.  *
  11.  * Unless required by applicable law or agreed to in writing, software
  12.  * distributed under the License is distributed on an "AS IS" BASIS,
  13.  * WITHOUT WARRANTIES OR CONDITIONS OF ANY KIND, either express or implied.
  14.  * See the License for the specific language governing permissions and
  15.  * limitations under the License.
  16.  */
  17. package org.orekit.propagation.analytical.tle;

  19. import org.hipparchus.geometry.euclidean.threed.Vector3D;
  20. import org.hipparchus.util.FastMath;
  21. import org.hipparchus.util.MathUtils;
  22. import org.hipparchus.util.SinCos;
  23. import org.orekit.annotation.DefaultDataContext;
  24. import org.orekit.attitudes.Attitude;
  25. import org.orekit.attitudes.AttitudeProvider;
  26. import org.orekit.data.DataContext;
  27. import org.orekit.errors.OrekitException;
  28. import org.orekit.errors.OrekitMessages;
  29. import org.orekit.frames.Frame;
  30. import org.orekit.frames.Frames;
  31. import org.orekit.orbits.CartesianOrbit;
  32. import org.orekit.orbits.Orbit;
  33. import org.orekit.propagation.Propagator;
  34. import org.orekit.propagation.SpacecraftState;
  35. import org.orekit.propagation.analytical.AbstractAnalyticalPropagator;
  36. import org.orekit.time.AbsoluteDate;
  37. import org.orekit.time.TimeScale;
  38. import org.orekit.utils.PVCoordinates;


  41. /** This class provides elements to propagate TLE's.
  42.  * <p>
  43.  * The models used are SGP4 and SDP4, initially proposed by NORAD as the unique convenient
  44.  * propagator for TLE's. Inputs and outputs of this propagator are only suited for
  45.  * NORAD two lines elements sets, since it uses estimations and mean values appropriate
  46.  * for TLE's only.
  47.  * </p>
  48.  * <p>
  49.  * Deep- or near- space propagator is selected internally according to NORAD recommendations
  50.  * so that the user has not to worry about the used computation methods. One instance is created
  51.  * for each TLE (this instance can only be get using {@link #selectExtrapolator(TLE)} method,
  52.  * and can compute {@link PVCoordinates position and velocity coordinates} at any
  53.  * time. Maximum accuracy is guaranteed in a 24h range period before and after the provided
  54.  * TLE epoch (of course this accuracy is not really measurable nor predictable: according to
  55.  * <a href="https://www.celestrak.com/">CelesTrak</a>, the precision is close to one kilometer
  56.  * and error won't probably rise above 2 km).
  57.  * </p>
  58.  * <p>This implementation is largely inspired from the paper and source code <a
  59.  * href="https://www.celestrak.com/publications/AIAA/2006-6753/">Revisiting Spacetrack
  60.  * Report #3</a> and is fully compliant with its results and tests cases.</p>
  61.  * @author Felix R. Hoots, Ronald L. Roehrich, December 1980 (original fortran)
  62.  * @author David A. Vallado, Paul Crawford, Richard Hujsak, T.S. Kelso (C++ translation and improvements)
  63.  * @author Fabien Maussion (java translation)
  64.  * @see TLE
  65.  */
  66. public abstract class TLEPropagator extends AbstractAnalyticalPropagator {

  68.     // CHECKSTYLE: stop VisibilityModifier check

  70.     /** Initial state. */
  71.     protected final TLE tle;

  73.     /** UTC time scale. */
  74.     protected final TimeScale utc;

  76.     /** final RAAN. */
  77.     protected double xnode;

  79.     /** final semi major axis. */
  80.     protected double a;

  82.     /** final eccentricity. */
  83.     protected double e;

  85.     /** final inclination. */
  86.     protected double i;

  88.     /** final perigee argument. */
  89.     protected double omega;

  91.     /** L from SPTRCK #3. */
  92.     protected double xl;

  94.     /** original recovered semi major axis. */
  95.     protected double a0dp;

  97.     /** original recovered mean motion. */
  98.     protected double xn0dp;

  100.     /** cosinus original inclination. */
  101.     protected double cosi0;

  103.     /** cos io squared. */
  104.     protected double theta2;

  106.     /** sinus original inclination. */
  107.     protected double sini0;

  109.     /** common parameter for mean anomaly (M) computation. */
  110.     protected double xmdot;

  112.     /** common parameter for perigee argument (omega) computation. */
  113.     protected double omgdot;

  115.     /** common parameter for raan (OMEGA) computation. */
  116.     protected double xnodot;

  118.     /** original eccentricity squared. */
  119.     protected double e0sq;
  120.     /** 1 - e2. */
  121.     protected double beta02;

  123.     /** sqrt (1 - e2). */
  124.     protected double beta0;

  126.     /** perigee, expressed in KM and ALTITUDE. */
  127.     protected double perige;

  129.     /** eta squared. */
  130.     protected double etasq;

  132.     /** original eccentricity * eta. */
  133.     protected double eeta;

  135.     /** s* new value for the contant s. */
  136.     protected double s4;

  138.     /** tsi from SPTRCK #3. */
  139.     protected double tsi;

  141.     /** eta from SPTRCK #3. */
  142.     protected double eta;

  144.     /** coef for SGP C3 computation. */
  145.     protected double coef;

  147.     /** coef for SGP C5 computation. */
  148.     protected double coef1;

  150.     /** C1 from SPTRCK #3. */
  151.     protected double c1;

  153.     /** C2 from SPTRCK #3. */
  154.     protected double c2;

  156.     /** C4 from SPTRCK #3. */
  157.     protected double c4;

  159.     /** common parameter for raan (OMEGA) computation. */
  160.     protected double xnodcf;

  162.     /** 3/2 * C1. */
  163.     protected double t2cof;

  165.     // CHECKSTYLE: resume VisibilityModifier check

  167.     /** TLE frame. */
  168.     private final Frame teme;

  170.     /** Spacecraft mass (kg). */
  171.     private final double mass;

  173.     /** Protected constructor for derived classes.
  174.      *
  175.      * <p>This constructor uses the {@link DataContext#getDefault() default data context}.
  176.      *
  177.      * @param initialTLE the unique TLE to propagate
  178.      * @param attitudeProvider provider for attitude computation
  179.      * @param mass spacecraft mass (kg)
  180.      * @see #TLEPropagator(TLE, AttitudeProvider, double, Frame)
  181.      */
  182.     @DefaultDataContext
  183.     protected TLEPropagator(final TLE initialTLE, final AttitudeProvider attitudeProvider,
  184.                             final double mass) {
  185.         this(initialTLE, attitudeProvider, mass,
  186.                 DataContext.getDefault().getFrames().getTEME());
  187.     }

  189.     /** Protected constructor for derived classes.
  190.      * @param initialTLE the unique TLE to propagate
  191.      * @param attitudeProvider provider for attitude computation
  192.      * @param mass spacecraft mass (kg)
  193.      * @param teme the TEME frame to use for propagation.
  194.      * @since 10.1
  195.      */
  196.     protected TLEPropagator(final TLE initialTLE,
  197.                             final AttitudeProvider attitudeProvider,
  198.                             final double mass,
  199.                             final Frame teme) {
  200.         super(attitudeProvider);
  201.         setStartDate(initialTLE.getDate());
  202.         this.tle  = initialTLE;
  203.         this.teme = teme;
  204.         this.mass = mass;
  205.         this.utc = initialTLE.getUtc();
  206.         initializeCommons();
  207.         sxpInitialize();
  208.         // set the initial state
  209.         final Orbit orbit = propagateOrbit(initialTLE.getDate());
  210.         final Attitude attitude = attitudeProvider.getAttitude(orbit, orbit.getDate(), orbit.getFrame());
  211.         super.resetInitialState(new SpacecraftState(orbit, attitude, mass));
  212.     }

  214.     /** Selects the extrapolator to use with the selected TLE.
  215.      *
  216.      * <p>This method uses the {@link DataContext#getDefault() default data context}.
  217.      *
  218.      * @param tle the TLE to propagate.
  219.      * @return the correct propagator.
  220.      * @see #selectExtrapolator(TLE, Frames)
  221.      */
  222.     @DefaultDataContext
  223.     public static TLEPropagator selectExtrapolator(final TLE tle) {
  224.         return selectExtrapolator(tle, DataContext.getDefault().getFrames());
  225.     }

  227.     /** Selects the extrapolator to use with the selected TLE.
  228.      * @param tle the TLE to propagate.
  229.      * @param frames set of Frames to use in the propagator.
  230.      * @return the correct propagator.
  231.      * @since 10.1
  232.      */
  233.     public static TLEPropagator selectExtrapolator(final TLE tle, final Frames frames) {
  234.         return selectExtrapolator(
  235.                 tle,
  236.                 Propagator.getDefaultLaw(frames),
  237.                 DEFAULT_MASS,
  238.                 frames.getTEME());
  239.     }

  241.     /** Selects the extrapolator to use with the selected TLE.
  242.      *
  243.      * <p>This method uses the {@link DataContext#getDefault() default data context}.
  244.      *
  245.      * @param tle the TLE to propagate.
  246.      * @param attitudeProvider provider for attitude computation
  247.      * @param mass spacecraft mass (kg)
  248.      * @return the correct propagator.
  249.      * @see #selectExtrapolator(TLE, AttitudeProvider, double, Frame)
  250.      */
  251.     @DefaultDataContext
  252.     public static TLEPropagator selectExtrapolator(final TLE tle, final AttitudeProvider attitudeProvider,
  253.                                                    final double mass) {
  254.         return selectExtrapolator(tle, attitudeProvider, mass,
  255.                 DataContext.getDefault().getFrames().getTEME());
  256.     }

  258.     /** Selects the extrapolator to use with the selected TLE.
  259.      * @param tle the TLE to propagate.
  260.      * @param attitudeProvider provider for attitude computation
  261.      * @param mass spacecraft mass (kg)
  262.      * @param teme the TEME frame to use for propagation.
  263.      * @return the correct propagator.
  264.      * @since 10.1
  265.      */
  266.     public static TLEPropagator selectExtrapolator(final TLE tle,
  267.                                                    final AttitudeProvider attitudeProvider,
  268.                                                    final double mass,
  269.                                                    final Frame teme) {

  271.         final double a1 = FastMath.pow( TLEConstants.XKE / (tle.getMeanMotion() * 60.0), TLEConstants.TWO_THIRD);
  272.         final double cosi0 = FastMath.cos(tle.getI());
  273.         final double temp = TLEConstants.CK2 * 1.5 * (3 * cosi0 * cosi0 - 1.0) *
  274.                             FastMath.pow(1.0 - tle.getE() * tle.getE(), -1.5);
  275.         final double delta1 = temp / (a1 * a1);
  276.         final double a0 = a1 * (1.0 - delta1 * (TLEConstants.ONE_THIRD + delta1 * (delta1 * 134.0 / 81.0 + 1.0)));
  277.         final double delta0 = temp / (a0 * a0);

  279.         // recover original mean motion :
  280.         final double xn0dp = tle.getMeanMotion() * 60.0 / (delta0 + 1.0);

  282.         // Period >= 225 minutes is deep space
  283.         if (MathUtils.TWO_PI / (xn0dp * TLEConstants.MINUTES_PER_DAY) >= (1.0 / 6.4)) {
  284.             return new DeepSDP4(tle, attitudeProvider, mass, teme);
  285.         } else {
  286.             return new SGP4(tle, attitudeProvider, mass, teme);
  287.         }
  288.     }

  290.     /** Get the Earth gravity coefficient used for TLE propagation.
  291.      * @return the Earth gravity coefficient.
  292.      */
  293.     public static double getMU() {
  294.         return TLEConstants.MU;
  295.     }

  297.     /** Get the extrapolated position and velocity from an initial TLE.
  298.      * @param date the final date
  299.      * @return the final PVCoordinates
  300.      */
  301.     public PVCoordinates getPVCoordinates(final AbsoluteDate date) {

  303.         sxpPropagate(date.durationFrom(tle.getDate()) / 60.0);

  305.         // Compute PV with previous calculated parameters
  306.         return computePVCoordinates();
  307.     }

  309.     /** Computation of the first commons parameters.
  310.      */
  311.     private void initializeCommons() {

  313.         // Sine and cosine of inclination
  314.         final SinCos scI0 = FastMath.sinCos(tle.getI());

  316.         final double a1 = FastMath.pow(TLEConstants.XKE / (tle.getMeanMotion() * 60.0), TLEConstants.TWO_THIRD);
  317.         cosi0 = scI0.cos();
  318.         theta2 = cosi0 * cosi0;
  319.         final double x3thm1 = 3.0 * theta2 - 1.0;
  320.         e0sq = tle.getE() * tle.getE();
  321.         beta02 = 1.0 - e0sq;
  322.         beta0 = FastMath.sqrt(beta02);
  323.         final double tval = TLEConstants.CK2 * 1.5 * x3thm1 / (beta0 * beta02);
  324.         final double delta1 = tval / (a1 * a1);
  325.         final double a0 = a1 * (1.0 - delta1 * (TLEConstants.ONE_THIRD + delta1 * (1.0 + 134.0 / 81.0 * delta1)));
  326.         final double delta0 = tval / (a0 * a0);

  328.         // recover original mean motion and semi-major axis :
  329.         xn0dp = tle.getMeanMotion() * 60.0 / (delta0 + 1.0);
  330.         a0dp = a0 / (1.0 - delta0);

  332.         // Values of s and qms2t :
  333.         s4 = TLEConstants.S;  // unmodified value for s
  334.         double q0ms24 = TLEConstants.QOMS2T; // unmodified value for q0ms2T

  336.         perige = (a0dp * (1 - tle.getE()) - TLEConstants.NORMALIZED_EQUATORIAL_RADIUS) * TLEConstants.EARTH_RADIUS; // perige

  338.         //  For perigee below 156 km, the values of s and qoms2t are changed :
  339.         if (perige < 156.0) {
  340.             if (perige <= 98.0) {
  341.                 s4 = 20.0;
  342.             } else {
  343.                 s4 = perige - 78.0;
  344.             }
  345.             final double temp_val = (120.0 - s4) * TLEConstants.NORMALIZED_EQUATORIAL_RADIUS / TLEConstants.EARTH_RADIUS;
  346.             final double temp_val_squared = temp_val * temp_val;
  347.             q0ms24 = temp_val_squared * temp_val_squared;
  348.             s4 = s4 / TLEConstants.EARTH_RADIUS + TLEConstants.NORMALIZED_EQUATORIAL_RADIUS; // new value for q0ms2T and s
  349.         }

  351.         final double pinv = 1.0 / (a0dp * beta02);
  352.         final double pinvsq = pinv * pinv;
  353.         tsi = 1.0 / (a0dp - s4);
  354.         eta = a0dp * tle.getE() * tsi;
  355.         etasq = eta * eta;
  356.         eeta = tle.getE() * eta;

  358.         final double psisq = FastMath.abs(1.0 - etasq); // abs because pow 3.5 needs positive value
  359.         final double tsi_squared = tsi * tsi;
  360.         coef = q0ms24 * tsi_squared * tsi_squared;
  361.         coef1 = coef / FastMath.pow(psisq, 3.5);

  363.         // C2 and C1 coefficients computation :
  364.         c2 = coef1 * xn0dp * (a0dp * (1.0 + 1.5 * etasq + eeta * (4.0 + etasq)) +
  365.              0.75 * TLEConstants.CK2 * tsi / psisq * x3thm1 * (8.0 + 3.0 * etasq * (8.0 + etasq)));
  366.         c1 = tle.getBStar() * c2;
  367.         sini0 = scI0.sin();

  369.         final double x1mth2 = 1.0 - theta2;

  371.         // C4 coefficient computation :
  372.         c4 = 2.0 * xn0dp * coef1 * a0dp * beta02 * (eta * (2.0 + 0.5 * etasq) +
  373.              tle.getE() * (0.5 + 2.0 * etasq) -
  374.              2 * TLEConstants.CK2 * tsi / (a0dp * psisq) *
  375.              (-3.0 * x3thm1 * (1.0 - 2.0 * eeta + etasq * (1.5 - 0.5 * eeta)) +
  376.               0.75 * x1mth2 * (2.0 * etasq - eeta * (1.0 + etasq)) * FastMath.cos(2.0 * tle.getPerigeeArgument())));

  378.         final double theta4 = theta2 * theta2;
  379.         final double temp1 = 3 * TLEConstants.CK2 * pinvsq * xn0dp;
  380.         final double temp2 = temp1 * TLEConstants.CK2 * pinvsq;
  381.         final double temp3 = 1.25 * TLEConstants.CK4 * pinvsq * pinvsq * xn0dp;

  383.         // atmospheric and gravitation coefs :(Mdf and OMEGAdf)
  384.         xmdot = xn0dp +
  385.                 0.5 * temp1 * beta0 * x3thm1 +
  386.                 0.0625 * temp2 * beta0 * (13.0 - 78.0 * theta2 + 137.0 * theta4);

  388.         final double x1m5th = 1.0 - 5.0 * theta2;

  390.         omgdot = -0.5 * temp1 * x1m5th +
  391.                  0.0625 * temp2 * (7.0 - 114.0 * theta2 + 395.0 * theta4) +
  392.                  temp3 * (3.0 - 36.0 * theta2 + 49.0 * theta4);

  394.         final double xhdot1 = -temp1 * cosi0;

  396.         xnodot = xhdot1 + (0.5 * temp2 * (4.0 - 19.0 * theta2) + 2.0 * temp3 * (3.0 - 7.0 * theta2)) * cosi0;
  397.         xnodcf = 3.5 * beta02 * xhdot1 * c1;
  398.         t2cof = 1.5 * c1;

  400.     }

  402.     /** Retrieves the position and velocity.
  403.      * @return the computed PVCoordinates.
  404.      */
  405.     private PVCoordinates computePVCoordinates() {

  407.         // Sine and cosine of final perigee argument
  408.         final SinCos scOmega = FastMath.sinCos(omega);

  410.         // Long period periodics
  411.         final double axn = e * scOmega.cos();
  412.         double temp = 1.0 / (a * (1.0 - e * e));
  413.         final double xlcof = 0.125 * TLEConstants.A3OVK2 * sini0 * (3.0 + 5.0 * cosi0) / (1.0 + cosi0);
  414.         final double aycof = 0.25 * TLEConstants.A3OVK2 * sini0;
  415.         final double xll = temp * xlcof * axn;
  416.         final double aynl = temp * aycof;
  417.         final double xlt = xl + xll;
  418.         final double ayn = e * scOmega.sin() + aynl;
  419.         final double elsq = axn * axn + ayn * ayn;
  420.         final double capu = MathUtils.normalizeAngle(xlt - xnode, FastMath.PI);
  421.         double epw = capu;
  422.         double ecosE = 0;
  423.         double esinE = 0;
  424.         double sinEPW = 0;
  425.         double cosEPW = 0;

  427.         // Dundee changes:  items dependent on cosio get recomputed:
  428.         final double cosi0Sq = cosi0 * cosi0;
  429.         final double x3thm1 = 3.0 * cosi0Sq - 1.0;
  430.         final double x1mth2 = 1.0 - cosi0Sq;
  431.         final double x7thm1 = 7.0 * cosi0Sq - 1.0;

  433.         if (e > (1 - 1e-6)) {
  434.             throw new OrekitException(OrekitMessages.TOO_LARGE_ECCENTRICITY_FOR_PROPAGATION_MODEL, e);
  435.         }

  437.         // Solve Kepler's' Equation.
  438.         final double newtonRaphsonEpsilon = 1e-12;
  439.         for (int j = 0; j < 10; j++) {

  441.             boolean doSecondOrderNewtonRaphson = true;

  443.             final SinCos scEPW = FastMath.sinCos(epw);
  444.             sinEPW = scEPW.sin();
  445.             cosEPW = scEPW.cos();
  446.             ecosE = axn * cosEPW + ayn * sinEPW;
  447.             esinE = axn * sinEPW - ayn * cosEPW;
  448.             final double f = capu - epw + esinE;
  449.             if (FastMath.abs(f) < newtonRaphsonEpsilon) {
  450.                 break;
  451.             }
  452.             final double fdot = 1.0 - ecosE;
  453.             double delta_epw = f / fdot;
  454.             if (j == 0) {
  455.                 final double maxNewtonRaphson = 1.25 * FastMath.abs(e);
  456.                 doSecondOrderNewtonRaphson = false;
  457.                 if (delta_epw > maxNewtonRaphson) {
  458.                     delta_epw = maxNewtonRaphson;
  459.                 } else if (delta_epw < -maxNewtonRaphson) {
  460.                     delta_epw = -maxNewtonRaphson;
  461.                 } else {
  462.                     doSecondOrderNewtonRaphson = true;
  463.                 }
  464.             }
  465.             if (doSecondOrderNewtonRaphson) {
  466.                 delta_epw = f / (fdot + 0.5 * esinE * delta_epw);
  467.             }
  468.             epw += delta_epw;
  469.         }

  471.         // Short period preliminary quantities
  472.         temp = 1.0 - elsq;
  473.         final double pl = a * temp;
  474.         final double r = a * (1.0 - ecosE);
  475.         double temp2 = a / r;
  476.         final double betal = FastMath.sqrt(temp);
  477.         temp = esinE / (1.0 + betal);
  478.         final double cosu = temp2 * (cosEPW - axn + ayn * temp);
  479.         final double sinu = temp2 * (sinEPW - ayn - axn * temp);
  480.         final double u = FastMath.atan2(sinu, cosu);
  481.         final double sin2u = 2.0 * sinu * cosu;
  482.         final double cos2u = 2.0 * cosu * cosu - 1.0;
  483.         final double temp1 = TLEConstants.CK2 / pl;
  484.         temp2 = temp1 / pl;

  486.         // Update for short periodics
  487.         final double rk = r * (1.0 - 1.5 * temp2 * betal * x3thm1) + 0.5 * temp1 * x1mth2 * cos2u;
  488.         final double uk = u - 0.25 * temp2 * x7thm1 * sin2u;
  489.         final double xnodek = xnode + 1.5 * temp2 * cosi0 * sin2u;
  490.         final double xinck = i + 1.5 * temp2 * cosi0 * sini0 * cos2u;

  492.         // Orientation vectors
  493.         final SinCos scuk   = FastMath.sinCos(uk);
  494.         final SinCos scik   = FastMath.sinCos(xinck);
  495.         final SinCos scnok  = FastMath.sinCos(xnodek);
  496.         final double sinuk  = scuk.sin();
  497.         final double cosuk  = scuk.cos();
  498.         final double sinik  = scik.sin();
  499.         final double cosik  = scik.cos();
  500.         final double sinnok = scnok.sin();
  501.         final double cosnok = scnok.cos();
  502.         final double xmx = -sinnok * cosik;
  503.         final double xmy = cosnok * cosik;
  504.         final double ux  = xmx * sinuk + cosnok * cosuk;
  505.         final double uy  = xmy * sinuk + sinnok * cosuk;
  506.         final double uz  = sinik * sinuk;

  508.         // Position and velocity
  509.         final double cr = 1000 * rk * TLEConstants.EARTH_RADIUS;
  510.         final Vector3D pos = new Vector3D(cr * ux, cr * uy, cr * uz);

  512.         final double rdot   = TLEConstants.XKE * FastMath.sqrt(a) * esinE / r;
  513.         final double rfdot  = TLEConstants.XKE * FastMath.sqrt(pl) / r;
  514.         final double xn     = TLEConstants.XKE / (a * FastMath.sqrt(a));
  515.         final double rdotk  = rdot - xn * temp1 * x1mth2 * sin2u;
  516.         final double rfdotk = rfdot + xn * temp1 * (x1mth2 * cos2u + 1.5 * x3thm1);
  517.         final double vx     = xmx * cosuk - cosnok * sinuk;
  518.         final double vy     = xmy * cosuk - sinnok * sinuk;
  519.         final double vz     = sinik * cosuk;

  521.         final double cv = 1000.0 * TLEConstants.EARTH_RADIUS / 60.0;
  522.         final Vector3D vel = new Vector3D(cv * (rdotk * ux + rfdotk * vx),
  523.                                           cv * (rdotk * uy + rfdotk * vy),
  524.                                           cv * (rdotk * uz + rfdotk * vz));

  526.         return new PVCoordinates(pos, vel);

  528.     }

  530.     /** Initialization proper to each propagator (SGP or SDP).
  531.      */
  532.     protected abstract void sxpInitialize();

  534.     /** Propagation proper to each propagator (SGP or SDP).
  535.      * @param t the offset from initial epoch (min)
  536.      */
  537.     protected abstract void sxpPropagate(double t);

  539.     /** {@inheritDoc} */
  540.     public void resetInitialState(final SpacecraftState state) {
  541.         throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
  542.     }

  544.     /** {@inheritDoc} */
  545.     protected void resetIntermediateState(final SpacecraftState state, final boolean forward) {
  546.         throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
  547.     }

  549.     /** {@inheritDoc} */
  550.     protected double getMass(final AbsoluteDate date) {
  551.         return mass;
  552.     }

  554.     /** {@inheritDoc} */
  555.     protected Orbit propagateOrbit(final AbsoluteDate date) {
  556.         return new CartesianOrbit(getPVCoordinates(date), teme, date, TLEConstants.MU);
  557.     }

  559.     /** Get the underlying TLE.
  560.      * @return underlying TLE
  561.      */
  562.     public TLE getTLE() {
  563.         return tle;
  564.     }

  566.     /** {@inheritDoc} */
  567.     public Frame getFrame() {
  568.         return teme;
  569.     }

  571. }
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