Examples with EcksteinHechlerPropagator org.orekit.propagation.analytical.EcksteinHechlerPropagator used on opensource projects

Example 1 with EcksteinHechlerPropagator

use of org.orekit.propagation.analytical.EcksteinHechlerPropagator in project Orekit by CS-SI.

the class BoxAndSolarArraySpacecraftTest method setUp.

@Before
public void setUp() {
    try {
        Utils.setDataRoot("regular-data");
        mu = 3.9860047e14;
        double ae = 6.378137e6;
        double c20 = -1.08263e-3;
        double c30 = 2.54e-6;
        double c40 = 1.62e-6;
        double c50 = 2.3e-7;
        double c60 = -5.5e-7;
        AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 7, 1), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
        // Satellite position as circular parameters, raan chosen to have sun elevation with
        // respect to orbit plane roughly evolving roughly from 15 to 15.2 degrees in the test range
        Orbit circ = new CircularOrbit(7178000.0, 0.5e-4, -0.5e-4, FastMath.toRadians(50.), FastMath.toRadians(280), FastMath.toRadians(10.0), PositionAngle.MEAN, FramesFactory.getEME2000(), date, mu);
        propagator = new EcksteinHechlerPropagator(circ, new LofOffset(circ.getFrame(), LOFType.VVLH), ae, mu, c20, c30, c40, c50, c60);
    } catch (OrekitException oe) {
        Assert.fail(oe.getLocalizedMessage());
    }
}

Example 2 with EcksteinHechlerPropagator

use of org.orekit.propagation.analytical.EcksteinHechlerPropagator in project Orekit by CS-SI.

the class TrackCorridor method createPropagator.

/**
 * Create an orbit propagator for a circular orbit
 * @param a  semi-major axis (m)
 * @param ex e cos(ω), first component of circular eccentricity vector
 * @param ey e sin(ω), second component of circular eccentricity vector
 * @param i inclination (rad)
 * @param raan right ascension of ascending node (Ω, rad)
 * @param alpha  an + ω, mean latitude argument (rad)
 * @param date date of the orbital parameters
 * @return an orbit propagator
 * @exception OrekitException if propagator cannot be built
 */
private Propagator createPropagator(final AbsoluteDate date, final double a, final double ex, final double ey, final double i, final double raan, final double alpha) throws OrekitException {
    // create orbit
    Orbit initialOrbit = new CircularOrbit(a, ex, ey, i, raan, alpha, PositionAngle.MEAN, FramesFactory.getEME2000(), date, Constants.EIGEN5C_EARTH_MU);
    // create propagator
    Propagator propagator = new EcksteinHechlerPropagator(initialOrbit, new LofOffset(initialOrbit.getFrame(), LOFType.TNW), Constants.EIGEN5C_EARTH_EQUATORIAL_RADIUS, Constants.EIGEN5C_EARTH_MU, Constants.EIGEN5C_EARTH_C20, Constants.EIGEN5C_EARTH_C30, Constants.EIGEN5C_EARTH_C40, Constants.EIGEN5C_EARTH_C50, Constants.EIGEN5C_EARTH_C60);
    return propagator;
}

Example 3 with EcksteinHechlerPropagator

use of org.orekit.propagation.analytical.EcksteinHechlerPropagator in project Orekit by CS-SI.

the class CircularOrbitTest method doTestInterpolation.

private void doTestInterpolation(boolean useDerivatives, double shiftErrorWithin, double interpolationErrorWithin, double shiftErrorSlightlyPast, double interpolationErrorSlightlyPast, double shiftErrorFarPast, double interpolationErrorFarPast) throws OrekitException {
    final double ehMu = 3.9860047e14;
    final double ae = 6.378137e6;
    final double c20 = -1.08263e-3;
    final double c30 = 2.54e-6;
    final double c40 = 1.62e-6;
    final double c50 = 2.3e-7;
    final double c60 = -5.5e-7;
    final AbsoluteDate date = AbsoluteDate.J2000_EPOCH.shiftedBy(584.);
    final Vector3D position = new Vector3D(3220103., 69623., 6449822.);
    final Vector3D velocity = new Vector3D(6414.7, -2006., -3180.);
    final CircularOrbit initialOrbit = new CircularOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), date, ehMu);
    EcksteinHechlerPropagator propagator = new EcksteinHechlerPropagator(initialOrbit, ae, ehMu, c20, c30, c40, c50, c60);
    // set up a 5 points sample
    List<Orbit> sample = new ArrayList<Orbit>();
    for (double dt = 0; dt < 300.0; dt += 60.0) {
        Orbit orbit = propagator.propagate(date.shiftedBy(dt)).getOrbit();
        if (!useDerivatives) {
            // remove derivatives
            double[] stateVector = new double[6];
            orbit.getType().mapOrbitToArray(orbit, PositionAngle.TRUE, stateVector, null);
            orbit = orbit.getType().mapArrayToOrbit(stateVector, null, PositionAngle.TRUE, orbit.getDate(), orbit.getMu(), orbit.getFrame());
        }
        sample.add(orbit);
    }
    // well inside the sample, interpolation should be much better than Keplerian shift
    double maxShiftError = 0;
    double maxInterpolationError = 0;
    for (double dt = 0; dt < 241.0; dt += 1.0) {
        AbsoluteDate t = initialOrbit.getDate().shiftedBy(dt);
        Vector3D shifted = initialOrbit.shiftedBy(dt).getPVCoordinates().getPosition();
        Vector3D interpolated = initialOrbit.interpolate(t, sample).getPVCoordinates().getPosition();
        Vector3D propagated = propagator.propagate(t).getPVCoordinates().getPosition();
        maxShiftError = FastMath.max(maxShiftError, shifted.subtract(propagated).getNorm());
        maxInterpolationError = FastMath.max(maxInterpolationError, interpolated.subtract(propagated).getNorm());
    }
    Assert.assertEquals(shiftErrorWithin, maxShiftError, 0.01 * shiftErrorWithin);
    Assert.assertEquals(interpolationErrorWithin, maxInterpolationError, 0.01 * interpolationErrorWithin);
    // slightly past sample end, interpolation should quickly increase, but remain reasonable
    maxShiftError = 0;
    maxInterpolationError = 0;
    for (double dt = 240; dt < 300.0; dt += 1.0) {
        AbsoluteDate t = initialOrbit.getDate().shiftedBy(dt);
        Vector3D shifted = initialOrbit.shiftedBy(dt).getPVCoordinates().getPosition();
        Vector3D interpolated = initialOrbit.interpolate(t, sample).getPVCoordinates().getPosition();
        Vector3D propagated = propagator.propagate(t).getPVCoordinates().getPosition();
        maxShiftError = FastMath.max(maxShiftError, shifted.subtract(propagated).getNorm());
        maxInterpolationError = FastMath.max(maxInterpolationError, interpolated.subtract(propagated).getNorm());
    }
    Assert.assertEquals(shiftErrorSlightlyPast, maxShiftError, 0.01 * shiftErrorSlightlyPast);
    Assert.assertEquals(interpolationErrorSlightlyPast, maxInterpolationError, 0.01 * interpolationErrorSlightlyPast);
    // far past sample end, interpolation should become really wrong
    maxShiftError = 0;
    maxInterpolationError = 0;
    for (double dt = 300; dt < 1000; dt += 1.0) {
        AbsoluteDate t = initialOrbit.getDate().shiftedBy(dt);
        Vector3D shifted = initialOrbit.shiftedBy(dt).getPVCoordinates().getPosition();
        Vector3D interpolated = initialOrbit.interpolate(t, sample).getPVCoordinates().getPosition();
        Vector3D propagated = propagator.propagate(t).getPVCoordinates().getPosition();
        maxShiftError = FastMath.max(maxShiftError, shifted.subtract(propagated).getNorm());
        maxInterpolationError = FastMath.max(maxInterpolationError, interpolated.subtract(propagated).getNorm());
    }
    Assert.assertEquals(shiftErrorFarPast, maxShiftError, 0.01 * shiftErrorFarPast);
    Assert.assertEquals(interpolationErrorFarPast, maxInterpolationError, 0.01 * interpolationErrorFarPast);
}

Example 4 with EcksteinHechlerPropagator

use of org.orekit.propagation.analytical.EcksteinHechlerPropagator in project Orekit by CS-SI.

the class SpacecraftStateTest method setUp.

@Before
public void setUp() {
    try {
        Utils.setDataRoot("regular-data");
        double mu = 3.9860047e14;
        double ae = 6.378137e6;
        double c20 = -1.08263e-3;
        double c30 = 2.54e-6;
        double c40 = 1.62e-6;
        double c50 = 2.3e-7;
        double c60 = -5.5e-7;
        mass = 2500;
        double a = 7187990.1979844316;
        double e = 0.5e-4;
        double i = 1.7105407051081795;
        double omega = 1.9674147913622104;
        double OMEGA = FastMath.toRadians(261);
        double lv = 0;
        AbsoluteDate date = new AbsoluteDate(new DateComponents(2004, 01, 01), TimeComponents.H00, TimeScalesFactory.getUTC());
        orbit = new KeplerianOrbit(a, e, i, omega, OMEGA, lv, PositionAngle.TRUE, FramesFactory.getEME2000(), date, mu);
        OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
        attitudeLaw = new BodyCenterPointing(orbit.getFrame(), earth);
        propagator = new EcksteinHechlerPropagator(orbit, attitudeLaw, mass, ae, mu, c20, c30, c40, c50, c60);
    } catch (OrekitException oe) {
        Assert.fail(oe.getLocalizedMessage());
    }
}

Example 5 with EcksteinHechlerPropagator

use of org.orekit.propagation.analytical.EcksteinHechlerPropagator in project Orekit by CS-SI.

the class EphemerisEventsTest method buildEphem.

private Ephemeris buildEphem(OrbitType type) throws IllegalArgumentException, OrekitException {
    double mass = 2500;
    double a = 7187990.1979844316;
    double e = 0.5e-4;
    double i = 1.7105407051081795;
    double omega = 1.9674147913622104;
    double OMEGA = FastMath.toRadians(261);
    double lv = 0;
    double mu = 3.9860047e14;
    double ae = 6.378137e6;
    double c20 = -1.08263e-3;
    double c30 = 2.54e-6;
    double c40 = 1.62e-6;
    double c50 = 2.3e-7;
    double c60 = -5.5e-7;
    double deltaT = finalDate.durationFrom(initDate);
    Orbit transPar = new KeplerianOrbit(a, e, i, omega, OMEGA, lv, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, mu);
    int nbIntervals = 720;
    Propagator propagator = new EcksteinHechlerPropagator(transPar, mass, ae, mu, c20, c30, c40, c50, c60);
    List<SpacecraftState> tab = new ArrayList<SpacecraftState>(nbIntervals + 1);
    for (int j = 0; j <= nbIntervals; j++) {
        SpacecraftState state = propagator.propagate(initDate.shiftedBy((j * deltaT) / nbIntervals));
        tab.add(new SpacecraftState(type.convertType(state.getOrbit()), state.getAttitude(), state.getMass()));
    }
    return new Ephemeris(tab, 2);
}