Examples with DateComponents org.orekit.time.DateComponents used on opensource projects

Example 96 with DateComponents

use of org.orekit.time.DateComponents in project Orekit by CS-SI.

the class DSSTPropagatorTest method getGEOState.

private SpacecraftState getGEOState() throws IllegalArgumentException, OrekitException {
    // No shadow at this date
    final AbsoluteDate initDate = new AbsoluteDate(new DateComponents(2003, 05, 21), new TimeComponents(1, 0, 0.), TimeScalesFactory.getUTC());
    final Orbit orbit = new EquinoctialOrbit(42164000, 10e-3, 10e-3, FastMath.tan(0.001745329) * FastMath.cos(2 * FastMath.PI / 3), FastMath.tan(0.001745329) * FastMath.sin(2 * FastMath.PI / 3), 0.1, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, 3.986004415E14);
    return new SpacecraftState(orbit);
}

Example 97 with DateComponents

use of org.orekit.time.DateComponents in project Orekit by CS-SI.

the class DSSTPropagatorTest method testImpulseManeuver.

@Test
public void testImpulseManeuver() throws OrekitException {
    final Orbit initialOrbit = new KeplerianOrbit(24532000.0, 0.72, 0.3, FastMath.PI, 0.4, 2.0, PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2008, 06, 23), new TimeComponents(14, 18, 37), TimeScalesFactory.getUTC()), 3.986004415e14);
    final double a = initialOrbit.getA();
    final double e = initialOrbit.getE();
    final double i = initialOrbit.getI();
    final double mu = initialOrbit.getMu();
    final double vApo = FastMath.sqrt(mu * (1 - e) / (a * (1 + e)));
    double dv = 0.99 * FastMath.tan(i) * vApo;
    // Set propagator with state
    setDSSTProp(new SpacecraftState(initialOrbit));
    // Add impulse maneuver
    dsstProp.setAttitudeProvider(new LofOffset(initialOrbit.getFrame(), LOFType.VVLH));
    dsstProp.addEventDetector(new ImpulseManeuver<NodeDetector>(new NodeDetector(initialOrbit, FramesFactory.getEME2000()), new Vector3D(dv, Vector3D.PLUS_J), 400.0));
    SpacecraftState propagated = dsstProp.propagate(initialOrbit.getDate().shiftedBy(8000));
    Assert.assertEquals(0.0028257, propagated.getI(), 1.0e-6);
}

Example 98 with DateComponents

use of org.orekit.time.DateComponents in project Orekit by CS-SI.

the class DSSTPropagatorTest method getLEOStatePropagatedBy30Minutes.

private SpacecraftState getLEOStatePropagatedBy30Minutes() throws IllegalArgumentException, OrekitException {
    final Vector3D position = new Vector3D(-6142438.668, 3492467.560, -25767.25680);
    final Vector3D velocity = new Vector3D(505.8479685, 942.7809215, 7435.922231);
    // Spring equinoxe 21st mars 2003 1h00m
    final AbsoluteDate initialDate = new AbsoluteDate(new DateComponents(2003, 03, 21), new TimeComponents(1, 0, 0.), TimeScalesFactory.getUTC());
    final CartesianOrbit osculatingOrbit = new CartesianOrbit(new PVCoordinates(position, velocity), FramesFactory.getTOD(IERSConventions.IERS_1996, false), initialDate, Constants.WGS84_EARTH_MU);
    // Adaptive step integrator
    // with a minimum step of 0.001 and a maximum step of 1000
    double minStep = 0.001;
    double maxstep = 1000.0;
    double positionTolerance = 10.0;
    OrbitType propagationType = OrbitType.EQUINOCTIAL;
    double[][] tolerances = NumericalPropagator.tolerances(positionTolerance, osculatingOrbit, propagationType);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxstep, tolerances[0], tolerances[1]);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(propagationType);
    NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(5, 5);
    ForceModel holmesFeatherstone = new HolmesFeatherstoneAttractionModel(FramesFactory.getITRF(IERSConventions.IERS_2010, true), provider);
    propagator.addForceModel(holmesFeatherstone);
    propagator.setInitialState(new SpacecraftState(osculatingOrbit));
    return propagator.propagate(new AbsoluteDate(initialDate, 1800.));
}

Example 99 with DateComponents

use of org.orekit.time.DateComponents in project Orekit by CS-SI.

the class TabulatedEphemerisTest method checkInterpolation.

private void checkInterpolation(StateFilter f, double expectedDP, double expectedDV) throws 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;
    final AbsoluteDate initDate = new AbsoluteDate(new DateComponents(2004, 01, 01), TimeComponents.H00, TimeScalesFactory.getUTC());
    final AbsoluteDate finalDate = new AbsoluteDate(new DateComponents(2004, 01, 02), TimeComponents.H00, TimeScalesFactory.getUTC());
    double deltaT = finalDate.durationFrom(initDate);
    Orbit transPar = new KeplerianOrbit(a, e, i, omega, OMEGA, lv, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, mu);
    int nbIntervals = 720;
    EcksteinHechlerPropagator eck = new EcksteinHechlerPropagator(transPar, mass, ae, mu, c20, c30, c40, c50, c60);
    AdditionalStateProvider provider = new AdditionalStateProvider() {

        public String getName() {
            return "dt";
        }

        public double[] getAdditionalState(SpacecraftState state) {
            return new double[] { state.getDate().durationFrom(initDate) };
        }
    };
    eck.addAdditionalStateProvider(provider);
    try {
        eck.addAdditionalStateProvider(provider);
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(OrekitMessages.ADDITIONAL_STATE_NAME_ALREADY_IN_USE, oe.getSpecifier());
    }
    List<SpacecraftState> tab = new ArrayList<SpacecraftState>(nbIntervals + 1);
    for (int j = 0; j <= nbIntervals; j++) {
        AbsoluteDate current = initDate.shiftedBy((j * deltaT) / nbIntervals);
        tab.add(f.filter(eck.propagate(current)));
    }
    try {
        new Ephemeris(tab, nbIntervals + 2);
        Assert.fail("an exception should have been thrown");
    } catch (MathIllegalArgumentException miae) {
    // expected
    }
    Ephemeris te = new Ephemeris(tab, 2);
    Assert.assertEquals(0.0, te.getMaxDate().durationFrom(finalDate), 1.0e-9);
    Assert.assertEquals(0.0, te.getMinDate().durationFrom(initDate), 1.0e-9);
    double maxP = 0;
    double maxV = 0;
    for (double dt = 0; dt < 3600; dt += 1) {
        AbsoluteDate date = initDate.shiftedBy(dt);
        CartesianOrbit c1 = (CartesianOrbit) eck.propagate(date).getOrbit();
        CartesianOrbit c2 = (CartesianOrbit) te.propagate(date).getOrbit();
        maxP = FastMath.max(maxP, Vector3D.distance(c1.getPVCoordinates().getPosition(), c2.getPVCoordinates().getPosition()));
        maxV = FastMath.max(maxV, Vector3D.distance(c1.getPVCoordinates().getVelocity(), c2.getPVCoordinates().getVelocity()));
    }
    Assert.assertEquals(expectedDP, maxP, 0.1 * expectedDP);
    Assert.assertEquals(expectedDV, maxV, 0.1 * expectedDV);
}

Example 100 with DateComponents

use of org.orekit.time.DateComponents in project Orekit by CS-SI.

the class TLETest method testBug77TooLargeSecondDerivative.

@Test
public void testBug77TooLargeSecondDerivative() throws OrekitException {
    try {
        TLE tle = new TLE(5555, 'U', 1971, 86, "J", 0, 908, new AbsoluteDate(new DateComponents(2012, 26), new TimeComponents(0.96078249 * Constants.JULIAN_DAY), TimeScalesFactory.getUTC()), taylorConvert(12.26882470, 1), taylorConvert(-0.00000004, 2), taylorConvert(0.99999e11, 3), 0.0075476, FastMath.toRadians(74.0161), FastMath.toRadians(328.9888), FastMath.toRadians(228.9750), FastMath.toRadians(30.6709), 80454, 0.01234e-9);
        tle.getLine1();
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(OrekitMessages.TLE_INVALID_PARAMETER, oe.getSpecifier());
        Assert.assertEquals(5555, ((Integer) oe.getParts()[0]).intValue());
        Assert.assertEquals("meanMotionSecondDerivative", oe.getParts()[1]);
    }
}