Examples with EquinoctialOrbit org.orekit.orbits.EquinoctialOrbit used on opensource projects

Example 6 with EquinoctialOrbit

use of org.orekit.orbits.EquinoctialOrbit in project Orekit by CS-SI.

the class DSSTPropagatorTest method getLEOState.

private SpacecraftState getLEOState() 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 initDate = new AbsoluteDate(new DateComponents(2003, 03, 21), new TimeComponents(1, 0, 0.), TimeScalesFactory.getUTC());
    return new SpacecraftState(new EquinoctialOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), initDate, 3.986004415E14));
}

Example 7 with EquinoctialOrbit

use of org.orekit.orbits.EquinoctialOrbit in project Orekit by CS-SI.

the class DSSTPropagator method computeOsculatingOrbit.

/**
 * Compute osculating state from mean state.
 * <p>
 * Compute and add the short periodic variation to the mean {@link SpacecraftState}.
 * </p>
 * @param meanState initial mean state
 * @param shortPeriodTerms short period terms
 * @return osculating state
 * @throws OrekitException if the computation of the short-periodic variation fails
 */
private static EquinoctialOrbit computeOsculatingOrbit(final SpacecraftState meanState, final List<ShortPeriodTerms> shortPeriodTerms) throws OrekitException {
    final double[] mean = new double[6];
    final double[] meanDot = new double[6];
    OrbitType.EQUINOCTIAL.mapOrbitToArray(meanState.getOrbit(), PositionAngle.MEAN, mean, meanDot);
    final double[] y = mean.clone();
    for (final ShortPeriodTerms spt : shortPeriodTerms) {
        final double[] shortPeriodic = spt.value(meanState.getOrbit());
        for (int i = 0; i < shortPeriodic.length; i++) {
            y[i] += shortPeriodic[i];
        }
    }
    return (EquinoctialOrbit) OrbitType.EQUINOCTIAL.mapArrayToOrbit(y, meanDot, PositionAngle.MEAN, meanState.getDate(), meanState.getMu(), meanState.getFrame());
}

Example 8 with EquinoctialOrbit

use of org.orekit.orbits.EquinoctialOrbit in project Orekit by CS-SI.

the class DSSTPropagator method computeOsculatingState.

/**
 * Conversion from mean to osculating orbit.
 * <p>
 * Compute osculating state <b>in a DSST sense</b>, corresponding to the
 * mean SpacecraftState in input, and according to the Force models taken
 * into account.
 * </p><p>
 * Since the osculating state is obtained by adding short-periodic variation
 * of each force model, the resulting output will depend on the
 * force models parameterized in input.
 * </p>
 * @param mean Mean state to convert
 * @param forces Forces to take into account
 * @param attitudeProvider attitude provider (may be null if there are no Gaussian force models
 * like atmospheric drag, radiation pressure or specific user-defined models)
 * @return osculating state in a DSST sense
 * @throws OrekitException if computation of short periodics fails
 */
public static SpacecraftState computeOsculatingState(final SpacecraftState mean, final AttitudeProvider attitudeProvider, final Collection<DSSTForceModel> forces) throws OrekitException {
    // Create the auxiliary object
    final AuxiliaryElements aux = new AuxiliaryElements(mean.getOrbit(), I);
    // Set the force models
    final List<ShortPeriodTerms> shortPeriodTerms = new ArrayList<ShortPeriodTerms>();
    for (final DSSTForceModel force : forces) {
        force.registerAttitudeProvider(attitudeProvider);
        shortPeriodTerms.addAll(force.initialize(aux, false));
        force.updateShortPeriodTerms(mean);
    }
    final EquinoctialOrbit osculatingOrbit = computeOsculatingOrbit(mean, shortPeriodTerms);
    return new SpacecraftState(osculatingOrbit, mean.getAttitude(), mean.getMass(), mean.getAdditionalStates());
}

Example 9 with EquinoctialOrbit

use of org.orekit.orbits.EquinoctialOrbit in project Orekit by CS-SI.

the class ThirdBodyAttractionTest method testSunContrib.

@Test(expected = OrekitException.class)
public void testSunContrib() throws OrekitException {
    // initialization
    AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 07, 01), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
    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(), date, mu);
    double period = 2 * FastMath.PI * orbit.getA() * FastMath.sqrt(orbit.getA() / orbit.getMu());
    // set up propagator
    NumericalPropagator calc = new NumericalPropagator(new GraggBulirschStoerIntegrator(10.0, period, 0, 1.0e-5));
    calc.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getSun()));
    // set up step handler to perform checks
    calc.setMasterMode(FastMath.floor(period), new ReferenceChecker(date) {

        protected double hXRef(double t) {
            return -1.06757e-3 + 0.221415e-11 * t + 18.9421e-5 * FastMath.cos(3.9820426e-7 * t) - 7.59983e-5 * FastMath.sin(3.9820426e-7 * t);
        }

        protected double hYRef(double t) {
            return 1.43526e-3 + 7.49765e-11 * t + 6.9448e-5 * FastMath.cos(3.9820426e-7 * t) + 17.6083e-5 * FastMath.sin(3.9820426e-7 * t);
        }
    });
    AbsoluteDate finalDate = date.shiftedBy(365 * period);
    calc.setInitialState(new SpacecraftState(orbit));
    calc.propagate(finalDate);
}

Example 10 with EquinoctialOrbit

use of org.orekit.orbits.EquinoctialOrbit in project Orekit by CS-SI.

the class SolarRadiationPressureTest method testRoughOrbitalModifs.

@Test
public void testRoughOrbitalModifs() throws ParseException, OrekitException, FileNotFoundException {
    // initialization
    AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 7, 1), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
    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(), date, mu);
    final double period = orbit.getKeplerianPeriod();
    Assert.assertEquals(86164, period, 1);
    PVCoordinatesProvider sun = CelestialBodyFactory.getSun();
    // creation of the force model
    OneAxisEllipsoid earth = new OneAxisEllipsoid(6378136.46, 1.0 / 298.25765, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
    SolarRadiationPressure SRP = new SolarRadiationPressure(sun, earth.getEquatorialRadius(), new IsotropicRadiationCNES95Convention(500.0, 0.7, 0.7));
    // creation of the propagator
    double[] absTolerance = { 0.1, 1.0e-9, 1.0e-9, 1.0e-5, 1.0e-5, 1.0e-5, 0.001 };
    double[] relTolerance = { 1.0e-4, 1.0e-4, 1.0e-4, 1.0e-6, 1.0e-6, 1.0e-6, 1.0e-7 };
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(900.0, 60000, absTolerance, relTolerance);
    integrator.setInitialStepSize(3600);
    final NumericalPropagator calc = new NumericalPropagator(integrator);
    calc.addForceModel(SRP);
    // Step Handler
    calc.setMasterMode(FastMath.floor(period), new SolarStepHandler());
    AbsoluteDate finalDate = date.shiftedBy(10 * period);
    calc.setInitialState(new SpacecraftState(orbit, 1500.0));
    calc.propagate(finalDate);
    Assert.assertTrue(calc.getCalls() < 7100);
}