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

Example 11 with TimeComponents

use of org.orekit.time.TimeComponents 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 12 with TimeComponents

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

the class JPLEphemeridesLoader method extractDate.

/**
 * Extract a date from a record.
 * @param record record to parse
 * @param offset offset of the double within the record
 * @return extracted date
 */
private AbsoluteDate extractDate(final byte[] record, final int offset) {
    final double t = extractDouble(record, offset);
    int jDay = (int) FastMath.floor(t);
    double seconds = (t + 0.5 - jDay) * Constants.JULIAN_DAY;
    if (seconds >= Constants.JULIAN_DAY) {
        ++jDay;
        seconds -= Constants.JULIAN_DAY;
    }
    return new AbsoluteDate(new DateComponents(DateComponents.JULIAN_EPOCH, jDay), new TimeComponents(seconds), timeScale);
}

Example 13 with TimeComponents

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

the class ThirdBodyAttractionTest method testGlobalStateJacobian.

@Test
public void testGlobalStateJacobian() throws OrekitException {
    // initialization
    AbsoluteDate date = new AbsoluteDate(new DateComponents(2003, 03, 01), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
    double i = FastMath.toRadians(98.7);
    double omega = FastMath.toRadians(93.0);
    double OMEGA = FastMath.toRadians(15.0 * 22.5);
    Orbit orbit = new KeplerianOrbit(7201009.7124401, 1e-3, i, omega, OMEGA, 0, PositionAngle.MEAN, FramesFactory.getEME2000(), date, Constants.EIGEN5C_EARTH_MU);
    OrbitType integrationType = OrbitType.CARTESIAN;
    double[][] tolerances = NumericalPropagator.tolerances(0.01, orbit, integrationType);
    NumericalPropagator propagator = new NumericalPropagator(new DormandPrince853Integrator(1.0e-3, 120, tolerances[0], tolerances[1]));
    propagator.setOrbitType(integrationType);
    final CelestialBody moon = CelestialBodyFactory.getMoon();
    final ThirdBodyAttraction forceModel = new ThirdBodyAttraction(moon);
    propagator.addForceModel(forceModel);
    SpacecraftState state0 = new SpacecraftState(orbit);
    checkStateJacobian(propagator, state0, date.shiftedBy(3.5 * 3600.0), 1e4, tolerances[0], 2.0e-9);
}

Example 14 with TimeComponents

use of org.orekit.time.TimeComponents 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 15 with TimeComponents

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

the class ConstantThrustManeuverTest method testForwardAndBackward.

@Test
public void testForwardAndBackward() throws OrekitException {
    final double isp = 318;
    final double mass = 2500;
    final double a = 24396159;
    final double e = 0.72831215;
    final double i = FastMath.toRadians(7);
    final double omega = FastMath.toRadians(180);
    final double OMEGA = FastMath.toRadians(261);
    final double lv = 0;
    final double duration = 3653.99;
    final double f = 420;
    final double delta = FastMath.toRadians(-7.4978);
    final double alpha = FastMath.toRadians(351);
    final AttitudeProvider law = new InertialProvider(new Rotation(new Vector3D(alpha, delta), Vector3D.PLUS_I));
    final AbsoluteDate initDate = new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC());
    final Orbit orbit = new KeplerianOrbit(a, e, i, omega, OMEGA, lv, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, mu);
    final SpacecraftState initialState = new SpacecraftState(orbit, law.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass);
    final AbsoluteDate fireDate = new AbsoluteDate(new DateComponents(2004, 01, 02), new TimeComponents(04, 15, 34.080), TimeScalesFactory.getUTC());
    final ConstantThrustManeuver maneuver = new ConstantThrustManeuver(fireDate, duration, f, isp, Vector3D.PLUS_I);
    Assert.assertEquals(f, maneuver.getThrust(), 1.0e-10);
    Assert.assertEquals(isp, maneuver.getISP(), 1.0e-10);
    double[][] tol = NumericalPropagator.tolerances(1.0, orbit, OrbitType.KEPLERIAN);
    AdaptiveStepsizeIntegrator integrator1 = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator1.setInitialStepSize(60);
    final NumericalPropagator propagator1 = new NumericalPropagator(integrator1);
    propagator1.setInitialState(initialState);
    propagator1.setAttitudeProvider(law);
    propagator1.addForceModel(maneuver);
    final SpacecraftState finalState = propagator1.propagate(fireDate.shiftedBy(3800));
    AdaptiveStepsizeIntegrator integrator2 = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator2.setInitialStepSize(60);
    final NumericalPropagator propagator2 = new NumericalPropagator(integrator2);
    propagator2.setInitialState(finalState);
    propagator2.setAttitudeProvider(law);
    propagator2.addForceModel(maneuver);
    final SpacecraftState recoveredState = propagator2.propagate(orbit.getDate());
    final Vector3D refPosition = initialState.getPVCoordinates().getPosition();
    final Vector3D recoveredPosition = recoveredState.getPVCoordinates().getPosition();
    Assert.assertEquals(0.0, Vector3D.distance(refPosition, recoveredPosition), 30.0);
    Assert.assertEquals(initialState.getMass(), recoveredState.getMass(), 1.5e-10);
}