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

Example 21 with DateComponents

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

the class ConstantThrustManeuverTest method testJacobianVs80Implementation.

@Test
public void testJacobianVs80Implementation() 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 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 AttitudeProvider law = new LofOffset(orbit.getFrame(), LOFType.VVLH);
    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);
    // before maneuver (Jacobian wrt. state is zero)
    checkStateJacobianVs80Implementation(initialState, maneuver, law, 1.0e-50, false);
    // in maneuver
    SpacecraftState startState = initialState.shiftedBy(fireDate.durationFrom(initDate));
    maneuver.getEventsDetectors().findFirst().get().eventOccurred(startState, true);
    SpacecraftState midState = startState.shiftedBy(duration / 2.0);
    checkStateJacobianVs80Implementation(midState, maneuver, law, 1.0e-20, false);
}

Example 22 with DateComponents

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

the class SmallManeuverAnalyticalModelTest method testLowEarthOrbit1.

@Test
public void testLowEarthOrbit1() throws OrekitException {
    Orbit leo = new CircularOrbit(7200000.0, -1.0e-5, 2.0e-4, FastMath.toRadians(98.0), FastMath.toRadians(123.456), 0.0, PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
    double mass = 5600.0;
    AbsoluteDate t0 = leo.getDate().shiftedBy(1000.0);
    Vector3D dV = new Vector3D(-0.01, 0.02, 0.03);
    double f = 20.0;
    double isp = 315.0;
    BoundedPropagator withoutManeuver = getEphemeris(leo, mass, t0, Vector3D.ZERO, f, isp);
    BoundedPropagator withManeuver = getEphemeris(leo, mass, t0, dV, f, isp);
    SmallManeuverAnalyticalModel model = new SmallManeuverAnalyticalModel(withoutManeuver.propagate(t0), dV, isp);
    Assert.assertEquals(t0, model.getDate());
    for (AbsoluteDate t = withoutManeuver.getMinDate(); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t.shiftedBy(60.0)) {
        PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, leo.getFrame());
        PVCoordinates pvWith = withManeuver.getPVCoordinates(t, leo.getFrame());
        PVCoordinates pvModel = model.apply(withoutManeuver.propagate(t)).getPVCoordinates(leo.getFrame());
        double nominalDeltaP = new PVCoordinates(pvWith, pvWithout).getPosition().getNorm();
        double modelError = new PVCoordinates(pvWith, pvModel).getPosition().getNorm();
        if (t.compareTo(t0) < 0) {
            // before maneuver, all positions should be equal
            Assert.assertEquals(0, nominalDeltaP, 1.0e-10);
            Assert.assertEquals(0, modelError, 1.0e-10);
        } else {
            // despite nominal deltaP exceeds 1 kilometer after less than 3 orbits
            if (t.durationFrom(t0) > 0.1 * leo.getKeplerianPeriod()) {
                Assert.assertTrue(modelError < 0.009 * nominalDeltaP);
            }
            Assert.assertTrue(modelError < 0.8);
        }
    }
}

Example 23 with DateComponents

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

the class SmallManeuverAnalyticalModelTest method testEccentricOrbit.

@Test
public void testEccentricOrbit() throws OrekitException {
    Orbit heo = new KeplerianOrbit(90000000.0, 0.92, FastMath.toRadians(98.0), FastMath.toRadians(12.3456), FastMath.toRadians(123.456), FastMath.toRadians(1.23456), PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
    double mass = 5600.0;
    AbsoluteDate t0 = heo.getDate().shiftedBy(1000.0);
    Vector3D dV = new Vector3D(-0.01, 0.02, 0.03);
    double f = 20.0;
    double isp = 315.0;
    BoundedPropagator withoutManeuver = getEphemeris(heo, mass, t0, Vector3D.ZERO, f, isp);
    BoundedPropagator withManeuver = getEphemeris(heo, mass, t0, dV, f, isp);
    SmallManeuverAnalyticalModel model = new SmallManeuverAnalyticalModel(withoutManeuver.propagate(t0), dV, isp);
    Assert.assertEquals(t0, model.getDate());
    for (AbsoluteDate t = withoutManeuver.getMinDate(); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t.shiftedBy(600.0)) {
        PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, heo.getFrame());
        PVCoordinates pvWith = withManeuver.getPVCoordinates(t, heo.getFrame());
        PVCoordinates pvModel = model.apply(withoutManeuver.propagate(t)).getPVCoordinates(heo.getFrame());
        double nominalDeltaP = new PVCoordinates(pvWith, pvWithout).getPosition().getNorm();
        double modelError = new PVCoordinates(pvWith, pvModel).getPosition().getNorm();
        if (t.compareTo(t0) < 0) {
            // before maneuver, all positions should be equal
            Assert.assertEquals(0, nominalDeltaP, 1.0e-10);
            Assert.assertEquals(0, modelError, 1.0e-10);
        } else {
            // despite nominal deltaP exceeds 300 kilometers at perigee, after 3 orbits
            if (t.durationFrom(t0) > 0.01 * heo.getKeplerianPeriod()) {
                Assert.assertTrue(modelError < 0.005 * nominalDeltaP);
            }
            Assert.assertTrue(modelError < 1700);
        }
    }
}

Example 24 with DateComponents

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

the class SolarRadiationPressureTest method testGlobalStateJacobianIsotropicClassical.

@Test
public void testGlobalStateJacobianIsotropicClassical() 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);
    SolarRadiationPressure forceModel = new SolarRadiationPressure(CelestialBodyFactory.getSun(), Constants.WGS84_EARTH_EQUATORIAL_RADIUS, new IsotropicRadiationClassicalConvention(2.5, 0.7, 0.2));
    propagator.addForceModel(forceModel);
    SpacecraftState state0 = new SpacecraftState(orbit);
    checkStateJacobian(propagator, state0, date.shiftedBy(3.5 * 3600.0), 1e6, tolerances[0], 2.0e-5);
}

Example 25 with DateComponents

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

the class SolarRadiationPressureTest method testLocalJacobianIsotropicClassicalVs80Implementation.

@Test
public void testLocalJacobianIsotropicClassicalVs80Implementation() 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);
    final SolarRadiationPressure forceModel = new SolarRadiationPressure(CelestialBodyFactory.getSun(), Constants.WGS84_EARTH_EQUATORIAL_RADIUS, new IsotropicRadiationClassicalConvention(2.5, 0.7, 0.2));
    checkStateJacobianVs80Implementation(new SpacecraftState(orbit), forceModel, new LofOffset(orbit.getFrame(), LOFType.VVLH), 1.0e-15, false);
}