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

Example 51 with KeplerianOrbit

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

the class ImpulseManeuverTest method testBackward.

@Test
public void testBackward() throws OrekitException {
    final AbsoluteDate iniDate = new AbsoluteDate(2003, 5, 1, 17, 30, 0.0, TimeScalesFactory.getUTC());
    final Orbit initialOrbit = new KeplerianOrbit(7e6, 1.0e-4, FastMath.toRadians(98.5), FastMath.toRadians(87.0), FastMath.toRadians(216.1807), FastMath.toRadians(319.779), PositionAngle.MEAN, FramesFactory.getEME2000(), iniDate, Constants.EIGEN5C_EARTH_MU);
    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit, new LofOffset(initialOrbit.getFrame(), LOFType.VNC));
    DateDetector dateDetector = new DateDetector(iniDate.shiftedBy(-300));
    Vector3D deltaV = new Vector3D(12.0, 1.0, -4.0);
    final double isp = 300;
    ImpulseManeuver<DateDetector> maneuver = new ImpulseManeuver<DateDetector>(dateDetector, deltaV, isp).withMaxCheck(3600.0).withThreshold(1.0e-6);
    propagator.addEventDetector(maneuver);
    SpacecraftState finalState = propagator.propagate(initialOrbit.getDate().shiftedBy(-900));
    Assert.assertTrue(finalState.getMass() > propagator.getInitialState().getMass());
    Assert.assertTrue(finalState.getDate().compareTo(propagator.getInitialState().getDate()) < 0);
}

Example 52 with KeplerianOrbit

use of org.orekit.orbits.KeplerianOrbit 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 53 with KeplerianOrbit

use of org.orekit.orbits.KeplerianOrbit 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 54 with KeplerianOrbit

use of org.orekit.orbits.KeplerianOrbit 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);
}

Example 55 with KeplerianOrbit

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

the class SolarRadiationPressureTest method testGlobalStateJacobianIsotropicSingle.

@Test
public void testGlobalStateJacobianIsotropicSingle() 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 IsotropicRadiationSingleCoefficient(2.5, 0.7));
    propagator.addForceModel(forceModel);
    SpacecraftState state0 = new SpacecraftState(orbit);
    checkStateJacobian(propagator, state0, date.shiftedBy(3.5 * 3600.0), 1e3, tolerances[0], 2.0e-5);
}