Examples with SpacecraftState org.orekit.propagation.SpacecraftState used on opensource projects

Example 41 with SpacecraftState

use of org.orekit.propagation.SpacecraftState in project Orekit by CS-SI.

the class KalmanOrbitDeterminationTest method createOrbit.

/**
 * Create an orbit from input parameters
 * @param parser input file parser
 * @param mu     central attraction coefficient
 * @throws NoSuchElementException if input parameters are missing
 * @throws OrekitException if inertial frame cannot be created
 */
private Orbit createOrbit(final KeyValueFileParser<ParameterKey> parser, final double mu) throws NoSuchElementException, OrekitException {
    final Frame frame;
    if (!parser.containsKey(ParameterKey.INERTIAL_FRAME)) {
        frame = FramesFactory.getEME2000();
    } else {
        frame = parser.getInertialFrame(ParameterKey.INERTIAL_FRAME);
    }
    // Orbit definition
    PositionAngle angleType = PositionAngle.MEAN;
    if (parser.containsKey(ParameterKey.ORBIT_ANGLE_TYPE)) {
        angleType = PositionAngle.valueOf(parser.getString(ParameterKey.ORBIT_ANGLE_TYPE).toUpperCase());
    }
    if (parser.containsKey(ParameterKey.ORBIT_KEPLERIAN_A)) {
        return new KeplerianOrbit(parser.getDouble(ParameterKey.ORBIT_KEPLERIAN_A), parser.getDouble(ParameterKey.ORBIT_KEPLERIAN_E), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_I), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_PA), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_RAAN), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_ANOMALY), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_EQUINOCTIAL_A)) {
        return new EquinoctialOrbit(parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_A), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_EX), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_EY), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_HX), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_HY), parser.getAngle(ParameterKey.ORBIT_EQUINOCTIAL_LAMBDA), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_CIRCULAR_A)) {
        return new CircularOrbit(parser.getDouble(ParameterKey.ORBIT_CIRCULAR_A), parser.getDouble(ParameterKey.ORBIT_CIRCULAR_EX), parser.getDouble(ParameterKey.ORBIT_CIRCULAR_EY), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_I), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_RAAN), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_ALPHA), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_TLE_LINE_1)) {
        final String line1 = parser.getString(ParameterKey.ORBIT_TLE_LINE_1);
        final String line2 = parser.getString(ParameterKey.ORBIT_TLE_LINE_2);
        final TLE tle = new TLE(line1, line2);
        TLEPropagator propagator = TLEPropagator.selectExtrapolator(tle);
        // propagator.setEphemerisMode();
        AbsoluteDate initDate = tle.getDate();
        SpacecraftState initialState = propagator.getInitialState();
        // Transformation from TEME to frame.
        Transform t = FramesFactory.getTEME().getTransformTo(FramesFactory.getEME2000(), initDate.getDate());
        return new CartesianOrbit(t.transformPVCoordinates(initialState.getPVCoordinates()), frame, initDate, mu);
    } else {
        final double[] pos = { parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PX), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PY), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PZ) };
        final double[] vel = { parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VX), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VY), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VZ) };
        return new CartesianOrbit(new PVCoordinates(new Vector3D(pos), new Vector3D(vel)), frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    }
}

Example 42 with SpacecraftState

use of org.orekit.propagation.SpacecraftState in project Orekit by CS-SI.

the class AbstractForceModelTest method integrateShiftedState.

private double[] integrateShiftedState(final NumericalPropagator propagator, final SpacecraftState state0, final AbsoluteDate targetDate, final int index, final double h) throws OrekitException {
    OrbitType orbitType = propagator.getOrbitType();
    PositionAngle angleType = propagator.getPositionAngleType();
    double[] a = new double[6];
    double[] aDot = new double[6];
    orbitType.mapOrbitToArray(state0.getOrbit(), angleType, a, aDot);
    a[index] += h;
    SpacecraftState shiftedState = new SpacecraftState(orbitType.mapArrayToOrbit(a, aDot, angleType, state0.getDate(), state0.getMu(), state0.getFrame()), state0.getAttitude(), state0.getMass());
    propagator.setInitialState(shiftedState);
    SpacecraftState integratedState = propagator.propagate(targetDate);
    orbitType.mapOrbitToArray(integratedState.getOrbit(), angleType, a, null);
    return a;
}

Example 43 with SpacecraftState

use of org.orekit.propagation.SpacecraftState in project Orekit by CS-SI.

the class BoxAndSolarArraySpacecraftTest method testWrongParameterLift.

@Test
public void testWrongParameterLift() throws OrekitException {
    SpacecraftState state = propagator.getInitialState();
    CelestialBody sun = CelestialBodyFactory.getSun();
    BoxAndSolarArraySpacecraft s = new BoxAndSolarArraySpacecraft(0, 0, 0, sun, 20.0, Vector3D.PLUS_J, 2.0, 0.4, 1.0, 0.0);
    try {
        s.dragAcceleration(state.getDate(), state.getFrame(), state.getPVCoordinates().getPosition(), state.getAttitude().getRotation(), state.getMass(), 1.0e-6, Vector3D.PLUS_I, getDragParameters(s), "wrong");
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(OrekitMessages.UNSUPPORTED_PARAMETER_NAME, oe.getSpecifier());
        Assert.assertEquals(3, oe.getParts().length);
        Assert.assertEquals("wrong", (String) oe.getParts()[0]);
        Assert.assertEquals(DragSensitive.DRAG_COEFFICIENT, (String) oe.getParts()[1]);
        Assert.assertEquals(DragSensitive.LIFT_RATIO, (String) oe.getParts()[2]);
    }
}

Example 44 with SpacecraftState

use of org.orekit.propagation.SpacecraftState in project Orekit by CS-SI.

the class BoxAndSolarArraySpacecraftTest method testBackwardIllumination.

@Test
public void testBackwardIllumination() throws OrekitException {
    SpacecraftState state = propagator.getInitialState();
    CelestialBody sun = CelestialBodyFactory.getSun();
    BoxAndSolarArraySpacecraft s = new BoxAndSolarArraySpacecraft(0, 0, 0, sun, 20.0, Vector3D.PLUS_J, 0.0, 1.0, 0.0);
    Vector3D n = s.getNormal(state.getDate(), state.getFrame(), state.getPVCoordinates().getPosition(), state.getAttitude().getRotation());
    FieldVector3D<DerivativeStructure> aPlus = s.radiationPressureAcceleration(state.getDate(), state.getFrame(), state.getPVCoordinates().getPosition(), state.getAttitude().getRotation(), state.getMass(), n, getRadiationParameters(s), RadiationSensitive.ABSORPTION_COEFFICIENT);
    FieldVector3D<DerivativeStructure> aMinus = s.radiationPressureAcceleration(state.getDate(), state.getFrame(), state.getPVCoordinates().getPosition(), state.getAttitude().getRotation(), state.getMass(), n.negate(), getRadiationParameters(s), RadiationSensitive.ABSORPTION_COEFFICIENT);
    Assert.assertEquals(0.0, aPlus.add(aMinus).getNorm().getReal(), Double.MIN_VALUE);
}

Example 45 with SpacecraftState

use of org.orekit.propagation.SpacecraftState in project Orekit by CS-SI.

the class BoxAndSolarArraySpacecraftTest method testNormalOptimalRotationField.

@Test
public void testNormalOptimalRotationField() throws OrekitException {
    AbsoluteDate initialDate = propagator.getInitialState().getDate();
    CelestialBody sun = CelestialBodyFactory.getSun();
    BoxAndSolarArraySpacecraft s = new BoxAndSolarArraySpacecraft(0, 0, 0, sun, 20.0, Vector3D.PLUS_J, 0.0, 1.0, 0.0);
    Field<Decimal64> field = Decimal64Field.getInstance();
    for (double dt = 0; dt < 4000; dt += 60) {
        AbsoluteDate date = initialDate.shiftedBy(dt);
        SpacecraftState state = propagator.propagate(date);
        FieldVector3D<Decimal64> normal = s.getNormal(new FieldAbsoluteDate<>(field, state.getDate()), state.getFrame(), new FieldVector3D<>(field, state.getPVCoordinates().getPosition()), new FieldRotation<>(field, state.getAttitude().getRotation()));
        Assert.assertEquals(0, FieldVector3D.dotProduct(normal, Vector3D.PLUS_J).getReal(), 1.0e-16);
    }
}