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

Example 31 with Orbit

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

the class NumericalPropagatorTest method testPropagationTypesElliptical.

@Test
public void testPropagationTypesElliptical() throws OrekitException, ParseException, IOException {
    // setup
    AbsoluteDate initDate = new AbsoluteDate();
    SpacecraftState initialState;
    final Vector3D position = new Vector3D(7.0e6, 1.0e6, 4.0e6);
    final Vector3D velocity = new Vector3D(-500.0, 8000.0, 1000.0);
    initDate = AbsoluteDate.J2000_EPOCH;
    final Orbit orbit = new EquinoctialOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), initDate, mu);
    initialState = new SpacecraftState(orbit);
    OrbitType type = OrbitType.EQUINOCTIAL;
    double[][] tolerance = NumericalPropagator.tolerances(0.001, orbit, type);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 200, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(60);
    propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(type);
    propagator.setInitialState(initialState);
    ForceModel gravityField = new HolmesFeatherstoneAttractionModel(FramesFactory.getITRF(IERSConventions.IERS_2010, true), GravityFieldFactory.getNormalizedProvider(5, 5));
    propagator.addForceModel(gravityField);
    // Propagation of the initial at t + dt
    final PVCoordinates pv = initialState.getPVCoordinates();
    final double dP = 0.001;
    final double dV = initialState.getMu() * dP / (pv.getPosition().getNormSq() * pv.getVelocity().getNorm());
    final PVCoordinates pvcM = propagateInType(initialState, dP, OrbitType.CARTESIAN, PositionAngle.MEAN);
    final PVCoordinates pviM = propagateInType(initialState, dP, OrbitType.CIRCULAR, PositionAngle.MEAN);
    final PVCoordinates pveM = propagateInType(initialState, dP, OrbitType.EQUINOCTIAL, PositionAngle.MEAN);
    final PVCoordinates pvkM = propagateInType(initialState, dP, OrbitType.KEPLERIAN, PositionAngle.MEAN);
    final PVCoordinates pvcE = propagateInType(initialState, dP, OrbitType.CARTESIAN, PositionAngle.ECCENTRIC);
    final PVCoordinates pviE = propagateInType(initialState, dP, OrbitType.CIRCULAR, PositionAngle.ECCENTRIC);
    final PVCoordinates pveE = propagateInType(initialState, dP, OrbitType.EQUINOCTIAL, PositionAngle.ECCENTRIC);
    final PVCoordinates pvkE = propagateInType(initialState, dP, OrbitType.KEPLERIAN, PositionAngle.ECCENTRIC);
    final PVCoordinates pvcT = propagateInType(initialState, dP, OrbitType.CARTESIAN, PositionAngle.TRUE);
    final PVCoordinates pviT = propagateInType(initialState, dP, OrbitType.CIRCULAR, PositionAngle.TRUE);
    final PVCoordinates pveT = propagateInType(initialState, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE);
    final PVCoordinates pvkT = propagateInType(initialState, dP, OrbitType.KEPLERIAN, PositionAngle.TRUE);
    Assert.assertEquals(0, pvcM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 3.0);
    Assert.assertEquals(0, pvcM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 2.0);
    Assert.assertEquals(0, pviM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.6);
    Assert.assertEquals(0, pviM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.4);
    Assert.assertEquals(0, pvkM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.5);
    Assert.assertEquals(0, pvkM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.3);
    Assert.assertEquals(0, pveM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.2);
    Assert.assertEquals(0, pveM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.2);
    Assert.assertEquals(0, pvcE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 3.0);
    Assert.assertEquals(0, pvcE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 2.0);
    Assert.assertEquals(0, pviE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.03);
    Assert.assertEquals(0, pviE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.04);
    Assert.assertEquals(0, pvkE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.4);
    Assert.assertEquals(0, pvkE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.3);
    Assert.assertEquals(0, pveE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.2);
    Assert.assertEquals(0, pveE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.07);
    Assert.assertEquals(0, pvcT.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 3.0);
    Assert.assertEquals(0, pvcT.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 2.0);
    Assert.assertEquals(0, pviT.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.3);
    Assert.assertEquals(0, pviT.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.2);
    Assert.assertEquals(0, pvkT.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.4);
    Assert.assertEquals(0, pvkT.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.2);
}

Example 32 with Orbit

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

the class PartialDerivativesTest method testWrongParametersDimension.

@Test
public void testWrongParametersDimension() throws OrekitException {
    Orbit initialOrbit = new KeplerianOrbit(8000000.0, 0.01, 0.1, 0.7, 0, 1.2, PositionAngle.TRUE, FramesFactory.getEME2000(), AbsoluteDate.J2000_EPOCH, Constants.EIGEN5C_EARTH_MU);
    double dP = 0.001;
    ForceModel sunAttraction = new ThirdBodyAttraction(CelestialBodyFactory.getSun());
    sunAttraction.getParameterDriver("Sun attraction coefficient").setSelected(true);
    ForceModel moonAttraction = new ThirdBodyAttraction(CelestialBodyFactory.getMoon());
    NumericalPropagator propagator = setUpPropagator(initialOrbit, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE, sunAttraction, moonAttraction);
    PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
    try {
        partials.setInitialJacobians(new SpacecraftState(initialOrbit), new double[6][6], new double[6][3]);
        partials.computeDerivatives(new SpacecraftState(initialOrbit), new double[6]);
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(OrekitMessages.INITIAL_MATRIX_AND_PARAMETERS_NUMBER_MISMATCH, oe.getSpecifier());
    }
}

Example 33 with Orbit

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

the class PartialDerivativesTest method testTooLargeDimension.

@Test(expected = OrekitException.class)
public void testTooLargeDimension() throws OrekitException {
    Orbit initialOrbit = new KeplerianOrbit(8000000.0, 0.01, 0.1, 0.7, 0, 1.2, PositionAngle.TRUE, FramesFactory.getEME2000(), AbsoluteDate.J2000_EPOCH, Constants.EIGEN5C_EARTH_MU);
    double dP = 0.001;
    NumericalPropagator propagator = setUpPropagator(initialOrbit, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE);
    PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
    partials.setInitialJacobians(new SpacecraftState(initialOrbit), new double[8][6], new double[6][2]);
}

Example 34 with Orbit

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

the class PartialDerivativesTest method testNotInitialized.

@Test(expected = OrekitException.class)
public void testNotInitialized() throws OrekitException {
    Orbit initialOrbit = new KeplerianOrbit(8000000.0, 0.01, 0.1, 0.7, 0, 1.2, PositionAngle.TRUE, FramesFactory.getEME2000(), AbsoluteDate.J2000_EPOCH, Constants.EIGEN5C_EARTH_MU);
    double dP = 0.001;
    NumericalPropagator propagator = setUpPropagator(initialOrbit, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE);
    new PartialDerivativesEquations("partials", propagator).getMapper();
}

Example 35 with Orbit

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

the class PartialDerivativesTest method testTooSmallDimension.

@Test(expected = OrekitException.class)
public void testTooSmallDimension() throws OrekitException {
    Orbit initialOrbit = new KeplerianOrbit(8000000.0, 0.01, 0.1, 0.7, 0, 1.2, PositionAngle.TRUE, FramesFactory.getEME2000(), AbsoluteDate.J2000_EPOCH, Constants.EIGEN5C_EARTH_MU);
    double dP = 0.001;
    NumericalPropagator propagator = setUpPropagator(initialOrbit, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE);
    PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
    partials.setInitialJacobians(new SpacecraftState(initialOrbit), new double[5][6], new double[6][2]);
}