Examples with NumericalPropagator org.orekit.propagation.numerical.NumericalPropagator used on opensource projects

Example 66 with NumericalPropagator

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

the class PolynomialParametricAccelerationTest method doTestEquivalentManeuver.

private void doTestEquivalentManeuver(final double mass, final AttitudeProvider maneuverLaw, final ConstantThrustManeuver maneuver, final AttitudeProvider accelerationLaw, final PolynomialParametricAcceleration parametricAcceleration, final double positionTolerance) throws OrekitException {
    SpacecraftState initialState = new SpacecraftState(initialOrbit, maneuverLaw.getAttitude(initialOrbit, initialOrbit.getDate(), initialOrbit.getFrame()), mass);
    double[][] tolerance = NumericalPropagator.tolerances(10, initialOrbit, initialOrbit.getType());
    // propagator 0 uses a maneuver that is so efficient it does not consume any fuel
    // (hence mass remains constant)
    AdaptiveStepsizeIntegrator integrator0 = new DormandPrince853Integrator(0.001, 100, tolerance[0], tolerance[1]);
    integrator0.setInitialStepSize(60);
    final NumericalPropagator propagator0 = new NumericalPropagator(integrator0);
    propagator0.setInitialState(initialState);
    propagator0.setAttitudeProvider(maneuverLaw);
    propagator0.addForceModel(maneuver);
    // propagator 1 uses a constant acceleration
    AdaptiveStepsizeIntegrator integrator1 = new DormandPrince853Integrator(0.001, 100, tolerance[0], tolerance[1]);
    integrator1.setInitialStepSize(60);
    final NumericalPropagator propagator1 = new NumericalPropagator(integrator1);
    propagator1.setInitialState(initialState);
    propagator1.setAttitudeProvider(accelerationLaw);
    propagator1.addForceModel(parametricAcceleration);
    MultiSatStepHandler handler = (interpolators, isLast) -> {
        Vector3D p0 = interpolators.get(0).getCurrentState().getPVCoordinates().getPosition();
        Vector3D p1 = interpolators.get(1).getCurrentState().getPVCoordinates().getPosition();
        Assert.assertEquals(0.0, Vector3D.distance(p0, p1), positionTolerance);
    };
    PropagatorsParallelizer parallelizer = new PropagatorsParallelizer(Arrays.asList(propagator0, propagator1), handler);
    parallelizer.propagate(initialOrbit.getDate(), initialOrbit.getDate().shiftedBy(1000.0));
}

Example 67 with NumericalPropagator

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

the class DragForceTest method testIssue229.

@Test
public void testIssue229() throws OrekitException {
    AbsoluteDate initialDate = new AbsoluteDate(2004, 1, 1, 0, 0, 0., TimeScalesFactory.getUTC());
    Frame frame = FramesFactory.getEME2000();
    double rpe = 160.e3 + Constants.WGS84_EARTH_EQUATORIAL_RADIUS;
    double rap = 2000.e3 + Constants.WGS84_EARTH_EQUATORIAL_RADIUS;
    double inc = FastMath.toRadians(0.);
    double aop = FastMath.toRadians(0.);
    double raan = FastMath.toRadians(0.);
    double mean = FastMath.toRadians(180.);
    double mass = 100.;
    KeplerianOrbit orbit = new KeplerianOrbit(0.5 * (rpe + rap), (rap - rpe) / (rpe + rap), inc, aop, raan, mean, PositionAngle.MEAN, frame, initialDate, Constants.EIGEN5C_EARTH_MU);
    IsotropicDrag shape = new IsotropicDrag(10., 2.2);
    Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
    BodyShape earthShape = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, itrf);
    Atmosphere atmosphere = new SimpleExponentialAtmosphere(earthShape, 2.6e-10, 200000, 26000);
    double[][] tolerance = NumericalPropagator.tolerances(0.1, orbit, OrbitType.CARTESIAN);
    AbstractIntegrator integrator = new DormandPrince853Integrator(1.0e-3, 300, tolerance[0], tolerance[1]);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    propagator.setMu(orbit.getMu());
    propagator.addForceModel(new DragForce(atmosphere, shape));
    PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
    propagator.setInitialState(partials.setInitialJacobians(new SpacecraftState(orbit, mass)));
    SpacecraftState state = propagator.propagate(new AbsoluteDate(2004, 1, 1, 1, 30, 0., TimeScalesFactory.getUTC()));
    double delta = 0.1;
    Orbit shifted = new CartesianOrbit(new TimeStampedPVCoordinates(orbit.getDate(), orbit.getPVCoordinates().getPosition().add(new Vector3D(delta, 0, 0)), orbit.getPVCoordinates().getVelocity()), orbit.getFrame(), orbit.getMu());
    propagator.setInitialState(partials.setInitialJacobians(new SpacecraftState(shifted, mass)));
    SpacecraftState newState = propagator.propagate(new AbsoluteDate(2004, 1, 1, 1, 30, 0., TimeScalesFactory.getUTC()));
    double[] dPVdX = new double[] { (newState.getPVCoordinates().getPosition().getX() - state.getPVCoordinates().getPosition().getX()) / delta, (newState.getPVCoordinates().getPosition().getY() - state.getPVCoordinates().getPosition().getY()) / delta, (newState.getPVCoordinates().getPosition().getZ() - state.getPVCoordinates().getPosition().getZ()) / delta, (newState.getPVCoordinates().getVelocity().getX() - state.getPVCoordinates().getVelocity().getX()) / delta, (newState.getPVCoordinates().getVelocity().getY() - state.getPVCoordinates().getVelocity().getY()) / delta, (newState.getPVCoordinates().getVelocity().getZ() - state.getPVCoordinates().getVelocity().getZ()) / delta };
    double[][] dYdY0 = new double[6][6];
    partials.getMapper().getStateJacobian(state, dYdY0);
    for (int i = 0; i < 6; ++i) {
        Assert.assertEquals(dPVdX[i], dYdY0[i][0], 6.2e-6 * FastMath.abs(dPVdX[i]));
    }
}

Example 68 with NumericalPropagator

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

the class MarshallSolarActivityFutureEstimationTest method getNumericalPropagator.

/**
 * Configure a numerical propagator.
 *
 * @param sun   Sun.
 * @param earth Earth.
 * @param ic    initial condition.
 * @return a propagator.
 * @throws OrekitException on error.
 */
private NumericalPropagator getNumericalPropagator(CelestialBody sun, OneAxisEllipsoid earth, SpacecraftState ic) throws OrekitException {
    // some non-integer step size to induce truncation error in flux interpolation
    final ODEIntegrator integrator = new ClassicalRungeKuttaIntegrator(120 + 0.1);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    DTM2000InputParameters flux = getFlux();
    final Atmosphere atmosphere = new DTM2000(flux, sun, earth);
    final IsotropicDrag satellite = new IsotropicDrag(1, 3.2);
    propagator.addForceModel(new DragForce(atmosphere, satellite));
    propagator.setInitialState(ic);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    return propagator;
}

Example 69 with NumericalPropagator

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

the class HolmesFeatherstoneAttractionModelTest method RealFieldExpectErrorTest.

/**
 *Same test as the previous one but not adding the ForceModel to the NumericalPropagator
 *    it is a test to validate the previous test.
 *    (to test if the ForceModel it's actually
 *    doing something in the Propagator and the FieldPropagator)
 */
@Test
public void RealFieldExpectErrorTest() throws OrekitException {
    DSFactory factory = new DSFactory(6, 0);
    DerivativeStructure a_0 = factory.variable(0, 7201009.7124401);
    DerivativeStructure e_0 = factory.variable(1, 1e-3);
    DerivativeStructure i_0 = factory.variable(2, 98.7 * FastMath.PI / 180);
    DerivativeStructure R_0 = factory.variable(3, 15.0 * 22.5 * FastMath.PI / 180);
    DerivativeStructure O_0 = factory.variable(4, 93.0 * FastMath.PI / 180);
    DerivativeStructure n_0 = factory.variable(5, 0.1);
    Field<DerivativeStructure> field = a_0.getField();
    DerivativeStructure zero = field.getZero();
    FieldAbsoluteDate<DerivativeStructure> J2000 = new FieldAbsoluteDate<>(field);
    Frame EME = FramesFactory.getEME2000();
    FieldKeplerianOrbit<DerivativeStructure> FKO = new FieldKeplerianOrbit<>(a_0, e_0, i_0, R_0, O_0, n_0, PositionAngle.MEAN, EME, J2000, Constants.EIGEN5C_EARTH_MU);
    FieldSpacecraftState<DerivativeStructure> initialState = new FieldSpacecraftState<>(FKO);
    SpacecraftState iSR = initialState.toSpacecraftState();
    OrbitType type = OrbitType.EQUINOCTIAL;
    double[][] tolerance = NumericalPropagator.tolerances(10.0, FKO.toOrbit(), type);
    AdaptiveStepsizeFieldIntegrator<DerivativeStructure> integrator = new DormandPrince853FieldIntegrator<>(field, 0.001, 200, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(zero.add(60));
    AdaptiveStepsizeIntegrator RIntegrator = new DormandPrince853Integrator(0.001, 200, tolerance[0], tolerance[1]);
    RIntegrator.setInitialStepSize(60);
    FieldNumericalPropagator<DerivativeStructure> FNP = new FieldNumericalPropagator<>(field, integrator);
    FNP.setOrbitType(type);
    FNP.setInitialState(initialState);
    NumericalPropagator NP = new NumericalPropagator(RIntegrator);
    NP.setOrbitType(type);
    NP.setInitialState(iSR);
    double[][] c = new double[3][1];
    c[0][0] = 0.0;
    c[2][0] = normalizedC20;
    double[][] s = new double[3][1];
    NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(6378136.460, mu, TideSystem.UNKNOWN, c, s);
    HolmesFeatherstoneAttractionModel forceModel = new HolmesFeatherstoneAttractionModel(itrf, provider);
    // FNP.addForceModel(forceModel);
    NP.addForceModel(forceModel);
    FieldAbsoluteDate<DerivativeStructure> target = J2000.shiftedBy(100.);
    FieldSpacecraftState<DerivativeStructure> finalState_DS = FNP.propagate(target);
    SpacecraftState finalState_R = NP.propagate(target.toAbsoluteDate());
    FieldPVCoordinates<DerivativeStructure> finPVC_DS = finalState_DS.getPVCoordinates();
    PVCoordinates finPVC_R = finalState_R.getPVCoordinates();
    Assert.assertFalse(FastMath.abs(finPVC_DS.toPVCoordinates().getPosition().getX() - finPVC_R.getPosition().getX()) < FastMath.abs(finPVC_R.getPosition().getX()) * 1e-11);
    Assert.assertFalse(FastMath.abs(finPVC_DS.toPVCoordinates().getPosition().getY() - finPVC_R.getPosition().getY()) < FastMath.abs(finPVC_R.getPosition().getY()) * 1e-11);
    Assert.assertFalse(FastMath.abs(finPVC_DS.toPVCoordinates().getPosition().getZ() - finPVC_R.getPosition().getZ()) < FastMath.abs(finPVC_R.getPosition().getZ()) * 1e-11);
}

Example 70 with NumericalPropagator

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

the class HolmesFeatherstoneAttractionModelTest method testZonalWithCunninghamReference.

// test the difference with the Cunningham model
@Test
@Deprecated
public void testZonalWithCunninghamReference() throws OrekitException {
    // initialization
    AbsoluteDate date = new AbsoluteDate(new DateComponents(2000, 07, 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, mu);
    propagator = new NumericalPropagator(new ClassicalRungeKuttaIntegrator(1000));
    propagator.addForceModel(new HolmesFeatherstoneAttractionModel(itrf, GravityFieldFactory.getNormalizedProvider(ae, mu, TideSystem.UNKNOWN, new double[][] { { 0.0 }, { 0.0 }, { normalizedC20 }, { normalizedC30 }, { normalizedC40 }, { normalizedC50 }, { normalizedC60 } }, new double[][] { { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 } })));
    propagator.setInitialState(new SpacecraftState(orbit));
    SpacecraftState hfOrb = propagator.propagate(date.shiftedBy(Constants.JULIAN_DAY));
    propagator.removeForceModels();
    propagator.addForceModel(new CunninghamAttractionModel(itrf, GravityFieldFactory.getUnnormalizedProvider(ae, mu, TideSystem.UNKNOWN, new double[][] { { 0.0 }, { 0.0 }, { unnormalizedC20 }, { unnormalizedC30 }, { unnormalizedC40 }, { unnormalizedC50 }, { unnormalizedC60 } }, new double[][] { { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 }, { 0.0 } })));
    propagator.setInitialState(new SpacecraftState(orbit));
    SpacecraftState cOrb = propagator.propagate(date.shiftedBy(Constants.JULIAN_DAY));
    Vector3D dif = hfOrb.getPVCoordinates().getPosition().subtract(cOrb.getPVCoordinates().getPosition());
    Assert.assertEquals(0, dif.getNorm(), 2e-9);
    Assert.assertTrue(propagator.getCalls() < 400);
}