Examples with DormandPrince853Integrator org.hipparchus.ode.nonstiff.DormandPrince853Integrator used on opensource projects

Example 61 with DormandPrince853Integrator

use of org.hipparchus.ode.nonstiff.DormandPrince853Integrator in project Orekit by CS-SI.

the class NumericalPropagatorTest method createPropagator.

private static synchronized NumericalPropagator createPropagator(SpacecraftState spacecraftState, OrbitType orbitType, PositionAngle angleType) throws OrekitException {
    final double minStep = 0.001;
    final double maxStep = 120.0;
    final double positionTolerance = 0.1;
    final int degree = 20;
    final int order = 20;
    final double spacecraftArea = 1.0;
    final double spacecraftDragCoefficient = 2.0;
    final double spacecraftReflectionCoefficient = 2.0;
    // propagator main configuration
    final double[][] tol = NumericalPropagator.tolerances(positionTolerance, spacecraftState.getOrbit(), orbitType);
    final ODEIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tol[0], tol[1]);
    final NumericalPropagator np = new NumericalPropagator(integrator);
    np.setOrbitType(orbitType);
    np.setPositionAngleType(angleType);
    np.setInitialState(spacecraftState);
    // Earth gravity field
    final OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
    final NormalizedSphericalHarmonicsProvider harmonicsGravityProvider = GravityFieldFactory.getNormalizedProvider(degree, order);
    np.addForceModel(new HolmesFeatherstoneAttractionModel(earth.getBodyFrame(), harmonicsGravityProvider));
    // Sun and Moon attraction
    np.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getSun()));
    np.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getMoon()));
    // atmospheric drag
    MarshallSolarActivityFutureEstimation msafe = new MarshallSolarActivityFutureEstimation("Jan2000F10-edited-data\\.txt", MarshallSolarActivityFutureEstimation.StrengthLevel.AVERAGE);
    DataProvidersManager.getInstance().feed(msafe.getSupportedNames(), msafe);
    DTM2000 atmosphere = new DTM2000(msafe, CelestialBodyFactory.getSun(), earth);
    np.addForceModel(new DragForce(atmosphere, new IsotropicDrag(spacecraftArea, spacecraftDragCoefficient)));
    // solar radiation pressure
    np.addForceModel(new SolarRadiationPressure(CelestialBodyFactory.getSun(), earth.getEquatorialRadius(), new IsotropicRadiationSingleCoefficient(spacecraftArea, spacecraftReflectionCoefficient)));
    return np;
}

Example 62 with DormandPrince853Integrator

use of org.hipparchus.ode.nonstiff.DormandPrince853Integrator in project Orekit by CS-SI.

the class NumericalPropagatorTest method testEphemerisDatesBackward.

@Test
public void testEphemerisDatesBackward() throws OrekitException {
    // setup
    TimeScale tai = TimeScalesFactory.getTAI();
    AbsoluteDate initialDate = new AbsoluteDate("2015-07-05", tai);
    AbsoluteDate startDate = new AbsoluteDate("2015-07-03", tai).shiftedBy(-0.1);
    AbsoluteDate endDate = new AbsoluteDate("2015-07-04", tai);
    Frame eci = FramesFactory.getGCRF();
    KeplerianOrbit orbit = new KeplerianOrbit(600e3 + Constants.WGS84_EARTH_EQUATORIAL_RADIUS, 0, 0, 0, 0, 0, PositionAngle.TRUE, eci, initialDate, mu);
    OrbitType type = OrbitType.CARTESIAN;
    double[][] tol = NumericalPropagator.tolerances(1e-3, orbit, type);
    NumericalPropagator prop = new NumericalPropagator(new DormandPrince853Integrator(0.1, 500, tol[0], tol[1]));
    prop.setOrbitType(type);
    prop.resetInitialState(new SpacecraftState(new CartesianOrbit(orbit)));
    // action
    prop.setEphemerisMode();
    prop.propagate(endDate, startDate);
    BoundedPropagator ephemeris = prop.getGeneratedEphemeris();
    // verify
    TimeStampedPVCoordinates actualPV = ephemeris.getPVCoordinates(startDate, eci);
    TimeStampedPVCoordinates expectedPV = orbit.getPVCoordinates(startDate, eci);
    MatcherAssert.assertThat(actualPV.getPosition(), OrekitMatchers.vectorCloseTo(expectedPV.getPosition(), 1.0));
    MatcherAssert.assertThat(actualPV.getVelocity(), OrekitMatchers.vectorCloseTo(expectedPV.getVelocity(), 1.0));
    MatcherAssert.assertThat(ephemeris.getMinDate().durationFrom(startDate), OrekitMatchers.closeTo(0, 0));
    MatcherAssert.assertThat(ephemeris.getMaxDate().durationFrom(endDate), OrekitMatchers.closeTo(0, 0));
    // test date
    AbsoluteDate date = endDate.shiftedBy(-0.11);
    Assert.assertEquals(ephemeris.propagate(date).getDate().durationFrom(date), 0, 0);
}

Example 63 with DormandPrince853Integrator

use of org.hipparchus.ode.nonstiff.DormandPrince853Integrator in project Orekit by CS-SI.

the class PartialDerivativesTest method testJacobianIssue18.

@Test
public void testJacobianIssue18() throws OrekitException {
    // Body mu
    final double mu = 3.9860047e14;
    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 double delta = FastMath.toRadians(-7.4978);
    final double alpha = FastMath.toRadians(351);
    final AttitudeProvider law = new InertialProvider(new Rotation(new Vector3D(alpha, delta), Vector3D.PLUS_I));
    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 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);
    double[] absTolerance = { 0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001 };
    double[] relTolerance = { 1.0e-7, 1.0e-4, 1.0e-4, 1.0e-7, 1.0e-7, 1.0e-7, 1.0e-7 };
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 1000, absTolerance, relTolerance);
    integrator.setInitialStepSize(60);
    final NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setAttitudeProvider(law);
    propagator.addForceModel(maneuver);
    maneuver.getParameterDriver("thrust").setSelected(true);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    PartialDerivativesEquations PDE = new PartialDerivativesEquations("derivatives", propagator);
    Assert.assertEquals(1, PDE.getSelectedParameters().getNbParams());
    propagator.setInitialState(PDE.setInitialJacobians(initialState));
    final AbsoluteDate finalDate = fireDate.shiftedBy(3800);
    final SpacecraftState finalorb = propagator.propagate(finalDate);
    Assert.assertEquals(0, finalDate.durationFrom(finalorb.getDate()), 1.0e-11);
}

Example 64 with DormandPrince853Integrator

use of org.hipparchus.ode.nonstiff.DormandPrince853Integrator in project Orekit by CS-SI.

the class PartialDerivativesTest method setUpPropagator.

private NumericalPropagator setUpPropagator(Orbit orbit, double dP, OrbitType orbitType, PositionAngle angleType, ForceModel... models) throws OrekitException {
    final double minStep = 0.001;
    final double maxStep = 1000;
    double[][] tol = NumericalPropagator.tolerances(dP, orbit, orbitType);
    NumericalPropagator propagator = new NumericalPropagator(new DormandPrince853Integrator(minStep, maxStep, tol[0], tol[1]));
    propagator.setOrbitType(orbitType);
    propagator.setPositionAngleType(angleType);
    for (ForceModel model : models) {
        propagator.addForceModel(model);
    }
    return propagator;
}

Example 65 with DormandPrince853Integrator

use of org.hipparchus.ode.nonstiff.DormandPrince853Integrator in project Orekit by CS-SI.

the class DSSTPropagatorTest method testIssue157.

@Test
public void testIssue157() throws OrekitException {
    Utils.setDataRoot("regular-data:potential/icgem-format");
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("^eigen-6s-truncated$", false));
    UnnormalizedSphericalHarmonicsProvider nshp = GravityFieldFactory.getUnnormalizedProvider(8, 8);
    Orbit orbit = new KeplerianOrbit(13378000, 0.05, 0, 0, FastMath.PI, 0, PositionAngle.MEAN, FramesFactory.getTOD(false), new AbsoluteDate(2003, 5, 6, TimeScalesFactory.getUTC()), nshp.getMu());
    double period = orbit.getKeplerianPeriod();
    double[][] tolerance = DSSTPropagator.tolerances(1.0, orbit);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(period / 100, period * 100, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(10 * period);
    DSSTPropagator propagator = new DSSTPropagator(integrator, true);
    OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getGTOD(false));
    CelestialBody sun = CelestialBodyFactory.getSun();
    CelestialBody moon = CelestialBodyFactory.getMoon();
    propagator.addForceModel(new DSSTZonal(nshp, 8, 7, 17));
    propagator.addForceModel(new DSSTTesseral(earth.getBodyFrame(), Constants.WGS84_EARTH_ANGULAR_VELOCITY, nshp, 8, 8, 4, 12, 8, 8, 4));
    propagator.addForceModel(new DSSTThirdBody(sun));
    propagator.addForceModel(new DSSTThirdBody(moon));
    propagator.addForceModel(new DSSTAtmosphericDrag(new HarrisPriester(sun, earth), 2.1, 180));
    propagator.addForceModel(new DSSTSolarRadiationPressure(1.2, 180, sun, earth.getEquatorialRadius()));
    propagator.setInitialState(new SpacecraftState(orbit, 45.0), true);
    SpacecraftState finalState = propagator.propagate(orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY));
    // the following comparison is in fact meaningless
    // the initial orbit is osculating the final orbit is a mean orbit
    // and they are not considered at the same epoch
    // we keep it only as is was an historical test
    Assert.assertEquals(2189.4, orbit.getA() - finalState.getA(), 1.0);
    propagator.setInitialState(new SpacecraftState(orbit, 45.0), false);
    finalState = propagator.propagate(orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY));
    // the following comparison is realistic
    // both the initial orbit and final orbit are mean orbits
    Assert.assertEquals(1478.05, orbit.getA() - finalState.getA(), 1.0);
}