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

Example 36 with AdaptiveStepsizeIntegrator

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

the class ThirdBodyAttractionTest 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, 5);
    DerivativeStructure a_0 = factory.variable(0, 7e7);
    DerivativeStructure e_0 = factory.variable(1, 0.4);
    DerivativeStructure i_0 = factory.variable(2, 85 * FastMath.PI / 180);
    DerivativeStructure R_0 = factory.variable(3, 0.7);
    DerivativeStructure O_0 = factory.variable(4, 0.5);
    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.KEPLERIAN;
    double[][] tolerance = NumericalPropagator.tolerances(0.001, 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);
    final ThirdBodyAttraction forceModel = new ThirdBodyAttraction(CelestialBodyFactory.getSun());
    FNP.addForceModel(forceModel);
    // NOT ADDING THE FORCE MODEL TO THE NUMERICAL PROPAGATOR   NP.addForceModel(forceModel);
    FieldAbsoluteDate<DerivativeStructure> target = J2000.shiftedBy(1000.);
    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 37 with AdaptiveStepsizeIntegrator

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

the class NumericalPropagatorTest method testEventDetectionBug.

@Test
public void testEventDetectionBug() throws OrekitException, IOException, ParseException {
    TimeScale utc = TimeScalesFactory.getUTC();
    AbsoluteDate initialDate = new AbsoluteDate(2005, 1, 1, 0, 0, 0.0, utc);
    double duration = 100000.;
    AbsoluteDate endDate = new AbsoluteDate(initialDate, duration);
    // Initialization of the frame EME2000
    Frame EME2000 = FramesFactory.getEME2000();
    // Initial orbit
    double a = 35786000. + 6378137.0;
    double e = 0.70;
    double rApogee = a * (1 + e);
    double vApogee = FastMath.sqrt(mu * (1 - e) / (a * (1 + e)));
    Orbit geo = new CartesianOrbit(new PVCoordinates(new Vector3D(rApogee, 0., 0.), new Vector3D(0., vApogee, 0.)), EME2000, initialDate, mu);
    duration = geo.getKeplerianPeriod();
    endDate = new AbsoluteDate(initialDate, duration);
    // Numerical Integration
    final double minStep = 0.001;
    final double maxStep = 1000;
    final double initStep = 60;
    final double[] absTolerance = { 0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001 };
    final 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(minStep, maxStep, absTolerance, relTolerance);
    integrator.setInitialStepSize(initStep);
    // Numerical propagator based on the integrator
    propagator = new NumericalPropagator(integrator);
    double mass = 1000.;
    SpacecraftState initialState = new SpacecraftState(geo, mass);
    propagator.setInitialState(initialState);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    // Set the events Detectors
    ApsideDetector event1 = new ApsideDetector(geo);
    propagator.addEventDetector(event1);
    // Set the propagation mode
    propagator.setSlaveMode();
    // Propagate
    SpacecraftState finalState = propagator.propagate(endDate);
    // we should stop long before endDate
    Assert.assertTrue(endDate.durationFrom(finalState.getDate()) > 40000.0);
}

Example 38 with AdaptiveStepsizeIntegrator

use of org.hipparchus.ode.nonstiff.AdaptiveStepsizeIntegrator 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 39 with AdaptiveStepsizeIntegrator

use of org.hipparchus.ode.nonstiff.AdaptiveStepsizeIntegrator 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);
}

Example 40 with AdaptiveStepsizeIntegrator

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

the class DSSTPropagatorTest method testShortPeriodCoefficients.

@Test
public void testShortPeriodCoefficients() throws OrekitException {
    Utils.setDataRoot("regular-data:potential/icgem-format");
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("^eigen-6s-truncated$", false));
    UnnormalizedSphericalHarmonicsProvider nshp = GravityFieldFactory.getUnnormalizedProvider(4, 4);
    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, false);
    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, 4, 3, 9));
    propagator.addForceModel(new DSSTTesseral(earth.getBodyFrame(), Constants.WGS84_EARTH_ANGULAR_VELOCITY, nshp, 4, 4, 4, 8, 4, 4, 2));
    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()));
    final AbsoluteDate finalDate = orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY);
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateNoConfig = propagator.propagate(finalDate);
    Assert.assertEquals(0, stateNoConfig.getAdditionalStates().size());
    propagator.setSelectedCoefficients(new HashSet<String>());
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateConfigEmpty = propagator.propagate(finalDate);
    Assert.assertEquals(234, stateConfigEmpty.getAdditionalStates().size());
    final Set<String> selected = new HashSet<String>();
    selected.add("DSST-3rd-body-Moon-s[7]");
    selected.add("DSST-central-body-tesseral-c[-2][3]");
    propagator.setSelectedCoefficients(selected);
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateConfigeSelected = propagator.propagate(finalDate);
    Assert.assertEquals(selected.size(), stateConfigeSelected.getAdditionalStates().size());
    propagator.setSelectedCoefficients(null);
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateConfigNull = propagator.propagate(finalDate);
    Assert.assertEquals(0, stateConfigNull.getAdditionalStates().size());
}