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

Example 51 with OrbitType

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

the class FieldNumericalPropagatorTest method doTestResetStateEvent.

private <T extends RealFieldElement<T>> void doTestResetStateEvent(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    // setup
    final FieldAbsoluteDate<T> initDate = FieldAbsoluteDate.getJ2000Epoch(field);
    FieldSpacecraftState<T> initialState;
    FieldNumericalPropagator<T> propagator;
    final FieldVector3D<T> position = new FieldVector3D<>(zero.add(7.0e6), zero.add(1.0e6), zero.add(4.0e6));
    final FieldVector3D<T> velocity = new FieldVector3D<>(zero.add(-500.0), zero.add(8000.0), zero.add(1000.0));
    final FieldOrbit<T> orbit = new FieldEquinoctialOrbit<>(new FieldPVCoordinates<>(position, velocity), FramesFactory.getEME2000(), initDate, mu);
    initialState = new FieldSpacecraftState<>(orbit);
    OrbitType type = OrbitType.EQUINOCTIAL;
    double[][] tolerance = NumericalPropagator.tolerances(0.001, orbit.toOrbit(), type);
    AdaptiveStepsizeFieldIntegrator<T> integrator = new DormandPrince853FieldIntegrator<>(field, 0.001, 200, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(zero.add(60));
    propagator = new FieldNumericalPropagator<>(field, integrator);
    propagator.setOrbitType(type);
    propagator.setInitialState(initialState);
    final FieldAbsoluteDate<T> resetDate = initDate.shiftedBy(1000);
    CheckingHandler<FieldDateDetector<T>, T> checking = new CheckingHandler<FieldDateDetector<T>, T>(Action.RESET_STATE) {

        public FieldSpacecraftState<T> resetState(FieldDateDetector<T> detector, FieldSpacecraftState<T> oldState) {
            return new FieldSpacecraftState<>(oldState.getOrbit(), oldState.getAttitude(), oldState.getMass().subtract(200.0));
        }
    };
    propagator.addEventDetector(new FieldDateDetector<>(resetDate).withHandler(checking));
    checking.assertEvent(false);
    final FieldSpacecraftState<T> finalState = propagator.propagate(initDate.shiftedBy(3200));
    checking.assertEvent(true);
    Assert.assertEquals(initialState.getMass().getReal() - 200, finalState.getMass().getReal(), 1.0e-10);
}

Example 52 with OrbitType

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

the class FieldNumericalPropagatorTest method doTestContinueEvent.

private <T extends RealFieldElement<T>> void doTestContinueEvent(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    // setup
    final FieldAbsoluteDate<T> initDate = FieldAbsoluteDate.getJ2000Epoch(field);
    FieldSpacecraftState<T> initialState;
    FieldNumericalPropagator<T> propagator;
    final FieldVector3D<T> position = new FieldVector3D<>(zero.add(7.0e6), zero.add(1.0e6), zero.add(4.0e6));
    final FieldVector3D<T> velocity = new FieldVector3D<>(zero.add(-500.0), zero.add(8000.0), zero.add(1000.0));
    final FieldOrbit<T> orbit = new FieldEquinoctialOrbit<>(new FieldPVCoordinates<>(position, velocity), FramesFactory.getEME2000(), initDate, mu);
    initialState = new FieldSpacecraftState<>(orbit);
    OrbitType type = OrbitType.EQUINOCTIAL;
    double[][] tolerance = NumericalPropagator.tolerances(0.001, orbit.toOrbit(), type);
    AdaptiveStepsizeFieldIntegrator<T> integrator = new DormandPrince853FieldIntegrator<>(field, 0.001, 200, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(zero.add(60));
    propagator = new FieldNumericalPropagator<>(field, integrator);
    propagator.setOrbitType(type);
    propagator.setInitialState(initialState);
    final FieldAbsoluteDate<T> resetDate = initDate.shiftedBy(1000);
    CheckingHandler<FieldDateDetector<T>, T> checking = new CheckingHandler<FieldDateDetector<T>, T>(Action.CONTINUE);
    propagator.addEventDetector(new FieldDateDetector<>(resetDate).withHandler(checking));
    final double dt = 3200;
    checking.assertEvent(false);
    Assert.assertEquals(0.0, propagator.getInitialState().getDate().durationFrom(initDate).getReal(), 1.0e-10);
    propagator.setResetAtEnd(false);
    final FieldSpacecraftState<T> finalState = propagator.propagate(initDate.shiftedBy(dt));
    Assert.assertEquals(0.0, propagator.getInitialState().getDate().durationFrom(initDate).getReal(), 1.0e-10);
    checking.assertEvent(true);
    final double n = FastMath.sqrt(initialState.getMu() / initialState.getA().getReal()) / initialState.getA().getReal();
    Assert.assertEquals(initialState.getA().getReal(), finalState.getA().getReal(), 1.0e-10);
    Assert.assertEquals(initialState.getEquinoctialEx().getReal(), finalState.getEquinoctialEx().getReal(), 1.0e-10);
    Assert.assertEquals(initialState.getEquinoctialEy().getReal(), finalState.getEquinoctialEy().getReal(), 1.0e-10);
    Assert.assertEquals(initialState.getHx().getReal(), finalState.getHx().getReal(), 1.0e-10);
    Assert.assertEquals(initialState.getHy().getReal(), finalState.getHy().getReal(), 1.0e-10);
    Assert.assertEquals(initialState.getLM().getReal() + n * dt, finalState.getLM().getReal(), 6.0e-10);
}

Example 53 with OrbitType

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

the class FieldNumericalPropagatorTest method doTestCartesian.

private <T extends RealFieldElement<T>> void doTestCartesian(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    // setup
    final FieldAbsoluteDate<T> initDate = FieldAbsoluteDate.getJ2000Epoch(field);
    FieldSpacecraftState<T> initialState;
    FieldNumericalPropagator<T> propagator;
    final FieldVector3D<T> position = new FieldVector3D<>(zero.add(7.0e6), zero.add(1.0e6), zero.add(4.0e6));
    final FieldVector3D<T> velocity = new FieldVector3D<>(zero.add(-500.0), zero.add(8000.0), zero.add(1000.0));
    final FieldOrbit<T> orbit = new FieldEquinoctialOrbit<>(new FieldPVCoordinates<>(position, velocity), FramesFactory.getEME2000(), initDate, mu);
    initialState = new FieldSpacecraftState<>(orbit);
    OrbitType type = OrbitType.EQUINOCTIAL;
    double[][] tolerance = NumericalPropagator.tolerances(0.001, orbit.toOrbit(), type);
    AdaptiveStepsizeFieldIntegrator<T> integrator = new DormandPrince853FieldIntegrator<>(field, 0.001, 200, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(zero.add(60));
    propagator = new FieldNumericalPropagator<>(field, integrator);
    propagator.setOrbitType(type);
    propagator.setInitialState(initialState);
    // Propagation of the initial at t + dt
    final T dt = zero.add(3200);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    final FieldPVCoordinates<T> finalState = propagator.propagate(initDate.shiftedBy(dt)).getPVCoordinates();
    final FieldVector3D<T> pFin = finalState.getPosition();
    final FieldVector3D<T> vFin = finalState.getVelocity();
    // Check results
    final FieldPVCoordinates<T> reference = initialState.shiftedBy(dt).getPVCoordinates();
    final FieldVector3D<T> pRef = reference.getPosition();
    final FieldVector3D<T> vRef = reference.getVelocity();
    Assert.assertEquals(0, pRef.subtract(pFin).getNorm().getReal(), 2e-4);
    Assert.assertEquals(0, vRef.subtract(vFin).getNorm().getReal(), 7e-8);
    try {
        propagator.getGeneratedEphemeris();
        Assert.fail("an exception should have been thrown");
    } catch (IllegalStateException ise) {
    // expected
    }
}

Example 54 with OrbitType

use of org.orekit.orbits.OrbitType 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 55 with OrbitType

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

the class PartialDerivativesTest method doTestParametersDerivatives.

private void doTestParametersDerivatives(String parameterName, double tolerance, OrbitType... orbitTypes) throws OrekitException {
    OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
    ForceModel drag = new DragForce(new HarrisPriester(CelestialBodyFactory.getSun(), earth), new IsotropicDrag(2.5, 1.2));
    NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(5, 5);
    ForceModel gravityField = new HolmesFeatherstoneAttractionModel(FramesFactory.getITRF(IERSConventions.IERS_2010, true), provider);
    Orbit baseOrbit = new KeplerianOrbit(7000000.0, 0.01, 0.1, 0.7, 0, 1.2, PositionAngle.TRUE, FramesFactory.getEME2000(), AbsoluteDate.J2000_EPOCH, provider.getMu());
    double dt = 900;
    double dP = 1.0;
    for (OrbitType orbitType : orbitTypes) {
        final Orbit initialOrbit = orbitType.convertType(baseOrbit);
        for (PositionAngle angleType : PositionAngle.values()) {
            NumericalPropagator propagator = setUpPropagator(initialOrbit, dP, orbitType, angleType, drag, gravityField);
            propagator.setMu(provider.getMu());
            for (final ForceModel forceModel : propagator.getAllForceModels()) {
                for (final ParameterDriver driver : forceModel.getParametersDrivers()) {
                    driver.setValue(driver.getReferenceValue());
                    driver.setSelected(driver.getName().equals(parameterName));
                }
            }
            PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
            final SpacecraftState initialState = partials.setInitialJacobians(new SpacecraftState(initialOrbit));
            propagator.setInitialState(initialState);
            final JacobiansMapper mapper = partials.getMapper();
            PickUpHandler pickUp = new PickUpHandler(mapper, null);
            propagator.setMasterMode(pickUp);
            propagator.propagate(initialState.getDate().shiftedBy(dt));
            double[][] dYdP = pickUp.getdYdP();
            // compute reference Jacobian using finite differences
            double[][] dYdPRef = new double[6][1];
            NumericalPropagator propagator2 = setUpPropagator(initialOrbit, dP, orbitType, angleType, drag, gravityField);
            propagator2.setMu(provider.getMu());
            ParameterDriversList bound = new ParameterDriversList();
            for (final ForceModel forceModel : propagator2.getAllForceModels()) {
                for (final ParameterDriver driver : forceModel.getParametersDrivers()) {
                    if (driver.getName().equals(parameterName)) {
                        driver.setSelected(true);
                        bound.add(driver);
                    } else {
                        driver.setSelected(false);
                    }
                }
            }
            ParameterDriver selected = bound.getDrivers().get(0);
            double p0 = selected.getReferenceValue();
            double h = selected.getScale();
            selected.setValue(p0 - 4 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sM4h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 - 3 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sM3h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 - 2 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sM2h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 - 1 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sM1h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 + 1 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sP1h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 + 2 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sP2h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 + 3 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sP3h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            selected.setValue(p0 + 4 * h);
            propagator2.resetInitialState(arrayToState(stateToArray(initialState, orbitType, angleType, true), orbitType, angleType, initialState.getFrame(), initialState.getDate(), // the mu may have been reset above
            propagator2.getMu(), initialState.getAttitude()));
            SpacecraftState sP4h = propagator2.propagate(initialOrbit.getDate().shiftedBy(dt));
            fillJacobianColumn(dYdPRef, 0, orbitType, angleType, h, sM4h, sM3h, sM2h, sM1h, sP1h, sP2h, sP3h, sP4h);
            for (int i = 0; i < 6; ++i) {
                Assert.assertEquals(dYdPRef[i][0], dYdP[i][0], FastMath.abs(dYdPRef[i][0] * tolerance));
            }
        }
    }
}