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

Example 11 with FieldEquinoctialOrbit

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

the class FieldNumericalPropagatorTest method doTestAdditionalStateEvent.

private <T extends RealFieldElement<T>> void doTestAdditionalStateEvent(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);
    propagator.addAdditionalEquations(new FieldAdditionalEquations<T>() {

        public String getName() {
            return "linear";
        }

        public T[] computeDerivatives(FieldSpacecraftState<T> s, T[] pDot) {
            pDot[0] = zero.add(1.0);
            return MathArrays.buildArray(field, 7);
        }
    });
    try {
        propagator.addAdditionalEquations(new FieldAdditionalEquations<T>() {

            public String getName() {
                return "linear";
            }

            public T[] computeDerivatives(FieldSpacecraftState<T> s, T[] pDot) {
                pDot[0] = zero.add(1.0);
                return MathArrays.buildArray(field, 7);
            }
        });
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(oe.getSpecifier(), OrekitMessages.ADDITIONAL_STATE_NAME_ALREADY_IN_USE);
    }
    try {
        propagator.addAdditionalStateProvider(new FieldAdditionalStateProvider<T>() {

            public String getName() {
                return "linear";
            }

            public T[] getAdditionalState(FieldSpacecraftState<T> state) {
                return null;
            }
        });
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(oe.getSpecifier(), OrekitMessages.ADDITIONAL_STATE_NAME_ALREADY_IN_USE);
    }
    propagator.addAdditionalStateProvider(new FieldAdditionalStateProvider<T>() {

        public String getName() {
            return "constant";
        }

        public T[] getAdditionalState(FieldSpacecraftState<T> state) {
            T[] ret = MathArrays.buildArray(field, 1);
            ret[0] = zero.add(1.0);
            return ret;
        }
    });
    Assert.assertTrue(propagator.isAdditionalStateManaged("linear"));
    Assert.assertTrue(propagator.isAdditionalStateManaged("constant"));
    Assert.assertFalse(propagator.isAdditionalStateManaged("non-managed"));
    Assert.assertEquals(2, propagator.getManagedAdditionalStates().length);
    propagator.setInitialState(propagator.getInitialState().addAdditionalState("linear", zero.add(1.5)));
    CheckingHandler<AdditionalStateLinearDetector<T>, T> checking = new CheckingHandler<AdditionalStateLinearDetector<T>, T>(Action.STOP);
    propagator.addEventDetector(new AdditionalStateLinearDetector<T>(zero.add(10.0), zero.add(1.0e-8)).withHandler(checking));
    final double dt = 3200;
    checking.assertEvent(false);
    final FieldSpacecraftState<T> finalState = propagator.propagate(initDate.shiftedBy(dt));
    checking.assertEvent(true);
    Assert.assertEquals(3.0, finalState.getAdditionalState("linear")[0].getReal(), 1.0e-8);
    Assert.assertEquals(1.5, finalState.getDate().durationFrom(initDate).getReal(), 1.0e-8);
}

Example 12 with FieldEquinoctialOrbit

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

the class FieldNumericalPropagatorTest method doTestPropagationTypesHyperbolic.

private <T extends RealFieldElement<T>> void doTestPropagationTypesHyperbolic(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);
    FieldSpacecraftState<T> state = new FieldSpacecraftState<>(new FieldKeplerianOrbit<>(zero.add(-10000000.0), zero.add(2.5), zero.add(0.3), zero, zero, zero, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, mu));
    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 FieldPVCoordinates<T> pv = state.getPVCoordinates();
    final T dP = zero.add(0.001);
    final T dV = dP.multiply(state.getMu()).divide(pv.getPosition().getNormSq().multiply(pv.getVelocity().getNorm()));
    final FieldPVCoordinates<T> pvcM = propagateInType(state, dP, OrbitType.CARTESIAN, PositionAngle.MEAN, propagator);
    final FieldPVCoordinates<T> pvkM = propagateInType(state, dP, OrbitType.KEPLERIAN, PositionAngle.MEAN, propagator);
    final FieldPVCoordinates<T> pvcE = propagateInType(state, dP, OrbitType.CARTESIAN, PositionAngle.ECCENTRIC, propagator);
    final FieldPVCoordinates<T> pvkE = propagateInType(state, dP, OrbitType.KEPLERIAN, PositionAngle.ECCENTRIC, propagator);
    final FieldPVCoordinates<T> pvcT = propagateInType(state, dP, OrbitType.CARTESIAN, PositionAngle.TRUE, propagator);
    final FieldPVCoordinates<T> pvkT = propagateInType(state, dP, OrbitType.KEPLERIAN, PositionAngle.TRUE, propagator);
    Assert.assertEquals(0, pvcM.getPosition().subtract(pvkT.getPosition()).getNorm().getReal() / dP.getReal(), 0.3);
    Assert.assertEquals(0, pvcM.getVelocity().subtract(pvkT.getVelocity()).getNorm().getReal() / dV.getReal(), 0.4);
    Assert.assertEquals(0, pvkM.getPosition().subtract(pvkT.getPosition()).getNorm().getReal() / dP.getReal(), 0.2);
    Assert.assertEquals(0, pvkM.getVelocity().subtract(pvkT.getVelocity()).getNorm().getReal() / dV.getReal(), 0.3);
    Assert.assertEquals(0, pvcE.getPosition().subtract(pvkT.getPosition()).getNorm().getReal() / dP.getReal(), 0.3);
    Assert.assertEquals(0, pvcE.getVelocity().subtract(pvkT.getVelocity()).getNorm().getReal() / dV.getReal(), 0.4);
    Assert.assertEquals(0, pvkE.getPosition().subtract(pvkT.getPosition()).getNorm().getReal() / dP.getReal(), 0.009);
    Assert.assertEquals(0, pvkE.getVelocity().subtract(pvkT.getVelocity()).getNorm().getReal() / dV.getReal(), 0.006);
    Assert.assertEquals(0, pvcT.getPosition().subtract(pvkT.getPosition()).getNorm().getReal() / dP.getReal(), 0.3);
    Assert.assertEquals(0, pvcT.getVelocity().subtract(pvkT.getVelocity()).getNorm().getReal() / dV.getReal(), 0.4);
}

Example 13 with FieldEquinoctialOrbit

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

the class FieldNumericalPropagatorTest method doTestKepler.

private <T extends RealFieldElement<T>> void doTestKepler(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 double dt = 3200;
    final FieldSpacecraftState<T> finalState = propagator.propagate(initDate.shiftedBy(-60), initDate.shiftedBy(dt));
    // Check results
    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(), 2.0e-9);
}

Example 14 with FieldEquinoctialOrbit

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

the class FieldNumericalPropagatorTest method createPropagator.

private <T extends RealFieldElement<T>> FieldNumericalPropagator<T> createPropagator(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    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));
    FieldAbsoluteDate<T> initDate = FieldAbsoluteDate.getJ2000Epoch(field);
    final FieldOrbit<T> orbit = new FieldEquinoctialOrbit<>(new FieldPVCoordinates<>(position, velocity), FramesFactory.getEME2000(), initDate, mu);
    FieldSpacecraftState<T> initialState = new FieldSpacecraftState<>(orbit);
    OrbitType type = OrbitType.EQUINOCTIAL;
    double[][] tolerance = FieldNumericalPropagator.tolerances(zero.add(0.001), orbit, type);
    AdaptiveStepsizeFieldIntegrator<T> integrator = new DormandPrince853FieldIntegrator<>(field, 0.001, 200, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(zero.add(60));
    FieldNumericalPropagator<T> propagator = new FieldNumericalPropagator<>(field, integrator);
    propagator.setOrbitType(type);
    propagator.setInitialState(initialState);
    return propagator;
}

Example 15 with FieldEquinoctialOrbit

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

the class FieldEcksteinHechlerPropagatorTest method doAlmostSphericalBody.

private <T extends RealFieldElement<T>> void doAlmostSphericalBody(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    FieldAbsoluteDate<T> date = new FieldAbsoluteDate<>(field);
    // Definition of initial conditions
    // ---------------------------------
    // with e around e = 1.4e-4 and i = 1.7 rad
    FieldVector3D<T> position = new FieldVector3D<>(zero.add(3220103.), zero.add(69623.), zero.add(6449822.));
    FieldVector3D<T> velocity = new FieldVector3D<>(zero.add(6414.7), zero.add(-2006.), zero.add(-3180.));
    FieldAbsoluteDate<T> initDate = date.shiftedBy(584.);
    FieldOrbit<T> initialOrbit = new FieldEquinoctialOrbit<>(new FieldPVCoordinates<>(position, velocity), FramesFactory.getEME2000(), initDate, provider.getMu());
    // Initialisation to simulate a Keplerian extrapolation
    // To be noticed: in order to simulate a Keplerian extrapolation with the
    // analytical
    // extrapolator, one should put the zonal coefficients to 0. But due to
    // numerical pbs
    // one must put a non 0 value.
    UnnormalizedSphericalHarmonicsProvider kepProvider = GravityFieldFactory.getUnnormalizedProvider(6.378137e6, 3.9860047e14, TideSystem.UNKNOWN, new double[][] { { 0 }, { 0 }, { 0.1e-10 }, { 0.1e-13 }, { 0.1e-13 }, { 0.1e-14 }, { 0.1e-14 } }, new double[][] { { 0 }, { 0 }, { 0 }, { 0 }, { 0 }, { 0 }, { 0 } });
    // Extrapolators definitions
    // -------------------------
    FieldEcksteinHechlerPropagator<T> extrapolatorAna = new FieldEcksteinHechlerPropagator<>(initialOrbit, kepProvider);
    FieldKeplerianPropagator<T> extrapolatorKep = new FieldKeplerianPropagator<>(initialOrbit);
    // Extrapolation at a final date different from initial date
    // ---------------------------------------------------------
    // extrapolation duration in seconds
    double delta_t = 100.0;
    FieldAbsoluteDate<T> extrapDate = initDate.shiftedBy(delta_t);
    FieldSpacecraftState<T> finalOrbitAna = extrapolatorAna.propagate(extrapDate);
    FieldSpacecraftState<T> finalOrbitKep = extrapolatorKep.propagate(extrapDate);
    Assert.assertEquals(finalOrbitAna.getDate().durationFrom(extrapDate).getReal(), 0.0, Utils.epsilonTest);
    // comparison of each orbital parameters
    Assert.assertEquals(finalOrbitAna.getA().getReal(), finalOrbitKep.getA().getReal(), 10 * Utils.epsilonTest * finalOrbitKep.getA().getReal());
    Assert.assertEquals(finalOrbitAna.getEquinoctialEx().getReal(), finalOrbitKep.getEquinoctialEx().getReal(), Utils.epsilonE * finalOrbitKep.getE().getReal());
    Assert.assertEquals(finalOrbitAna.getEquinoctialEy().getReal(), finalOrbitKep.getEquinoctialEy().getReal(), Utils.epsilonE * finalOrbitKep.getE().getReal());
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbitAna.getHx().getReal(), finalOrbitKep.getHx().getReal()), finalOrbitKep.getHx().getReal(), Utils.epsilonAngle * FastMath.abs(finalOrbitKep.getI().getReal()));
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbitAna.getHy().getReal(), finalOrbitKep.getHy().getReal()), finalOrbitKep.getHy().getReal(), Utils.epsilonAngle * FastMath.abs(finalOrbitKep.getI().getReal()));
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbitAna.getLv().getReal(), finalOrbitKep.getLv().getReal()), finalOrbitKep.getLv().getReal(), Utils.epsilonAngle * FastMath.abs(finalOrbitKep.getLv().getReal()));
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbitAna.getLE().getReal(), finalOrbitKep.getLE().getReal()), finalOrbitKep.getLE().getReal(), Utils.epsilonAngle * FastMath.abs(finalOrbitKep.getLE().getReal()));
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbitAna.getLM().getReal(), finalOrbitKep.getLM().getReal()), finalOrbitKep.getLM().getReal(), Utils.epsilonAngle * FastMath.abs(finalOrbitKep.getLM().getReal()));
}