Examples with Attitude org.orekit.attitudes.Attitude used on opensource projects

Example 6 with Attitude

use of org.orekit.attitudes.Attitude in project Orekit by CS-SI.

the class ImpulseManeuverTest method testAdditionalStateNumerical.

@Test
public void testAdditionalStateNumerical() throws OrekitException {
    final double mu = CelestialBodyFactory.getEarth().getGM();
    final double initialX = 7100e3;
    final double initialY = 0.0;
    final double initialZ = 1300e3;
    final double initialVx = 0;
    final double initialVy = 8000;
    final double initialVz = 1000;
    final Vector3D position = new Vector3D(initialX, initialY, initialZ);
    final Vector3D velocity = new Vector3D(initialVx, initialVy, initialVz);
    final AbsoluteDate epoch = new AbsoluteDate(2010, 1, 1, 0, 0, 0, TimeScalesFactory.getUTC());
    final TimeStampedPVCoordinates pv = new TimeStampedPVCoordinates(epoch, position, velocity, Vector3D.ZERO);
    final Orbit initialOrbit = new CartesianOrbit(pv, FramesFactory.getEME2000(), mu);
    final double totalPropagationTime = 10.0;
    final double deltaX = 0.01;
    final double deltaY = 0.02;
    final double deltaZ = 0.03;
    final double isp = 300;
    final Vector3D deltaV = new Vector3D(deltaX, deltaY, deltaZ);
    final AttitudeProvider attitudeProvider = new LofOffset(initialOrbit.getFrame(), LOFType.VNC);
    final Attitude initialAttitude = attitudeProvider.getAttitude(initialOrbit, initialOrbit.getDate(), initialOrbit.getFrame());
    double[][] tolerances = NumericalPropagator.tolerances(10.0, initialOrbit, initialOrbit.getType());
    DormandPrince853Integrator integrator = new DormandPrince853Integrator(1.0e-3, 60, tolerances[0], tolerances[1]);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(initialOrbit.getType());
    PartialDerivativesEquations pde = new PartialDerivativesEquations("derivatives", propagator);
    final SpacecraftState initialState = pde.setInitialJacobians(new SpacecraftState(initialOrbit, initialAttitude));
    propagator.resetInitialState(initialState);
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(0.5 * totalPropagationTime));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, RotationConvention.VECTOR_OPERATOR, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector, attitudeOverride, deltaV, isp).withThreshold(1.0e-3);
    propagator.addEventDetector(burnAtEpoch);
    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));
    Assert.assertEquals(1, finalState.getAdditionalStates().size());
    Assert.assertEquals(36, finalState.getAdditionalState("derivatives").length);
    double[][] stateTransitionMatrix = new double[6][6];
    pde.getMapper().getStateJacobian(finalState, stateTransitionMatrix);
    for (int i = 0; i < 6; ++i) {
        for (int j = 0; j < 6; ++j) {
            double sIJ = stateTransitionMatrix[i][j];
            if (j == i) {
                // dPi/dPj and dVi/dVj are roughly 1 for small propagation times
                Assert.assertEquals(1.0, sIJ, 2.0e-4);
            } else if (j == i + 3) {
                // dVi/dPi is roughly the propagation time for small propagation times
                Assert.assertEquals(totalPropagationTime, sIJ, 4.0e-5 * totalPropagationTime);
            } else {
                // other derivatives are almost zero for small propagation times
                Assert.assertEquals(0, sIJ, 1.0e-4);
            }
        }
    }
}

Example 7 with Attitude

use of org.orekit.attitudes.Attitude in project Orekit by CS-SI.

the class ImpulseManeuverTest method testAdditionalStateKeplerian.

@Test
public void testAdditionalStateKeplerian() throws OrekitException {
    final double mu = CelestialBodyFactory.getEarth().getGM();
    final double initialX = 7100e3;
    final double initialY = 0.0;
    final double initialZ = 1300e3;
    final double initialVx = 0;
    final double initialVy = 8000;
    final double initialVz = 1000;
    final Vector3D position = new Vector3D(initialX, initialY, initialZ);
    final Vector3D velocity = new Vector3D(initialVx, initialVy, initialVz);
    final AbsoluteDate epoch = new AbsoluteDate(2010, 1, 1, 0, 0, 0, TimeScalesFactory.getUTC());
    final TimeStampedPVCoordinates pv = new TimeStampedPVCoordinates(epoch, position, velocity, Vector3D.ZERO);
    final Orbit initialOrbit = new CartesianOrbit(pv, FramesFactory.getEME2000(), mu);
    final double totalPropagationTime = 10;
    final double deltaX = 0.01;
    final double deltaY = 0.02;
    final double deltaZ = 0.03;
    final double isp = 300;
    final Vector3D deltaV = new Vector3D(deltaX, deltaY, deltaZ);
    final AttitudeProvider attitudeProvider = new LofOffset(initialOrbit.getFrame(), LOFType.VNC);
    final Attitude initialAttitude = attitudeProvider.getAttitude(initialOrbit, initialOrbit.getDate(), initialOrbit.getFrame());
    final SpacecraftState initialState = new SpacecraftState(initialOrbit, initialAttitude);
    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit);
    propagator.resetInitialState(initialState.addAdditionalState("testOnly", -1.0));
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(0.5 * totalPropagationTime));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, RotationConvention.VECTOR_OPERATOR, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector, attitudeOverride, deltaV, isp).withThreshold(1.0e-3);
    propagator.addEventDetector(burnAtEpoch);
    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));
    Assert.assertEquals(1, finalState.getAdditionalStates().size());
    Assert.assertEquals(-1.0, finalState.getAdditionalState("testOnly")[0], 1.0e-15);
}

Example 8 with Attitude

use of org.orekit.attitudes.Attitude in project Orekit by CS-SI.

the class KeplerianPropagatorTest method tesWrapedAttitudeException.

@Test(expected = OrekitException.class)
public void tesWrapedAttitudeException() throws OrekitException {
    final KeplerianOrbit orbit = new KeplerianOrbit(7.8e6, 0.032, 0.4, 0.1, 0.2, 0.3, PositionAngle.TRUE, FramesFactory.getEME2000(), AbsoluteDate.J2000_EPOCH, 3.986004415e14);
    KeplerianPropagator propagator = new KeplerianPropagator(orbit, new AttitudeProvider() {

        private static final long serialVersionUID = 1L;

        public Attitude getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) throws OrekitException {
            throw new OrekitException((Throwable) null, new DummyLocalizable("dummy error"));
        }

        public <T extends RealFieldElement<T>> FieldAttitude<T> getAttitude(FieldPVCoordinatesProvider<T> pvProv, FieldAbsoluteDate<T> date, Frame frame) throws OrekitException {
            throw new OrekitException((Throwable) null, new DummyLocalizable("dummy error"));
        }
    });
    propagator.propagate(orbit.getDate().shiftedBy(10.09));
}

Example 9 with Attitude

use of org.orekit.attitudes.Attitude in project Orekit by CS-SI.

the class AbstractAnalyticalPropagator method basicPropagate.

/**
 * Propagate an orbit without any fancy features.
 * <p>This method is similar in spirit to the {@link #propagate} method,
 * except that it does <strong>not</strong> call any handler during
 * propagation, nor any discrete events, not additional states. It always
 * stop exactly at the specified date.</p>
 * @param date target date for propagation
 * @return state at specified date
 * @exception OrekitException if propagation cannot reach specified date
 */
protected SpacecraftState basicPropagate(final AbsoluteDate date) throws OrekitException {
    try {
        // evaluate orbit
        final Orbit orbit = propagateOrbit(date);
        // evaluate attitude
        final Attitude attitude = getAttitudeProvider().getAttitude(pvProvider, date, orbit.getFrame());
        // build raw state
        return new SpacecraftState(orbit, attitude, getMass(date));
    } catch (OrekitException oe) {
        throw new OrekitException(oe);
    }
}

Example 10 with Attitude

use of org.orekit.attitudes.Attitude in project Orekit by CS-SI.

the class Ephemeris method basicPropagate.

@Override
public /**
 * {@inheritDoc}
 */
SpacecraftState basicPropagate(final AbsoluteDate date) throws OrekitException {
    final SpacecraftState evaluatedState;
    final AbsoluteDate central;
    if (date.compareTo(minDate) < 0 && FastMath.abs(date.durationFrom(minDate)) <= extrapolationThreshold) {
        // avoid TimeStampedCacheException as we are still within the tolerance before minDate
        central = minDate;
    } else if (date.compareTo(maxDate) > 0 && FastMath.abs(date.durationFrom(maxDate)) <= extrapolationThreshold) {
        // avoid TimeStampedCacheException as we are still within the tolerance after maxDate
        central = maxDate;
    } else {
        central = date;
    }
    final List<SpacecraftState> neighbors = cache.getNeighbors(central).collect(Collectors.toList());
    evaluatedState = neighbors.get(0).interpolate(date, neighbors);
    final AttitudeProvider attitudeProvider = getAttitudeProvider();
    if (attitudeProvider == null) {
        return evaluatedState;
    } else {
        pvProvider.setCurrentState(evaluatedState);
        final Attitude calculatedAttitude = attitudeProvider.getAttitude(pvProvider, date, evaluatedState.getFrame());
        return new SpacecraftState(evaluatedState.getOrbit(), calculatedAttitude, evaluatedState.getMass());
    }
}