Examples with Rotation org.hipparchus.geometry.euclidean.threed.Rotation used on opensource projects

Example 46 with Rotation

use of org.hipparchus.geometry.euclidean.threed.Rotation in project Orekit by CS-SI.

the class MODProviderTest method testEuler1976.

@Test
public void testEuler1976() throws OrekitException {
    TransformProvider eulerBasedProvider = new TransformProvider() {

        private static final long serialVersionUID = 1L;

        private final PolynomialNutation zetaA = new PolynomialNutation(0.0, 2306.2181 * Constants.ARC_SECONDS_TO_RADIANS, 0.30188 * Constants.ARC_SECONDS_TO_RADIANS, 0.017998 * Constants.ARC_SECONDS_TO_RADIANS);

        private final PolynomialNutation thetaA = new PolynomialNutation(0.0, 2004.3109 * Constants.ARC_SECONDS_TO_RADIANS, -0.42665 * Constants.ARC_SECONDS_TO_RADIANS, -0.041833 * Constants.ARC_SECONDS_TO_RADIANS);

        private final PolynomialNutation zA = new PolynomialNutation(0.0, 2306.2181 * Constants.ARC_SECONDS_TO_RADIANS, 1.09468 * Constants.ARC_SECONDS_TO_RADIANS, 0.018203 * Constants.ARC_SECONDS_TO_RADIANS);

        public Transform getTransform(AbsoluteDate date) {
            final double tc = IERSConventions.IERS_1996.evaluateTC(date);
            final Rotation r1 = new Rotation(Vector3D.PLUS_K, zA.value(tc), RotationConvention.VECTOR_OPERATOR);
            final Rotation r2 = new Rotation(Vector3D.PLUS_J, -thetaA.value(tc), RotationConvention.VECTOR_OPERATOR);
            final Rotation r3 = new Rotation(Vector3D.PLUS_K, zetaA.value(tc), RotationConvention.VECTOR_OPERATOR);
            return new Transform(date, r1.compose(r2.compose(r3, RotationConvention.VECTOR_OPERATOR), RotationConvention.VECTOR_OPERATOR));
        }

        public <T extends RealFieldElement<T>> FieldTransform<T> getTransform(final FieldAbsoluteDate<T> date) {
            throw new UnsupportedOperationException("never called in this test");
        }
    };
    MODProvider modProvider = new MODProvider(IERSConventions.IERS_1996);
    for (double dt = -5 * Constants.JULIAN_YEAR; dt < 5 * Constants.JULIAN_YEAR; dt += 10 * Constants.JULIAN_DAY) {
        AbsoluteDate date = AbsoluteDate.J2000_EPOCH.shiftedBy(dt);
        Transform t = new Transform(date, modProvider.getTransform(date).getInverse(), eulerBasedProvider.getTransform(date));
        Assert.assertEquals(0, t.getRotation().getAngle(), 1.01e-11);
    }
}

Example 47 with Rotation

use of org.hipparchus.geometry.euclidean.threed.Rotation in project Orekit by CS-SI.

the class GTODProviderTest method testAASReferenceLEO.

@Test
public void testAASReferenceLEO() throws OrekitException {
    // this reference test has been extracted from the following paper:
    // Implementation Issues Surrounding the New IAU Reference Systems for Astrodynamics
    // David A. Vallado, John H. Seago, P. Kenneth Seidelmann
    // http://www.centerforspace.com/downloads/files/pubs/AAS-06-134.pdf
    Utils.setLoaders(IERSConventions.IERS_1996, Utils.buildEOPList(IERSConventions.IERS_1996, ITRFVersion.ITRF_2008, new double[][] { { 53098, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53099, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53100, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53101, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53102, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53103, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53104, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN }, { 53105, -0.4399619, 0.0015563, -0.140682, 0.333309, -0.052195, -0.003875, Double.NaN, Double.NaN } }));
    AbsoluteDate t0 = new AbsoluteDate(new DateComponents(2004, 04, 06), new TimeComponents(07, 51, 28.386009), TimeScalesFactory.getUTC());
    // PEF iau76
    PVCoordinates pvPEF = new PVCoordinates(new Vector3D(-1033475.0313, 7901305.5856, 6380344.5328), new Vector3D(-3225.632747, -2872.442511, 5531.931288));
    // it seems the induced effect of pole nutation correction δΔψ on the equation of the equinoxes
    // was not taken into account in the reference paper, so we fix it here for the test
    final double dDeltaPsi = FramesFactory.getEOPHistory(IERSConventions.IERS_1996, true).getEquinoxNutationCorrection(t0)[0];
    final double epsilonA = IERSConventions.IERS_1996.getMeanObliquityFunction().value(t0);
    final Transform fix = new Transform(t0, new Rotation(Vector3D.PLUS_K, dDeltaPsi * FastMath.cos(epsilonA), RotationConvention.FRAME_TRANSFORM));
    // TOD iau76
    PVCoordinates pvTOD = new PVCoordinates(new Vector3D(5094514.7804, 6127366.4612, 6380344.5328), new Vector3D(-4746.088567, 786.077222, 5531.931288));
    Transform t = FramesFactory.getTOD(IERSConventions.IERS_1996, true).getTransformTo(FramesFactory.getGTOD(IERSConventions.IERS_1996, true), t0);
    checkPV(fix.transformPVCoordinates(pvPEF), t.transformPVCoordinates(pvTOD), 0.00942, 3.12e-5);
    // if we forget to apply nutation corrections, results are much worse, which is expected
    t = FramesFactory.getTOD(false).getTransformTo(FramesFactory.getGTOD(false), t0);
    checkPV(fix.transformPVCoordinates(pvPEF), t.transformPVCoordinates(pvTOD), 257.49, 0.13955);
}

Example 48 with Rotation

use of org.hipparchus.geometry.euclidean.threed.Rotation in project Orekit by CS-SI.

the class GroundStation method getOffsetToInertial.

/**
 * Get the transform between offset frame and inertial frame.
 * <p>
 * The offset frame takes the <em>current</em> position offset,
 * polar motion and the meridian shift into account. The frame
 * returned is disconnected from later changes in the parameters.
 * When the {@link ParameterDriver parameters} managing these
 * offsets are changed, the method must be called again to retrieve
 * a new offset frame.
 * </p>
 * @param inertial inertial frame to transform to
 * @param date date of the transform
 * @return offset frame defining vectors
 * @exception OrekitException if offset frame cannot be computed for current offset values
 */
public Transform getOffsetToInertial(final Frame inertial, final AbsoluteDate date) throws OrekitException {
    // take Earth offsets into account
    final Transform intermediateToBody = estimatedEarthFrameProvider.getTransform(date).getInverse();
    // take station offset into account
    final double x = parametricModel(eastOffsetDriver);
    final double y = parametricModel(northOffsetDriver);
    final double z = parametricModel(zenithOffsetDriver);
    final BodyShape baseShape = baseFrame.getParentShape();
    final Transform baseToBody = baseFrame.getTransformTo(baseShape.getBodyFrame(), date);
    Vector3D origin = baseToBody.transformPosition(new Vector3D(x, y, z));
    origin = origin.add(computeDisplacement(date, origin));
    final GeodeticPoint originGP = baseShape.transform(origin, baseShape.getBodyFrame(), date);
    final Transform offsetToIntermediate = new Transform(date, new Transform(date, new Rotation(Vector3D.PLUS_I, Vector3D.PLUS_K, originGP.getEast(), originGP.getZenith()), Vector3D.ZERO), new Transform(date, origin));
    // combine all transforms together
    final Transform bodyToInert = baseFrame.getParent().getTransformTo(inertial, date);
    return new Transform(date, offsetToIntermediate, new Transform(date, intermediateToBody, bodyToInert));
}

Example 49 with Rotation

use of org.hipparchus.geometry.euclidean.threed.Rotation in project Orekit by CS-SI.

the class EcksteinHechlerPropagatorTest method testInitializationCorrectness.

@Test
public void testInitializationCorrectness() throws OrekitException, IOException {
    // Definition of initial conditions
    AbsoluteDate date = AbsoluteDate.J2000_EPOCH.shiftedBy(154.);
    Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
    Frame eme2000 = FramesFactory.getEME2000();
    Vector3D pole = itrf.getTransformTo(eme2000, date).transformVector(Vector3D.PLUS_K);
    Frame poleAligned = new Frame(FramesFactory.getEME2000(), new Transform(date, new Rotation(pole, Vector3D.PLUS_K)), "pole aligned", true);
    CircularOrbit initial = new CircularOrbit(7208669.8179538045, 1.3740461966386876E-4, -3.2364250248363356E-5, FastMath.toRadians(97.40236024565775), FastMath.toRadians(166.15873160992115), FastMath.toRadians(90.1282370098961), PositionAngle.MEAN, poleAligned, date, provider.getMu());
    // find the default Eckstein-Hechler propagator initialized from the initial orbit
    EcksteinHechlerPropagator defaultEH = new EcksteinHechlerPropagator(initial, provider);
    // the osculating parameters recomputed by the default Eckstein-Hechler propagator are quite different
    // from initial orbit
    CircularOrbit defaultOrbit = (CircularOrbit) OrbitType.CIRCULAR.convertType(defaultEH.propagateOrbit(initial.getDate()));
    Assert.assertEquals(267.4, defaultOrbit.getA() - initial.getA(), 0.1);
    // the position on the other hand match perfectly
    Assert.assertEquals(0.0, Vector3D.distance(defaultOrbit.getPVCoordinates().getPosition(), initial.getPVCoordinates().getPosition()), 1.0e-8);
    // set up a reference numerical propagator starting for the specified start orbit
    // using the same force models (i.e. the first few zonal terms)
    double[][] tol = NumericalPropagator.tolerances(0.1, initial, OrbitType.CIRCULAR);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator.setInitialStepSize(60);
    NumericalPropagator num = new NumericalPropagator(integrator);
    num.addForceModel(new HolmesFeatherstoneAttractionModel(itrf, GravityFieldFactory.getNormalizedProvider(provider)));
    num.setInitialState(new SpacecraftState(initial));
    num.setOrbitType(OrbitType.CIRCULAR);
    // find the best Eckstein-Hechler propagator that match the orbit evolution
    PropagatorConverter converter = new FiniteDifferencePropagatorConverter(new EcksteinHechlerPropagatorBuilder(initial, provider, PositionAngle.TRUE, 1.0), 1.0e-6, 100);
    EcksteinHechlerPropagator fittedEH = (EcksteinHechlerPropagator) converter.convert(num, 3 * initial.getKeplerianPeriod(), 300);
    // the default Eckstein-Hechler propagator did however quite a good job, as it found
    // an orbit close to the best fitting
    CircularOrbit fittedOrbit = (CircularOrbit) OrbitType.CIRCULAR.convertType(fittedEH.propagateOrbit(initial.getDate()));
    Assert.assertEquals(0.623, defaultOrbit.getA() - fittedOrbit.getA(), 0.1);
    // the position on the other hand are slightly different
    // because the fitted orbit minimizes the residuals over a complete time span,
    // not on a single point
    Assert.assertEquals(58.0, Vector3D.distance(defaultOrbit.getPVCoordinates().getPosition(), fittedOrbit.getPVCoordinates().getPosition()), 0.1);
}

Example 50 with Rotation

use of org.hipparchus.geometry.euclidean.threed.Rotation in project Orekit by CS-SI.

the class EcksteinHechlerPropagatorTest method testIssue223.

@Test
public void testIssue223() throws OrekitException, IOException, ClassNotFoundException {
    // Definition of initial conditions
    AbsoluteDate date = AbsoluteDate.J2000_EPOCH.shiftedBy(154.);
    Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
    Frame eme2000 = FramesFactory.getEME2000();
    Vector3D pole = itrf.getTransformTo(eme2000, date).transformVector(Vector3D.PLUS_K);
    Frame poleAligned = new Frame(FramesFactory.getEME2000(), new Transform(date, new Rotation(pole, Vector3D.PLUS_K)), "pole aligned", true);
    CircularOrbit initial = new CircularOrbit(7208669.8179538045, 1.3740461966386876E-4, -3.2364250248363356E-5, FastMath.toRadians(97.40236024565775), FastMath.toRadians(166.15873160992115), FastMath.toRadians(90.1282370098961), PositionAngle.MEAN, poleAligned, date, provider.getMu());
    EcksteinHechlerPropagator propagator = new EcksteinHechlerPropagator(initial, provider);
    propagator.addAdditionalStateProvider(new SevenProvider());
    propagator.setEphemerisMode();
    propagator.propagate(initial.getDate().shiftedBy(40000));
    BoundedPropagator ephemeris = propagator.getGeneratedEphemeris();
    Assert.assertSame(poleAligned, ephemeris.getFrame());
    ByteArrayOutputStream bos = new ByteArrayOutputStream();
    ObjectOutputStream oos = new ObjectOutputStream(bos);
    oos.writeObject(ephemeris);
    Assert.assertTrue(bos.size() > 2450);
    Assert.assertTrue(bos.size() < 2550);
    ByteArrayInputStream bis = new ByteArrayInputStream(bos.toByteArray());
    ObjectInputStream ois = new ObjectInputStream(bis);
    BoundedPropagator deserialized = (BoundedPropagator) ois.readObject();
    Assert.assertEquals(initial.getA(), deserialized.getInitialState().getA(), 1.0e-10);
    Assert.assertEquals(initial.getEquinoctialEx(), deserialized.getInitialState().getEquinoctialEx(), 1.0e-10);
    SpacecraftState s = deserialized.propagate(initial.getDate().shiftedBy(20000));
    Map<String, double[]> additional = s.getAdditionalStates();
    Assert.assertEquals(1, additional.size());
    Assert.assertEquals(1, additional.get("seven").length);
    Assert.assertEquals(7, additional.get("seven")[0], 1.0e-15);
}