Example usage for org.apache.commons.math3.geometry.euclidean.threed RotationOrder XYX

Introduction

In this page you can find the example usage for org.apache.commons.math3.geometry.euclidean.threed RotationOrder XYX.

Prototype

RotationOrder XYX

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Document

Set of Euler angles.

Usage

From source file:org.orekit.attitudes.LofOffsetPointingTest.java

@Test
public void testMiss() throws OrekitException {
    final CircularOrbit circ = new CircularOrbit(7178000.0, 0.5e-4, -0.5e-4, FastMath.toRadians(0.),
            FastMath.toRadians(270.), FastMath.toRadians(5.300), PositionAngle.MEAN, FramesFactory.getEME2000(),
            date, mu);/* w ww.  j a v a 2 s  .c  om*/
    final LofOffset upsideDown = new LofOffset(circ.getFrame(), LOFType.VVLH, RotationOrder.XYX, FastMath.PI, 0,
            0);
    final LofOffsetPointing pointing = new LofOffsetPointing(circ.getFrame(), earthSpheric, upsideDown,
            Vector3D.PLUS_K);
    try {
        pointing.getTargetPV(circ, date, circ.getFrame());
        Assert.fail("an exception should have been thrown");
    } catch (OrekitException oe) {
        Assert.assertEquals(OrekitMessages.ATTITUDE_POINTING_LAW_DOES_NOT_POINT_TO_GROUND, oe.getSpecifier());
    }
}

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From source file:org.orekit.attitudes.LofOffsetPointingTest.java

@Test
public void testSpin() throws OrekitException {

    AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 01, 01), new TimeComponents(3, 25, 45.6789),
            TimeScalesFactory.getUTC());
    KeplerianOrbit orbit = new KeplerianOrbit(7178000.0, 1.e-4, FastMath.toRadians(50.),
            FastMath.toRadians(10.), FastMath.toRadians(20.), FastMath.toRadians(30.), PositionAngle.MEAN,
            FramesFactory.getEME2000(), date, 3.986004415e14);

    final AttitudeProvider law = new LofOffsetPointing(orbit.getFrame(), earthSpheric,
            new LofOffset(orbit.getFrame(), LOFType.VVLH, RotationOrder.XYX, 0.1, 0.2, 0.3), Vector3D.PLUS_K);

    Propagator propagator = new KeplerianPropagator(orbit, law);

    double h = 0.01;
    SpacecraftState sMinus = propagator.propagate(date.shiftedBy(-h));
    SpacecraftState s0 = propagator.propagate(date);
    SpacecraftState sPlus = propagator.propagate(date.shiftedBy(h));

    // check spin is consistent with attitude evolution
    double errorAngleMinus = Rotation.distance(sMinus.shiftedBy(h).getAttitude().getRotation(),
            s0.getAttitude().getRotation());
    double evolutionAngleMinus = Rotation.distance(sMinus.getAttitude().getRotation(),
            s0.getAttitude().getRotation());
    Assert.assertEquals(0.0, errorAngleMinus, 1.0e-6 * evolutionAngleMinus);
    double errorAnglePlus = Rotation.distance(s0.getAttitude().getRotation(),
            sPlus.shiftedBy(-h).getAttitude().getRotation());
    double evolutionAnglePlus = Rotation.distance(s0.getAttitude().getRotation(),
            sPlus.getAttitude().getRotation());
    Assert.assertEquals(0.0, errorAnglePlus, 1.0e-6 * evolutionAnglePlus);

    Vector3D spin0 = s0.getAttitude().getSpin();
    Vector3D reference = AngularCoordinates.estimateRate(sMinus.getAttitude().getRotation(),
            sPlus.getAttitude().getRotation(), 2 * h);
    Assert.assertTrue(spin0.getNorm() > 1.0e-3);
    Assert.assertEquals(0.0, spin0.subtract(reference).getNorm(), 1.0e-10);

}

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From source file:org.orekit.attitudes.LofOffsetTest.java

@Test
public void testSpin() throws OrekitException {

    final AttitudeProvider law = new LofOffset(orbit.getFrame(), LOFType.VVLH, RotationOrder.XYX, 0.1, 0.2,
            0.3);//from  w  w w . j  av  a2 s  .  c o  m

    AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 01, 01), new TimeComponents(3, 25, 45.6789),
            TimeScalesFactory.getUTC());
    KeplerianOrbit orbit = new KeplerianOrbit(7178000.0, 1.e-4, FastMath.toRadians(50.),
            FastMath.toRadians(10.), FastMath.toRadians(20.), FastMath.toRadians(30.), PositionAngle.MEAN,
            FramesFactory.getEME2000(), date, 3.986004415e14);

    Propagator propagator = new KeplerianPropagator(orbit, law);

    double h = 0.01;
    SpacecraftState sMinus = propagator.propagate(date.shiftedBy(-h));
    SpacecraftState s0 = propagator.propagate(date);
    SpacecraftState sPlus = propagator.propagate(date.shiftedBy(h));

    // check spin is consistent with attitude evolution
    double errorAngleMinus = Rotation.distance(sMinus.shiftedBy(h).getAttitude().getRotation(),
            s0.getAttitude().getRotation());
    double evolutionAngleMinus = Rotation.distance(sMinus.getAttitude().getRotation(),
            s0.getAttitude().getRotation());
    Assert.assertEquals(0.0, errorAngleMinus, 1.0e-6 * evolutionAngleMinus);
    double errorAnglePlus = Rotation.distance(s0.getAttitude().getRotation(),
            sPlus.shiftedBy(-h).getAttitude().getRotation());
    double evolutionAnglePlus = Rotation.distance(s0.getAttitude().getRotation(),
            sPlus.getAttitude().getRotation());
    Assert.assertEquals(0.0, errorAnglePlus, 1.0e-6 * evolutionAnglePlus);

    Vector3D spin0 = s0.getAttitude().getSpin();
    Vector3D reference = AngularCoordinates.estimateRate(sMinus.getAttitude().getRotation(),
            sPlus.getAttitude().getRotation(), 2 * h);
    Assert.assertEquals(0.0, spin0.subtract(reference).getNorm(), 1.0e-10);

}

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From source file:org.orekit.forces.maneuvers.ImpulseManeuverTest.java

@Test
public void testInertialManeuver() 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 state = new TimeStampedPVCoordinates(epoch, position, velocity,
            Vector3D.ZERO);/*w w w .  j  av  a 2 s . c o  m*/
    final Orbit initialOrbit = new CartesianOrbit(state, FramesFactory.getEME2000(), mu);

    final double totalPropagationTime = 0.00001;
    final double driftTimeInSec = totalPropagationTime / 2.0;
    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);

    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit,
            new LofOffset(initialOrbit.getFrame(), LOFType.VNC));
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(driftTimeInSec));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector,
            attitudeOverride, deltaV, isp).withThreshold(driftTimeInSec / 4);
    propagator.addEventDetector(burnAtEpoch);

    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));

    final double finalVxExpected = initialVx + deltaX;
    final double finalVyExpected = initialVy + deltaY;
    final double finalVzExpected = initialVz + deltaZ;
    final double maneuverTolerance = 1e-4;

    final Vector3D finalVelocity = finalState.getPVCoordinates().getVelocity();
    Assert.assertEquals(finalVxExpected, finalVelocity.getX(), maneuverTolerance);
    Assert.assertEquals(finalVyExpected, finalVelocity.getY(), maneuverTolerance);
    Assert.assertEquals(finalVzExpected, finalVelocity.getZ(), maneuverTolerance);

}

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