Examples with TimeComponents org.orekit.time.TimeComponents used on opensource projects

Example 76 with TimeComponents

use of org.orekit.time.TimeComponents in project Orekit by CS-SI.

the class FixedRateTest method testZeroRate.

@Test
public void testZeroRate() throws OrekitException {
    AbsoluteDate date = new AbsoluteDate(new DateComponents(2004, 3, 2), new TimeComponents(13, 17, 7.865), TimeScalesFactory.getUTC());
    final Frame frame = FramesFactory.getEME2000();
    FixedRate law = new FixedRate(new Attitude(date, frame, new Rotation(0.48, 0.64, 0.36, 0.48, false), Vector3D.ZERO, Vector3D.ZERO));
    PVCoordinates pv = new PVCoordinates(new Vector3D(28812595.32012577, 5948437.4640250085, 0), new Vector3D(0, 0, 3680.853673522056));
    Orbit orbit = new KeplerianOrbit(pv, frame, date, 3.986004415e14);
    Rotation attitude0 = law.getAttitude(orbit, date, frame).getRotation();
    Assert.assertEquals(0, Rotation.distance(attitude0, law.getReferenceAttitude().getRotation()), 1.0e-10);
    Rotation attitude1 = law.getAttitude(orbit.shiftedBy(10.0), date.shiftedBy(10.0), frame).getRotation();
    Assert.assertEquals(0, Rotation.distance(attitude1, law.getReferenceAttitude().getRotation()), 1.0e-10);
    Rotation attitude2 = law.getAttitude(orbit.shiftedBy(20.0), date.shiftedBy(20.0), frame).getRotation();
    Assert.assertEquals(0, Rotation.distance(attitude2, law.getReferenceAttitude().getRotation()), 1.0e-10);
}

Example 77 with TimeComponents

use of org.orekit.time.TimeComponents in project Orekit by CS-SI.

the class FixedRateTest method testSpin.

@Test
public void testSpin() throws OrekitException {
    AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 01, 01), new TimeComponents(3, 25, 45.6789), TimeScalesFactory.getUTC());
    final double rate = 2 * FastMath.PI / (12 * 60);
    AttitudeProvider law = new FixedRate(new Attitude(date, FramesFactory.getEME2000(), new Rotation(0.48, 0.64, 0.36, 0.48, false), new Vector3D(rate, Vector3D.PLUS_K), Vector3D.ZERO));
    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-14);
}

Example 78 with TimeComponents

use of org.orekit.time.TimeComponents in project Orekit by CS-SI.

the class AdapterPropagatorTest method testLowEarthOrbit.

@Test
public void testLowEarthOrbit() throws OrekitException, ParseException, IOException {
    Orbit leo = new CircularOrbit(7200000.0, -1.0e-5, 2.0e-4, FastMath.toRadians(98.0), FastMath.toRadians(123.456), 0.0, PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
    double mass = 5600.0;
    AbsoluteDate t0 = leo.getDate().shiftedBy(1000.0);
    Vector3D dV = new Vector3D(-0.1, 0.2, 0.3);
    double f = 20.0;
    double isp = 315.0;
    double vExhaust = Constants.G0_STANDARD_GRAVITY * isp;
    double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust);
    BoundedPropagator withoutManeuver = getEphemeris(leo, mass, 5, new LofOffset(leo.getFrame(), LOFType.LVLH), t0, Vector3D.ZERO, f, isp, false, false, null);
    BoundedPropagator withManeuver = getEphemeris(leo, mass, 5, new LofOffset(leo.getFrame(), LOFType.LVLH), t0, dV, f, isp, false, false, null);
    // we set up a model that reverts the maneuvers
    AdapterPropagator adapterPropagator = new AdapterPropagator(withManeuver);
    AdapterPropagator.DifferentialEffect effect = new SmallManeuverAnalyticalModel(adapterPropagator.propagate(t0), dV.negate(), isp);
    adapterPropagator.addEffect(effect);
    adapterPropagator.addAdditionalStateProvider(new AdditionalStateProvider() {

        public String getName() {
            return "dummy 3";
        }

        public double[] getAdditionalState(SpacecraftState state) {
            return new double[3];
        }
    });
    // the adapted propagators do not manage the additional states from the reference,
    // they simply forward them
    Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 1"));
    Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 2"));
    Assert.assertTrue(adapterPropagator.isAdditionalStateManaged("dummy 3"));
    for (AbsoluteDate t = t0.shiftedBy(0.5 * dt); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t.shiftedBy(60.0)) {
        PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, leo.getFrame());
        PVCoordinates pvReverted = adapterPropagator.getPVCoordinates(t, leo.getFrame());
        double revertError = new PVCoordinates(pvWithout, pvReverted).getPosition().getNorm();
        Assert.assertEquals(0, revertError, 0.45);
        Assert.assertEquals(2, adapterPropagator.propagate(t).getAdditionalState("dummy 1").length);
        Assert.assertEquals(1, adapterPropagator.propagate(t).getAdditionalState("dummy 2").length);
        Assert.assertEquals(3, adapterPropagator.propagate(t).getAdditionalState("dummy 3").length);
    }
}

Example 79 with TimeComponents

use of org.orekit.time.TimeComponents in project Orekit by CS-SI.

the class AdapterPropagatorTest method testNonKeplerian.

@Test
public void testNonKeplerian() throws OrekitException, ParseException, IOException {
    Orbit leo = new CircularOrbit(7204319.233600575, 4.434564637450575E-4, 0.0011736728299091088, 1.7211611441767323, 5.5552084166959474, 24950.321259193086, PositionAngle.TRUE, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2003, 9, 16), new TimeComponents(23, 11, 20.264), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
    double mass = 4093.0;
    AbsoluteDate t0 = new AbsoluteDate(new DateComponents(2003, 9, 16), new TimeComponents(23, 14, 40.264), TimeScalesFactory.getUTC());
    Vector3D dV = new Vector3D(0.0, 3.0, 0.0);
    double f = 40.0;
    double isp = 300.0;
    double vExhaust = Constants.G0_STANDARD_GRAVITY * isp;
    double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust);
    // setup a specific coefficient file for gravity potential as it will also
    // try to read a corrupted one otherwise
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("g007_eigen_05c_coef", false));
    NormalizedSphericalHarmonicsProvider gravityField = GravityFieldFactory.getNormalizedProvider(8, 8);
    BoundedPropagator withoutManeuver = getEphemeris(leo, mass, 10, new LofOffset(leo.getFrame(), LOFType.VNC), t0, Vector3D.ZERO, f, isp, true, true, gravityField);
    BoundedPropagator withManeuver = getEphemeris(leo, mass, 10, new LofOffset(leo.getFrame(), LOFType.VNC), t0, dV, f, isp, true, true, gravityField);
    // we set up a model that reverts the maneuvers
    AdapterPropagator adapterPropagator = new AdapterPropagator(withManeuver);
    SpacecraftState state0 = adapterPropagator.propagate(t0);
    AdapterPropagator.DifferentialEffect directEffect = new SmallManeuverAnalyticalModel(state0, dV.negate(), isp);
    AdapterPropagator.DifferentialEffect derivedEffect = new J2DifferentialEffect(state0, directEffect, false, GravityFieldFactory.getUnnormalizedProvider(gravityField));
    adapterPropagator.addEffect(directEffect);
    adapterPropagator.addEffect(derivedEffect);
    adapterPropagator.addAdditionalStateProvider(new AdditionalStateProvider() {

        public String getName() {
            return "dummy 3";
        }

        public double[] getAdditionalState(SpacecraftState state) {
            return new double[3];
        }
    });
    // the adapted propagators do not manage the additional states from the reference,
    // they simply forward them
    Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 1"));
    Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 2"));
    Assert.assertTrue(adapterPropagator.isAdditionalStateManaged("dummy 3"));
    double maxDelta = 0;
    double maxNominal = 0;
    for (AbsoluteDate t = t0.shiftedBy(0.5 * dt); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t.shiftedBy(600.0)) {
        PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, leo.getFrame());
        PVCoordinates pvWith = withManeuver.getPVCoordinates(t, leo.getFrame());
        PVCoordinates pvReverted = adapterPropagator.getPVCoordinates(t, leo.getFrame());
        double nominal = new PVCoordinates(pvWithout, pvWith).getPosition().getNorm();
        double revertError = new PVCoordinates(pvWithout, pvReverted).getPosition().getNorm();
        maxDelta = FastMath.max(maxDelta, revertError);
        maxNominal = FastMath.max(maxNominal, nominal);
        Assert.assertEquals(2, adapterPropagator.propagate(t).getAdditionalState("dummy 1").length);
        Assert.assertEquals(1, adapterPropagator.propagate(t).getAdditionalState("dummy 2").length);
        Assert.assertEquals(3, adapterPropagator.propagate(t).getAdditionalState("dummy 3").length);
    }
    Assert.assertTrue(maxDelta < 120);
    Assert.assertTrue(maxNominal > 2800);
}

Example 80 with TimeComponents

use of org.orekit.time.TimeComponents in project Orekit by CS-SI.

the class AdapterPropagatorTest method testEccentricOrbit.

@Test
public void testEccentricOrbit() throws OrekitException, ParseException, IOException {
    Orbit heo = new KeplerianOrbit(90000000.0, 0.92, FastMath.toRadians(98.0), FastMath.toRadians(12.3456), FastMath.toRadians(123.456), FastMath.toRadians(1.23456), PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
    double mass = 5600.0;
    AbsoluteDate t0 = heo.getDate().shiftedBy(1000.0);
    Vector3D dV = new Vector3D(-0.01, 0.02, 0.03);
    double f = 20.0;
    double isp = 315.0;
    double vExhaust = Constants.G0_STANDARD_GRAVITY * isp;
    double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust);
    BoundedPropagator withoutManeuver = getEphemeris(heo, mass, 5, new LofOffset(heo.getFrame(), LOFType.LVLH), t0, Vector3D.ZERO, f, isp, false, false, null);
    BoundedPropagator withManeuver = getEphemeris(heo, mass, 5, new LofOffset(heo.getFrame(), LOFType.LVLH), t0, dV, f, isp, false, false, null);
    // we set up a model that reverts the maneuvers
    AdapterPropagator adapterPropagator = new AdapterPropagator(withManeuver);
    AdapterPropagator.DifferentialEffect effect = new SmallManeuverAnalyticalModel(adapterPropagator.propagate(t0), dV.negate(), isp);
    adapterPropagator.addEffect(effect);
    adapterPropagator.addAdditionalStateProvider(new AdditionalStateProvider() {

        public String getName() {
            return "dummy 3";
        }

        public double[] getAdditionalState(SpacecraftState state) {
            return new double[3];
        }
    });
    // the adapted propagators do not manage the additional states from the reference,
    // they simply forward them
    Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 1"));
    Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 2"));
    Assert.assertTrue(adapterPropagator.isAdditionalStateManaged("dummy 3"));
    for (AbsoluteDate t = t0.shiftedBy(0.5 * dt); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t.shiftedBy(300.0)) {
        PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, heo.getFrame());
        PVCoordinates pvReverted = adapterPropagator.getPVCoordinates(t, heo.getFrame());
        double revertError = Vector3D.distance(pvWithout.getPosition(), pvReverted.getPosition());
        Assert.assertEquals(0, revertError, 2.5e-5 * heo.getA());
        Assert.assertEquals(2, adapterPropagator.propagate(t).getAdditionalState("dummy 1").length);
        Assert.assertEquals(1, adapterPropagator.propagate(t).getAdditionalState("dummy 2").length);
        Assert.assertEquals(3, adapterPropagator.propagate(t).getAdditionalState("dummy 3").length);
    }
}