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

Example 1 with CircularOrbit

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

the class HarmonicParametricAccelerationTest method testCoefficientsDetermination.

@Test
public void testCoefficientsDetermination() throws OrekitException {
    final double mass = 2500;
    final Orbit orbit = new CircularOrbit(7500000.0, 1.0e-4, 1.0e-3, 1.7, 0.3, 0.5, PositionAngle.TRUE, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 2, 3), TimeComponents.H00, TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
    final double period = orbit.getKeplerianPeriod();
    AttitudeProvider maneuverLaw = new LofOffset(orbit.getFrame(), LOFType.VNC);
    SpacecraftState initialState = new SpacecraftState(orbit, maneuverLaw.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass);
    double dP = 10.0;
    double minStep = 0.001;
    double maxStep = 100;
    double[][] tolerance = NumericalPropagator.tolerances(dP, orbit, orbit.getType());
    // generate PV measurements corresponding to a tangential maneuver
    AdaptiveStepsizeIntegrator integrator0 = new DormandPrince853Integrator(minStep, maxStep, tolerance[0], tolerance[1]);
    integrator0.setInitialStepSize(60);
    final NumericalPropagator propagator0 = new NumericalPropagator(integrator0);
    propagator0.setInitialState(initialState);
    propagator0.setAttitudeProvider(maneuverLaw);
    ForceModel hpaRefX1 = new HarmonicParametricAcceleration(Vector3D.PLUS_I, true, "refX1", null, period, 1);
    ForceModel hpaRefY1 = new HarmonicParametricAcceleration(Vector3D.PLUS_J, true, "refY1", null, period, 1);
    ForceModel hpaRefZ2 = new HarmonicParametricAcceleration(Vector3D.PLUS_K, true, "refZ2", null, period, 2);
    hpaRefX1.getParametersDrivers()[0].setValue(2.4e-2);
    hpaRefX1.getParametersDrivers()[1].setValue(3.1);
    hpaRefY1.getParametersDrivers()[0].setValue(4.0e-2);
    hpaRefY1.getParametersDrivers()[1].setValue(0.3);
    hpaRefZ2.getParametersDrivers()[0].setValue(1.0e-2);
    hpaRefZ2.getParametersDrivers()[1].setValue(1.8);
    propagator0.addForceModel(hpaRefX1);
    propagator0.addForceModel(hpaRefY1);
    propagator0.addForceModel(hpaRefZ2);
    final List<ObservedMeasurement<?>> measurements = new ArrayList<>();
    propagator0.setMasterMode(10.0, (state, isLast) -> measurements.add(new PV(state.getDate(), state.getPVCoordinates().getPosition(), state.getPVCoordinates().getVelocity(), 1.0e-3, 1.0e-6, 1.0)));
    propagator0.propagate(orbit.getDate().shiftedBy(900));
    // set up an estimator to retrieve the maneuver as several harmonic accelerations in inertial frame
    final NumericalPropagatorBuilder propagatorBuilder = new NumericalPropagatorBuilder(orbit, new DormandPrince853IntegratorBuilder(minStep, maxStep, dP), PositionAngle.TRUE, dP);
    propagatorBuilder.addForceModel(new HarmonicParametricAcceleration(Vector3D.PLUS_I, true, "X1", null, period, 1));
    propagatorBuilder.addForceModel(new HarmonicParametricAcceleration(Vector3D.PLUS_J, true, "Y1", null, period, 1));
    propagatorBuilder.addForceModel(new HarmonicParametricAcceleration(Vector3D.PLUS_K, true, "Z2", null, period, 2));
    final BatchLSEstimator estimator = new BatchLSEstimator(new LevenbergMarquardtOptimizer(), propagatorBuilder);
    estimator.setParametersConvergenceThreshold(1.0e-2);
    estimator.setMaxIterations(20);
    estimator.setMaxEvaluations(100);
    for (final ObservedMeasurement<?> measurement : measurements) {
        estimator.addMeasurement(measurement);
    }
    // we will estimate only the force model parameters, not the orbit
    for (final ParameterDriver d : estimator.getOrbitalParametersDrivers(false).getDrivers()) {
        d.setSelected(false);
    }
    setParameter(estimator, "X1 γ", 1.0e-2);
    setParameter(estimator, "X1 φ", 4.0);
    setParameter(estimator, "Y1 γ", 1.0e-2);
    setParameter(estimator, "Y1 φ", 0.0);
    setParameter(estimator, "Z2 γ", 1.0e-2);
    setParameter(estimator, "Z2 φ", 1.0);
    estimator.estimate();
    Assert.assertTrue(estimator.getIterationsCount() < 15);
    Assert.assertTrue(estimator.getEvaluationsCount() < 15);
    Assert.assertEquals(0.0, estimator.getOptimum().getRMS(), 1.0e-5);
    Assert.assertEquals(hpaRefX1.getParametersDrivers()[0].getValue(), getParameter(estimator, "X1 γ"), 1.e-12);
    Assert.assertEquals(hpaRefX1.getParametersDrivers()[1].getValue(), getParameter(estimator, "X1 φ"), 1.e-12);
    Assert.assertEquals(hpaRefY1.getParametersDrivers()[0].getValue(), getParameter(estimator, "Y1 γ"), 1.e-12);
    Assert.assertEquals(hpaRefY1.getParametersDrivers()[1].getValue(), getParameter(estimator, "Y1 φ"), 1.e-12);
    Assert.assertEquals(hpaRefZ2.getParametersDrivers()[0].getValue(), getParameter(estimator, "Z2 γ"), 1.e-12);
    Assert.assertEquals(hpaRefZ2.getParametersDrivers()[1].getValue(), getParameter(estimator, "Z2 φ"), 1.e-12);
}

Example 2 with CircularOrbit

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

the class OrbitDeterminationTest method createOrbit.

/**
 * Create an orbit from input parameters
 * @param parser input file parser
 * @param mu     central attraction coefficient
 * @throws NoSuchElementException if input parameters are missing
 * @throws OrekitException if inertial frame cannot be created
 */
private Orbit createOrbit(final KeyValueFileParser<ParameterKey> parser, final double mu) throws NoSuchElementException, OrekitException {
    final Frame frame;
    if (!parser.containsKey(ParameterKey.INERTIAL_FRAME)) {
        frame = FramesFactory.getEME2000();
    } else {
        frame = parser.getInertialFrame(ParameterKey.INERTIAL_FRAME);
    }
    // Orbit definition
    PositionAngle angleType = PositionAngle.MEAN;
    if (parser.containsKey(ParameterKey.ORBIT_ANGLE_TYPE)) {
        angleType = PositionAngle.valueOf(parser.getString(ParameterKey.ORBIT_ANGLE_TYPE).toUpperCase());
    }
    if (parser.containsKey(ParameterKey.ORBIT_KEPLERIAN_A)) {
        return new KeplerianOrbit(parser.getDouble(ParameterKey.ORBIT_KEPLERIAN_A), parser.getDouble(ParameterKey.ORBIT_KEPLERIAN_E), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_I), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_PA), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_RAAN), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_ANOMALY), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_EQUINOCTIAL_A)) {
        return new EquinoctialOrbit(parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_A), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_EX), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_EY), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_HX), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_HY), parser.getAngle(ParameterKey.ORBIT_EQUINOCTIAL_LAMBDA), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_CIRCULAR_A)) {
        return new CircularOrbit(parser.getDouble(ParameterKey.ORBIT_CIRCULAR_A), parser.getDouble(ParameterKey.ORBIT_CIRCULAR_EX), parser.getDouble(ParameterKey.ORBIT_CIRCULAR_EY), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_I), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_RAAN), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_ALPHA), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_TLE_LINE_1)) {
        final String line1 = parser.getString(ParameterKey.ORBIT_TLE_LINE_1);
        final String line2 = parser.getString(ParameterKey.ORBIT_TLE_LINE_2);
        final TLE tle = new TLE(line1, line2);
        TLEPropagator propagator = TLEPropagator.selectExtrapolator(tle);
        // propagator.setEphemerisMode();
        AbsoluteDate initDate = tle.getDate();
        SpacecraftState initialState = propagator.getInitialState();
        // Transformation from TEME to frame.
        Transform t = FramesFactory.getTEME().getTransformTo(FramesFactory.getEME2000(), initDate.getDate());
        return new CartesianOrbit(t.transformPVCoordinates(initialState.getPVCoordinates()), frame, initDate, mu);
    } else {
        final double[] pos = { parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PX), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PY), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PZ) };
        final double[] vel = { parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VX), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VY), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VZ) };
        return new CartesianOrbit(new PVCoordinates(new Vector3D(pos), new Vector3D(vel)), frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    }
}

Example 3 with CircularOrbit

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

the class KalmanOrbitDeterminationTest method createOrbit.

/**
 * Create an orbit from input parameters
 * @param parser input file parser
 * @param mu     central attraction coefficient
 * @throws NoSuchElementException if input parameters are missing
 * @throws OrekitException if inertial frame cannot be created
 */
private Orbit createOrbit(final KeyValueFileParser<ParameterKey> parser, final double mu) throws NoSuchElementException, OrekitException {
    final Frame frame;
    if (!parser.containsKey(ParameterKey.INERTIAL_FRAME)) {
        frame = FramesFactory.getEME2000();
    } else {
        frame = parser.getInertialFrame(ParameterKey.INERTIAL_FRAME);
    }
    // Orbit definition
    PositionAngle angleType = PositionAngle.MEAN;
    if (parser.containsKey(ParameterKey.ORBIT_ANGLE_TYPE)) {
        angleType = PositionAngle.valueOf(parser.getString(ParameterKey.ORBIT_ANGLE_TYPE).toUpperCase());
    }
    if (parser.containsKey(ParameterKey.ORBIT_KEPLERIAN_A)) {
        return new KeplerianOrbit(parser.getDouble(ParameterKey.ORBIT_KEPLERIAN_A), parser.getDouble(ParameterKey.ORBIT_KEPLERIAN_E), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_I), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_PA), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_RAAN), parser.getAngle(ParameterKey.ORBIT_KEPLERIAN_ANOMALY), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_EQUINOCTIAL_A)) {
        return new EquinoctialOrbit(parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_A), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_EX), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_EY), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_HX), parser.getDouble(ParameterKey.ORBIT_EQUINOCTIAL_HY), parser.getAngle(ParameterKey.ORBIT_EQUINOCTIAL_LAMBDA), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_CIRCULAR_A)) {
        return new CircularOrbit(parser.getDouble(ParameterKey.ORBIT_CIRCULAR_A), parser.getDouble(ParameterKey.ORBIT_CIRCULAR_EX), parser.getDouble(ParameterKey.ORBIT_CIRCULAR_EY), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_I), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_RAAN), parser.getAngle(ParameterKey.ORBIT_CIRCULAR_ALPHA), angleType, frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    } else if (parser.containsKey(ParameterKey.ORBIT_TLE_LINE_1)) {
        final String line1 = parser.getString(ParameterKey.ORBIT_TLE_LINE_1);
        final String line2 = parser.getString(ParameterKey.ORBIT_TLE_LINE_2);
        final TLE tle = new TLE(line1, line2);
        TLEPropagator propagator = TLEPropagator.selectExtrapolator(tle);
        // propagator.setEphemerisMode();
        AbsoluteDate initDate = tle.getDate();
        SpacecraftState initialState = propagator.getInitialState();
        // Transformation from TEME to frame.
        Transform t = FramesFactory.getTEME().getTransformTo(FramesFactory.getEME2000(), initDate.getDate());
        return new CartesianOrbit(t.transformPVCoordinates(initialState.getPVCoordinates()), frame, initDate, mu);
    } else {
        final double[] pos = { parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PX), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PY), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_PZ) };
        final double[] vel = { parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VX), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VY), parser.getDouble(ParameterKey.ORBIT_CARTESIAN_VZ) };
        return new CartesianOrbit(new PVCoordinates(new Vector3D(pos), new Vector3D(vel)), frame, parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), mu);
    }
}

Example 4 with CircularOrbit

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

the class BoxAndSolarArraySpacecraftTest method setUp.

@Before
public void setUp() {
    try {
        Utils.setDataRoot("regular-data");
        mu = 3.9860047e14;
        double ae = 6.378137e6;
        double c20 = -1.08263e-3;
        double c30 = 2.54e-6;
        double c40 = 1.62e-6;
        double c50 = 2.3e-7;
        double c60 = -5.5e-7;
        AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 7, 1), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
        // Satellite position as circular parameters, raan chosen to have sun elevation with
        // respect to orbit plane roughly evolving roughly from 15 to 15.2 degrees in the test range
        Orbit circ = new CircularOrbit(7178000.0, 0.5e-4, -0.5e-4, FastMath.toRadians(50.), FastMath.toRadians(280), FastMath.toRadians(10.0), PositionAngle.MEAN, FramesFactory.getEME2000(), date, mu);
        propagator = new EcksteinHechlerPropagator(circ, new LofOffset(circ.getFrame(), LOFType.VVLH), ae, mu, c20, c30, c40, c50, c60);
    } catch (OrekitException oe) {
        Assert.fail(oe.getLocalizedMessage());
    }
}

Example 5 with CircularOrbit

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

the class DSSTPropagatorTest method testIssue364.

@Test
public void testIssue364() throws OrekitException {
    Utils.setDataRoot("regular-data");
    AbsoluteDate date = new AbsoluteDate("2003-06-18T00:00:00.000", TimeScalesFactory.getUTC());
    CircularOrbit orbit = new CircularOrbit(7389068.5, 0.0, 0.0, 1.709573, 1.308398, 0, PositionAngle.MEAN, FramesFactory.getTOD(IERSConventions.IERS_2010, false), date, Constants.WGS84_EARTH_MU);
    SpacecraftState osculatingState = new SpacecraftState(orbit, 1116.2829);
    List<DSSTForceModel> dsstForceModels = new ArrayList<DSSTForceModel>();
    dsstForceModels.add(new DSSTThirdBody(CelestialBodyFactory.getMoon()));
    dsstForceModels.add(new DSSTThirdBody(CelestialBodyFactory.getSun()));
    SpacecraftState meanState = DSSTPropagator.computeMeanState(osculatingState, null, dsstForceModels);
    Assert.assertEquals(0.421, osculatingState.getA() - meanState.getA(), 1.0e-3);
    Assert.assertEquals(-5.23e-8, osculatingState.getEquinoctialEx() - meanState.getEquinoctialEx(), 1.0e-10);
    Assert.assertEquals(15.22e-8, osculatingState.getEquinoctialEy() - meanState.getEquinoctialEy(), 1.0e-10);
    Assert.assertEquals(-3.15e-8, osculatingState.getHx() - meanState.getHx(), 1.0e-10);
    Assert.assertEquals(2.83e-8, osculatingState.getHy() - meanState.getHy(), 1.0e-10);
    Assert.assertEquals(15.96e-8, osculatingState.getLM() - meanState.getLM(), 1.0e-10);
}