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

Example 86 with Orbit

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

the class KeplerianPropagatorTest method testPropagationDate.

/**
 * Check that the date returned by {@link KeplerianPropagator#propagate(AbsoluteDate)}
 * is the same as the date passed to propagate().
 *
 * @throws OrekitException on error.
 */
@Test
public void testPropagationDate() throws OrekitException {
    // setup
    AbsoluteDate initDate = AbsoluteDate.J2000_EPOCH;
    // date s.t. target - date rounds down when represented as a double.
    AbsoluteDate target = initDate.shiftedBy(20.0).shiftedBy(FastMath.ulp(20.0) / 4);
    Orbit ic = new KeplerianOrbit(6378137 + 500e3, 1e-3, 0, 0, 0, 0, PositionAngle.TRUE, FramesFactory.getGCRF(), initDate, mu);
    Propagator propagator = new KeplerianPropagator(ic);
    // action
    SpacecraftState actual = propagator.propagate(target);
    // verify
    Assert.assertEquals(target, actual.getDate());
}

Example 87 with Orbit

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

the class KeplerianConverterTest method checkFit.

protected void checkFit(final Orbit orbit, final double duration, final double stepSize, final double threshold, final boolean positionOnly, final double expectedRMS, final String... freeParameters) throws OrekitException {
    Propagator p = new KeplerianPropagator(orbit);
    List<SpacecraftState> sample = new ArrayList<SpacecraftState>();
    for (double dt = 0; dt < duration; dt += stepSize) {
        sample.add(p.propagate(orbit.getDate().shiftedBy(dt)));
    }
    PropagatorBuilder builder = new KeplerianPropagatorBuilder(OrbitType.KEPLERIAN.convertType(orbit), PositionAngle.MEAN, 1.0);
    FiniteDifferencePropagatorConverter fitter = new FiniteDifferencePropagatorConverter(builder, threshold, 1000);
    fitter.convert(sample, positionOnly, freeParameters);
    Assert.assertEquals(expectedRMS, fitter.getRMS(), 0.01 * expectedRMS);
    KeplerianPropagator prop = (KeplerianPropagator) fitter.getAdaptedPropagator();
    Orbit fitted = prop.getInitialState().getOrbit();
    final double eps = 1.0e-12;
    Assert.assertEquals(orbit.getPVCoordinates().getPosition().getX(), fitted.getPVCoordinates().getPosition().getX(), eps * orbit.getPVCoordinates().getPosition().getX());
    Assert.assertEquals(orbit.getPVCoordinates().getPosition().getY(), fitted.getPVCoordinates().getPosition().getY(), eps * orbit.getPVCoordinates().getPosition().getY());
    Assert.assertEquals(orbit.getPVCoordinates().getPosition().getZ(), fitted.getPVCoordinates().getPosition().getZ(), eps * orbit.getPVCoordinates().getPosition().getZ());
    Assert.assertEquals(orbit.getPVCoordinates().getVelocity().getX(), fitted.getPVCoordinates().getVelocity().getX(), -eps * orbit.getPVCoordinates().getVelocity().getX());
    Assert.assertEquals(orbit.getPVCoordinates().getVelocity().getY(), fitted.getPVCoordinates().getVelocity().getY(), eps * orbit.getPVCoordinates().getVelocity().getY());
    Assert.assertEquals(orbit.getPVCoordinates().getVelocity().getZ(), fitted.getPVCoordinates().getVelocity().getZ(), eps * orbit.getPVCoordinates().getVelocity().getZ());
}

Example 88 with Orbit

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

the class OsculatingToMeanElementsConverterTest method testTrivial.

@Test
public void testTrivial() throws Exception {
    final AbsoluteDate date = new AbsoluteDate("2011-12-12T11:57:20.000", TimeScalesFactory.getUTC());
    final Orbit orbit1 = new CircularOrbit(7204535.848109436, -4.484755873986251E-4, 0.0011562979012178316, FastMath.toRadians(98.74341600466741), FastMath.toRadians(43.32990110790338), FastMath.toRadians(180.0), PositionAngle.MEAN, FramesFactory.getGCRF(), date, Constants.WGS84_EARTH_MU);
    final SpacecraftState initialState = new SpacecraftState(orbit1);
    // Set up the numerical propagator
    final double[][] tol = NumericalPropagator.tolerances(1.0, initialState.getOrbit(), initialState.getOrbit().getType());
    final double minStep = 1.;
    final double maxStep = 200.;
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tol[0], tol[1]);
    integrator.setInitialStepSize(100.);
    final NumericalPropagator prop = new NumericalPropagator(integrator);
    prop.setInitialState(initialState);
    final OsculatingToMeanElementsConverter converter = new OsculatingToMeanElementsConverter(initialState, 2, prop, 1.0);
    final SpacecraftState meanOrbit = converter.convert();
    final double eps = 1.e-15;
    Assert.assertEquals(orbit1.getA(), meanOrbit.getA(), eps * orbit1.getA());
    Assert.assertEquals(orbit1.getEquinoctialEx(), meanOrbit.getEquinoctialEx(), eps);
    Assert.assertEquals(orbit1.getEquinoctialEy(), meanOrbit.getEquinoctialEy(), eps);
    Assert.assertEquals(orbit1.getHx(), meanOrbit.getHx(), eps);
    Assert.assertEquals(orbit1.getHy(), meanOrbit.getHy(), eps);
    Assert.assertEquals(MathUtils.normalizeAngle(orbit1.getLM(), FastMath.PI), MathUtils.normalizeAngle(meanOrbit.getLM(), FastMath.PI), eps);
}

Example 89 with Orbit

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

the class ElevationDetectorTest method testIssue110.

@Test
public void testIssue110() throws OrekitException {
    // KEPLERIAN PROPAGATOR
    final Frame eme2000Frame = FramesFactory.getEME2000();
    final AbsoluteDate initDate = AbsoluteDate.J2000_EPOCH;
    final double a = 7000000.0;
    final Orbit initialOrbit = new KeplerianOrbit(a, 0.0, FastMath.PI / 2.2, 0.0, FastMath.PI / 2., 0.0, PositionAngle.TRUE, eme2000Frame, initDate, Constants.EGM96_EARTH_MU);
    final KeplerianPropagator kProp = new KeplerianPropagator(initialOrbit);
    // earth shape
    final OneAxisEllipsoid earthShape = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
    // Ground station
    final GeodeticPoint stat = new GeodeticPoint(FastMath.toRadians(35.), FastMath.toRadians(149.8), 0.);
    final TopocentricFrame station = new TopocentricFrame(earthShape, stat, "GSTATION");
    // detector creation
    // =================
    final double maxCheck = 600.;
    final double threshold = 1.0e-3;
    final EventDetector rawEvent = new ElevationDetector(maxCheck, threshold, station).withConstantElevation(FastMath.toRadians(5.0)).withHandler(new ContinueOnEvent<ElevationDetector>());
    final EventsLogger logger = new EventsLogger();
    kProp.addEventDetector(logger.monitorDetector(rawEvent));
    // PROPAGATION with DETECTION
    final AbsoluteDate finalDate = initDate.shiftedBy(30 * 60.);
    kProp.propagate(finalDate);
    Assert.assertEquals(2, logger.getLoggedEvents().size());
    Assert.assertTrue(logger.getLoggedEvents().get(0).isIncreasing());
    Assert.assertEquals(478.945, logger.getLoggedEvents().get(0).getState().getDate().durationFrom(initDate), 1.0e-3);
    Assert.assertFalse(logger.getLoggedEvents().get(1).isIncreasing());
    Assert.assertEquals(665.721, logger.getLoggedEvents().get(1).getState().getDate().durationFrom(initDate), 1.0e-3);
}

Example 90 with Orbit

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

the class ElevationDetectorTest method testIssue136.

@Test
public void testIssue136() throws OrekitException {
    // Initial state definition : date, orbit
    AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, TimeScalesFactory.getUTC());
    // inertial frame for orbit definition
    Frame inertialFrame = FramesFactory.getEME2000();
    Orbit initialOrbit = new KeplerianOrbit(6828137.005, 7.322641382145889e-10, 1.6967079057368113, 0.0, 1.658054062748353, 0.0001223149429077902, PositionAngle.MEAN, inertialFrame, initialDate, Constants.EIGEN5C_EARTH_MU);
    // Propagator : consider a simple Keplerian motion (could be more elaborate)
    Propagator kepler = new EcksteinHechlerPropagator(initialOrbit, Constants.EGM96_EARTH_EQUATORIAL_RADIUS, Constants.EGM96_EARTH_MU, Constants.EGM96_EARTH_C20, 0.0, 0.0, 0.0, 0.0);
    // Earth and frame
    // equatorial radius in meter
    double ae = 6378137.0;
    // flattening
    double f = 1.0 / 298.257223563;
    // terrestrial frame at an arbitrary date
    Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
    BodyShape earth = new OneAxisEllipsoid(ae, f, itrf);
    // Station
    final double longitude = FastMath.toRadians(-147.5);
    final double latitude = FastMath.toRadians(64);
    final double altitude = 160;
    final GeodeticPoint station1 = new GeodeticPoint(latitude, longitude, altitude);
    final TopocentricFrame sta1Frame = new TopocentricFrame(earth, station1, "station1");
    // Event definition
    final double maxcheck = 120.0;
    final double elevation = FastMath.toRadians(5.);
    final double threshold = 10.0;
    final EventDetector rawEvent = new ElevationDetector(maxcheck, threshold, sta1Frame).withConstantElevation(elevation).withHandler(new ContinueOnEvent<ElevationDetector>());
    final EventsLogger logger = new EventsLogger();
    kepler.addEventDetector(logger.monitorDetector(rawEvent));
    // Propagate from the initial date to the first raising or for the fixed duration
    kepler.propagate(initialDate.shiftedBy(60 * 60 * 24.0 * 40));
    int countIncreasing = 0;
    int countDecreasing = 0;
    for (LoggedEvent le : logger.getLoggedEvents()) {
        if (le.isIncreasing()) {
            ++countIncreasing;
        } else {
            ++countDecreasing;
        }
    }
    Assert.assertEquals(314, countIncreasing);
    Assert.assertEquals(314, countDecreasing);
}