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

Example 36 with EquinoctialOrbit

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

the class ElevationDetectorTest method testAgata.

@Test
public void testAgata() throws OrekitException {
    final TimeScale utc = TimeScalesFactory.getUTC();
    final Vector3D position = new Vector3D(-6142438.668, 3492467.56, -25767.257);
    final Vector3D velocity = new Vector3D(505.848, 942.781, 7435.922);
    final AbsoluteDate date = new AbsoluteDate(2003, 9, 16, utc);
    final Orbit orbit = new EquinoctialOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), date, mu);
    Propagator propagator = new EcksteinHechlerPropagator(orbit, ae, mu, c20, c30, c40, c50, c60);
    // 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);
    GeodeticPoint point = new GeodeticPoint(FastMath.toRadians(48.833), FastMath.toRadians(2.333), 0.0);
    TopocentricFrame topo = new TopocentricFrame(earth, point, "Gstation");
    Checking checking = new Checking(topo);
    ElevationDetector detector = new ElevationDetector(topo).withConstantElevation(FastMath.toRadians(5.0)).withHandler(checking);
    Assert.assertNull(detector.getElevationMask());
    Assert.assertNull(detector.getRefractionModel());
    Assert.assertSame(topo, detector.getTopocentricFrame());
    Assert.assertEquals(FastMath.toRadians(5.0), detector.getMinElevation(), 1.0e-15);
    AbsoluteDate startDate = new AbsoluteDate(2003, 9, 15, 12, 0, 0, utc);
    propagator.resetInitialState(propagator.propagate(startDate));
    propagator.addEventDetector(detector);
    propagator.setMasterMode(10.0, checking);
    propagator.propagate(startDate.shiftedBy(Constants.JULIAN_DAY));
}

Example 37 with EquinoctialOrbit

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

the class ElevationDetectorTest method testHorizon.

@Test
public void testHorizon() throws OrekitException {
    final TimeScale utc = TimeScalesFactory.getUTC();
    final Vector3D position = new Vector3D(-6142438.668, 3492467.56, -25767.257);
    final Vector3D velocity = new Vector3D(505.848, 942.781, 7435.922);
    final AbsoluteDate date = new AbsoluteDate(2003, 9, 16, utc);
    final Orbit orbit = new EquinoctialOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), date, mu);
    Propagator propagator = new EcksteinHechlerPropagator(orbit, ae, mu, c20, c30, c40, c50, c60);
    // 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);
    GeodeticPoint point = new GeodeticPoint(FastMath.toRadians(48.833), FastMath.toRadians(2.333), 0.0);
    TopocentricFrame topo = new TopocentricFrame(earth, point, "Gstation");
    AtmosphericRefractionModel refractionModel = new EarthStandardAtmosphereRefraction();
    ElevationDetector detector = new ElevationDetector(topo).withRefraction(refractionModel).withHandler(new StopOnIncreasing<ElevationDetector>());
    Assert.assertSame(refractionModel, detector.getRefractionModel());
    AbsoluteDate startDate = new AbsoluteDate(2003, 9, 15, 20, 0, 0, utc);
    propagator.resetInitialState(propagator.propagate(startDate));
    propagator.addEventDetector(detector);
    final SpacecraftState fs = propagator.propagate(startDate.shiftedBy(Constants.JULIAN_DAY));
    double elevation = topo.getElevation(fs.getPVCoordinates().getPosition(), fs.getFrame(), fs.getDate());
    Assert.assertEquals(FastMath.toRadians(-0.5746255623877098), elevation, 2.0e-5);
}

Example 38 with EquinoctialOrbit

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

the class EventDetectorTest method testNoisyGFunction.

@Test
public void testNoisyGFunction() throws OrekitException {
    // initial conditions
    Frame eme2000 = FramesFactory.getEME2000();
    TimeScale utc = TimeScalesFactory.getUTC();
    AbsoluteDate initialDate = new AbsoluteDate(2011, 5, 11, utc);
    AbsoluteDate startDate = new AbsoluteDate(2032, 10, 17, utc);
    AbsoluteDate interruptDate = new AbsoluteDate(2032, 10, 18, utc);
    AbsoluteDate targetDate = new AbsoluteDate(2211, 5, 11, utc);
    KeplerianPropagator k1 = new KeplerianPropagator(new EquinoctialOrbit(new PVCoordinates(new Vector3D(4008462.4706055815, -3155502.5373837613, -5044275.9880020910), new Vector3D(-5012.9298276860990, 1920.3567095973078, -5172.7403501801580)), eme2000, initialDate, Constants.WGS84_EARTH_MU));
    KeplerianPropagator k2 = new KeplerianPropagator(new EquinoctialOrbit(new PVCoordinates(new Vector3D(4008912.4039522274, -3155453.3125615157, -5044297.6484738905), new Vector3D(-5012.5883854112530, 1920.6332221785074, -5172.2177085540500)), eme2000, initialDate, Constants.WGS84_EARTH_MU));
    k2.addEventDetector(new CloseApproachDetector(2015.243454166727, 0.0001, 100, new ContinueOnEvent<CloseApproachDetector>(), k1));
    k2.addEventDetector(new DateDetector(Constants.JULIAN_DAY, 1.0e-6, interruptDate));
    SpacecraftState s = k2.propagate(startDate, targetDate);
    Assert.assertEquals(0.0, interruptDate.durationFrom(s.getDate()), 1.1e-6);
}

Example 39 with EquinoctialOrbit

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

the class EventsLoggerTest method setUp.

@Before
public void setUp() {
    try {
        Utils.setDataRoot("regular-data");
        mu = 3.9860047e14;
        final Vector3D position = new Vector3D(-6142438.668, 3492467.560, -25767.25680);
        final Vector3D velocity = new Vector3D(505.8479685, 942.7809215, 7435.922231);
        iniDate = new AbsoluteDate(1969, 7, 28, 4, 0, 0.0, TimeScalesFactory.getTT());
        final Orbit orbit = new EquinoctialOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), iniDate, mu);
        initialState = new SpacecraftState(orbit);
        double[] absTolerance = { 0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001 };
        double[] relTolerance = { 1.0e-7, 1.0e-4, 1.0e-4, 1.0e-7, 1.0e-7, 1.0e-7, 1.0e-7 };
        AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 1000, absTolerance, relTolerance);
        integrator.setInitialStepSize(60);
        propagator = new NumericalPropagator(integrator);
        propagator.setInitialState(initialState);
        count = 0;
        umbraDetector = buildDetector(true);
        penumbraDetector = buildDetector(false);
    } catch (OrekitException oe) {
        Assert.fail(oe.getLocalizedMessage());
    }
}

Example 40 with EquinoctialOrbit

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

the class KeplerianPropagatorTest method propagatedCartesian.

@Test
public void propagatedCartesian() throws OrekitException {
    // Definition of initial conditions with position and velocity
    // ------------------------------------------------------------
    Vector3D position = new Vector3D(7.0e6, 1.0e6, 4.0e6);
    Vector3D velocity = new Vector3D(-500.0, 8000.0, 1000.0);
    double mu = 3.9860047e14;
    AbsoluteDate initDate = AbsoluteDate.J2000_EPOCH.shiftedBy(584.);
    Orbit initialOrbit = new EquinoctialOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), initDate, mu);
    // Extrapolator definition
    // -----------------------
    KeplerianPropagator extrapolator = new KeplerianPropagator(initialOrbit);
    // Extrapolation at a final date different from initial date
    // ---------------------------------------------------------
    // extrapolation duration in seconds
    double delta_t = 100000.0;
    AbsoluteDate extrapDate = initDate.shiftedBy(delta_t);
    SpacecraftState finalOrbit = extrapolator.propagate(extrapDate);
    // computation of (M final - M initial) with another method
    double a = finalOrbit.getA();
    // another way to compute n
    double n = FastMath.sqrt(finalOrbit.getMu() / FastMath.pow(a, 3));
    Assert.assertEquals(n * delta_t, finalOrbit.getLM() - initialOrbit.getLM(), Utils.epsilonAngle);
    // computation of M final orbit
    double LM = finalOrbit.getLE() - finalOrbit.getEquinoctialEx() * FastMath.sin(finalOrbit.getLE()) + finalOrbit.getEquinoctialEy() * FastMath.cos(finalOrbit.getLE());
    Assert.assertEquals(LM, finalOrbit.getLM(), Utils.epsilonAngle);
    // test of tan ((LE - Lv)/2) :
    Assert.assertEquals(FastMath.tan((finalOrbit.getLE() - finalOrbit.getLv()) / 2.), tangLEmLv(finalOrbit.getLv(), finalOrbit.getEquinoctialEx(), finalOrbit.getEquinoctialEy()), Utils.epsilonAngle);
    // test of evolution of M vs E: LM = LE - ex*sin(LE) + ey*cos(LE)
    // with ex and ey the same for initial and final orbit
    double deltaM = finalOrbit.getLM() - initialOrbit.getLM();
    double deltaE = finalOrbit.getLE() - initialOrbit.getLE();
    double delta = finalOrbit.getEquinoctialEx() * (FastMath.sin(finalOrbit.getLE()) - FastMath.sin(initialOrbit.getLE())) - finalOrbit.getEquinoctialEy() * (FastMath.cos(finalOrbit.getLE()) - FastMath.cos(initialOrbit.getLE()));
    Assert.assertEquals(deltaM, deltaE - delta, Utils.epsilonAngle);
    // the orbital elements except for Mean/True/Eccentric latitude arguments are the same
    Assert.assertEquals(finalOrbit.getA(), initialOrbit.getA(), Utils.epsilonTest * initialOrbit.getA());
    Assert.assertEquals(finalOrbit.getEquinoctialEx(), initialOrbit.getEquinoctialEx(), Utils.epsilonE);
    Assert.assertEquals(finalOrbit.getEquinoctialEy(), initialOrbit.getEquinoctialEy(), Utils.epsilonE);
    Assert.assertEquals(finalOrbit.getHx(), initialOrbit.getHx(), Utils.epsilonAngle);
    Assert.assertEquals(finalOrbit.getHy(), initialOrbit.getHy(), Utils.epsilonAngle);
    // for final orbit
    double ex = finalOrbit.getEquinoctialEx();
    double ey = finalOrbit.getEquinoctialEy();
    double hx = finalOrbit.getHx();
    double hy = finalOrbit.getHy();
    double LE = finalOrbit.getLE();
    double ex2 = ex * ex;
    double ey2 = ey * ey;
    double hx2 = hx * hx;
    double hy2 = hy * hy;
    double h2p1 = 1. + hx2 + hy2;
    double beta = 1. / (1. + FastMath.sqrt(1. - ex2 - ey2));
    double x3 = -ex + (1. - beta * ey2) * FastMath.cos(LE) + beta * ex * ey * FastMath.sin(LE);
    double y3 = -ey + (1. - beta * ex2) * FastMath.sin(LE) + beta * ex * ey * FastMath.cos(LE);
    Vector3D U = new Vector3D((1. + hx2 - hy2) / h2p1, (2. * hx * hy) / h2p1, (-2. * hy) / h2p1);
    Vector3D V = new Vector3D((2. * hx * hy) / h2p1, (1. - hx2 + hy2) / h2p1, (2. * hx) / h2p1);
    Vector3D r = new Vector3D(finalOrbit.getA(), new Vector3D(x3, U, y3, V));
    Assert.assertEquals(finalOrbit.getPVCoordinates().getPosition().getNorm(), r.getNorm(), Utils.epsilonTest * r.getNorm());
}