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

Example 36 with KeplerianOrbit

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

the class OrekitStepHandlerTest method testIsInterpolated.

/**
 * Check {@link OrekitStepInterpolator#isPreviousStateInterpolated()} and {@link
 * OrekitStepInterpolator#isCurrentStateInterpolated()}.
 *
 * @throws OrekitException on error.
 */
@Test
public void testIsInterpolated() throws OrekitException {
    // setup
    NumericalPropagator propagator = new NumericalPropagator(new ClassicalRungeKuttaIntegrator(60));
    AbsoluteDate date = AbsoluteDate.J2000_EPOCH;
    Frame eci = FramesFactory.getGCRF();
    SpacecraftState ic = new SpacecraftState(new KeplerianOrbit(6378137 + 500e3, 1e-3, 0, 0, 0, 0, PositionAngle.TRUE, eci, date, Constants.EIGEN5C_EARTH_MU));
    propagator.setInitialState(ic);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    // detector triggers half way through second step
    DateDetector detector = new DateDetector(date.shiftedBy(90)).withHandler(new ContinueOnEvent<>());
    propagator.addEventDetector(detector);
    // action and verify
    Queue<Boolean> expected = new ArrayDeque<>(Arrays.asList(false, false, false, true, true, false));
    propagator.setMasterMode(new OrekitStepHandler() {

        @Override
        public void handleStep(OrekitStepInterpolator interpolator, boolean isLast) {
            assertEquals(expected.poll(), interpolator.isPreviousStateInterpolated());
            assertEquals(expected.poll(), interpolator.isCurrentStateInterpolated());
        }
    });
    final AbsoluteDate end = date.shiftedBy(120);
    assertEquals(end, propagator.propagate(end).getDate());
}

Example 37 with KeplerianOrbit

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

the class FieldNodeDetector method estimateNodesTimeSeparation.

/**
 * Find time separation between nodes.
 * <p>
 * The estimation of time separation is based on Keplerian motion, it is only
 * used as a rough guess for a safe setting of default max check interval for
 * event detection.
 * </p>
 * @param orbit initial orbit
 * @return minimum time separation between nodes
 */
private static double estimateNodesTimeSeparation(final Orbit orbit) {
    final KeplerianOrbit keplerian = (KeplerianOrbit) OrbitType.KEPLERIAN.convertType(orbit);
    // mean anomaly of ascending node
    final double ascendingM = new KeplerianOrbit(keplerian.getA(), keplerian.getE(), keplerian.getI(), keplerian.getPerigeeArgument(), keplerian.getRightAscensionOfAscendingNode(), -keplerian.getPerigeeArgument(), PositionAngle.TRUE, keplerian.getFrame(), keplerian.getDate(), keplerian.getMu()).getMeanAnomaly();
    // mean anomaly of descending node
    final double descendingM = new KeplerianOrbit(keplerian.getA(), keplerian.getE(), keplerian.getI(), keplerian.getPerigeeArgument(), keplerian.getRightAscensionOfAscendingNode(), FastMath.PI - keplerian.getPerigeeArgument(), PositionAngle.TRUE, keplerian.getFrame(), keplerian.getDate(), keplerian.getMu()).getMeanAnomaly();
    // differences between mean anomalies
    final double delta1 = MathUtils.normalizeAngle(ascendingM, descendingM + FastMath.PI) - descendingM;
    final double delta2 = 2 * FastMath.PI - delta1;
    // minimum time separation between the two nodes
    return FastMath.min(delta1, delta2) / keplerian.getKeplerianMeanMotion();
}

Example 38 with KeplerianOrbit

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

the class TLEPropagatorBuilder method buildPropagator.

/**
 * {@inheritDoc}
 */
public Propagator buildPropagator(final double[] normalizedParameters) throws OrekitException {
    // create the orbit
    setParameters(normalizedParameters);
    final Orbit orbit = createInitialOrbit();
    // we really need a Keplerian orbit type
    final KeplerianOrbit kep = (KeplerianOrbit) OrbitType.KEPLERIAN.convertType(orbit);
    final TLE tle = new TLE(satelliteNumber, classification, launchYear, launchNumber, launchPiece, TLE.DEFAULT, elementNumber, orbit.getDate(), kep.getKeplerianMeanMotion(), 0.0, 0.0, kep.getE(), MathUtils.normalizeAngle(orbit.getI(), FastMath.PI), MathUtils.normalizeAngle(kep.getPerigeeArgument(), FastMath.PI), MathUtils.normalizeAngle(kep.getRightAscensionOfAscendingNode(), FastMath.PI), MathUtils.normalizeAngle(kep.getMeanAnomaly(), FastMath.PI), revolutionNumberAtEpoch, bStar);
    return TLEPropagator.selectExtrapolator(tle);
}

Example 39 with KeplerianOrbit

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

the class RelativityTest method testSmallEffectOnOrbit.

/**
 * check against example in Tapley, Schutz, and Born, p 65-66. They predict a
 * progression of perigee of 11 arcsec/year. To get the same results we must set the
 * propagation tolerances to 1e-5.
 *
 * @throws OrekitException on error
 */
@Test
public void testSmallEffectOnOrbit() throws OrekitException {
    // setup
    final double gm = Constants.EIGEN5C_EARTH_MU;
    Orbit orbit = new KeplerianOrbit(7500e3, 0.025, FastMath.toRadians(41.2), 0, 0, 0, PositionAngle.TRUE, frame, date, gm);
    double[][] tol = NumericalPropagator.tolerances(0.00001, orbit, OrbitType.CARTESIAN);
    AbstractIntegrator integrator = new DormandPrince853Integrator(1, 3600, tol[0], tol[1]);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(OrbitType.CARTESIAN);
    propagator.addForceModel(new Relativity(gm));
    propagator.setInitialState(new SpacecraftState(orbit));
    // action: propagate a period
    AbsoluteDate end = orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY);
    PVCoordinates actual = propagator.getPVCoordinates(end, frame);
    // verify
    KeplerianOrbit endOrbit = new KeplerianOrbit(actual, frame, end, gm);
    KeplerianOrbit startOrbit = new KeplerianOrbit(orbit);
    double dp = endOrbit.getPerigeeArgument() - startOrbit.getPerigeeArgument();
    double dtYears = end.durationFrom(orbit.getDate()) / Constants.JULIAN_YEAR;
    double dpDeg = FastMath.toDegrees(dp);
    // change in argument of perigee in arcseconds per year
    double arcsecPerYear = dpDeg * 3600 / dtYears;
    Assert.assertEquals(11, arcsecPerYear, 0.5);
}

Example 40 with KeplerianOrbit

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

the class SolidTidesTest method testTideEffect1996.

@Test
public void testTideEffect1996() throws OrekitException {
    Frame eme2000 = FramesFactory.getEME2000();
    TimeScale utc = TimeScalesFactory.getUTC();
    AbsoluteDate date = new AbsoluteDate(2003, 07, 01, 13, 59, 27.816, utc);
    Orbit orbit = new KeplerianOrbit(7201009.7124401, 1e-3, FastMath.toRadians(98.7), FastMath.toRadians(93.0), FastMath.toRadians(15.0 * 22.5), 0, PositionAngle.MEAN, eme2000, date, Constants.EIGEN5C_EARTH_MU);
    doTestTideEffect(orbit, IERSConventions.IERS_1996, 44.09481, 0.00000);
}