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

Example 81 with Orbit

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

the class KeplerianPropagatorTest method testEphemerisModeWithHandler.

@Test
public void testEphemerisModeWithHandler() throws OrekitException {
    // setup
    AbsoluteDate initDate = AbsoluteDate.GPS_EPOCH;
    Orbit ic = new KeplerianOrbit(6378137 + 500e3, 1e-3, 0, 0, 0, 0, PositionAngle.TRUE, FramesFactory.getGCRF(), initDate, mu);
    Propagator propagator = new KeplerianPropagator(ic);
    AbsoluteDate end = initDate.shiftedBy(90 * 60);
    // action
    final List<SpacecraftState> states = new ArrayList<>();
    propagator.setEphemerisMode((interpolator, isLast) -> {
        states.add(interpolator.getCurrentState());
        states.add(interpolator.getPreviousState());
    });
    propagator.propagate(end);
    final BoundedPropagator ephemeris = propagator.getGeneratedEphemeris();
    // verify
    // got some data
    Assert.assertTrue(states.size() > 1);
    for (SpacecraftState state : states) {
        PVCoordinates actual = ephemeris.propagate(state.getDate()).getPVCoordinates();
        Assert.assertThat(actual, OrekitMatchers.pvIs(state.getPVCoordinates()));
    }
}

Example 82 with Orbit

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

the class KeplerianPropagatorTest method sameDateKeplerian.

@Test
public void sameDateKeplerian() throws OrekitException {
    // Definition of initial conditions with Keplerian parameters
    // -----------------------------------------------------------
    AbsoluteDate initDate = AbsoluteDate.J2000_EPOCH.shiftedBy(584.);
    Orbit initialOrbit = new KeplerianOrbit(7209668.0, 0.5e-4, 1.7, 2.1, 2.9, 6.2, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, mu);
    // Extrapolator definition
    // -----------------------
    KeplerianPropagator extrapolator = new KeplerianPropagator(initialOrbit);
    // Extrapolation at the initial date
    // ---------------------------------
    // extrapolation duration in seconds
    double delta_t = 0.0;
    AbsoluteDate extrapDate = initDate.shiftedBy(delta_t);
    SpacecraftState finalOrbit = extrapolator.propagate(extrapDate);
    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.epsilonTest * FastMath.max(100., FastMath.abs(n * delta_t)));
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbit.getLM(), initialOrbit.getLM()), initialOrbit.getLM(), Utils.epsilonAngle * FastMath.abs(initialOrbit.getLM()));
    Assert.assertEquals(finalOrbit.getA(), initialOrbit.getA(), Utils.epsilonTest * initialOrbit.getA());
    Assert.assertEquals(finalOrbit.getE(), initialOrbit.getE(), Utils.epsilonE * initialOrbit.getE());
    Assert.assertEquals(MathUtils.normalizeAngle(finalOrbit.getI(), initialOrbit.getI()), initialOrbit.getI(), Utils.epsilonAngle * FastMath.abs(initialOrbit.getI()));
}

Example 83 with Orbit

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

the class KeplerianPropagatorTest method testIssue107.

@Test
public void testIssue107() 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 CircularOrbit(new PVCoordinates(position, velocity), FramesFactory.getEME2000(), date, mu);
    Propagator propagator = new KeplerianPropagator(orbit) {

        private static final long serialVersionUID = 1L;

        AbsoluteDate lastDate = AbsoluteDate.PAST_INFINITY;

        protected SpacecraftState basicPropagate(final AbsoluteDate date) throws OrekitException {
            if (date.compareTo(lastDate) < 0) {
                throw new OrekitException(LocalizedCoreFormats.SIMPLE_MESSAGE, "no backward propagation allowed");
            }
            lastDate = date;
            return super.basicPropagate(date);
        }
    };
    SpacecraftState finalState = propagator.propagate(date.shiftedBy(3600.0));
    Assert.assertEquals(3600.0, finalState.getDate().durationFrom(date), 1.0e-15);
}

Example 84 with Orbit

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

the class KeplerianPropagatorTest method testIssue224.

@Test
public void testIssue224() throws OrekitException, IOException, ClassNotFoundException {
    // Inertial frame
    Frame inertialFrame = FramesFactory.getEME2000();
    // Initial date
    TimeScale utc = TimeScalesFactory.getUTC();
    AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, utc);
    // Central attraction coefficient
    double mu = 3.986004415e+14;
    // Initial orbit
    // semi major axis in meters
    double a = 42100;
    // eccentricity
    double e = 0.01;
    // inclination
    double i = FastMath.toRadians(6);
    // perigee argument
    double omega = FastMath.toRadians(180);
    // right ascention of ascending node
    double raan = FastMath.toRadians(261);
    // mean anomaly
    double lM = 0;
    Orbit initialOrbit = new KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN, inertialFrame, initialDate, mu);
    // Initial state definition
    SpacecraftState initialState = new SpacecraftState(initialOrbit);
    // Propagator
    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit, new LofOffset(inertialFrame, LOFType.VVLH));
    propagator.addAdditionalStateProvider(new SevenProvider());
    propagator.setEphemerisMode();
    // Impulsive burn 1
    final AbsoluteDate burn1Date = initialState.getDate().shiftedBy(200);
    ImpulseManeuver<DateDetector> impulsiveBurn1 = new ImpulseManeuver<DateDetector>(new DateDetector(burn1Date), new Vector3D(1000, 0, 0), 320);
    propagator.addEventDetector(impulsiveBurn1);
    // Impulsive burn 2
    final AbsoluteDate burn2Date = initialState.getDate().shiftedBy(300);
    ImpulseManeuver<DateDetector> impulsiveBurn2 = new ImpulseManeuver<DateDetector>(new DateDetector(burn2Date), new Vector3D(1000, 0, 0), 320);
    propagator.addEventDetector(impulsiveBurn2);
    propagator.propagate(initialState.getDate().shiftedBy(400));
    ByteArrayOutputStream bos = new ByteArrayOutputStream();
    ObjectOutputStream oos = new ObjectOutputStream(bos);
    oos.writeObject(propagator.getGeneratedEphemeris());
    Assert.assertTrue(bos.size() > 2400);
    Assert.assertTrue(bos.size() < 2500);
    ByteArrayInputStream bis = new ByteArrayInputStream(bos.toByteArray());
    ObjectInputStream ois = new ObjectInputStream(bis);
    BoundedPropagator ephemeris = (BoundedPropagator) ois.readObject();
    ephemeris.setMasterMode(10, new OrekitFixedStepHandler() {

        public void handleStep(SpacecraftState currentState, boolean isLast) {
            if (currentState.getDate().durationFrom(burn1Date) < -0.001) {
                Assert.assertEquals(42100.0, currentState.getA(), 1.0e-3);
            } else if (currentState.getDate().durationFrom(burn1Date) > 0.001 && currentState.getDate().durationFrom(burn2Date) < -0.001) {
                Assert.assertEquals(42979.962, currentState.getA(), 1.0e-3);
            } else if (currentState.getDate().durationFrom(burn2Date) > 0.001) {
                Assert.assertEquals(43887.339, currentState.getA(), 1.0e-3);
            }
        }
    });
    ephemeris.propagate(ephemeris.getMaxDate());
}

Example 85 with Orbit

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

the class KeplerianPropagatorTest method testIssue223.

@Test
public void testIssue223() throws OrekitException, IOException, ClassNotFoundException {
    // Inertial frame
    Frame inertialFrame = FramesFactory.getEME2000();
    // Initial date
    TimeScale utc = TimeScalesFactory.getUTC();
    AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, utc);
    // Central attraction coefficient
    double mu = 3.986004415e+14;
    // Initial orbit
    // semi major axis in meters
    double a = 42100;
    // eccentricity
    double e = 0.01;
    // inclination
    double i = FastMath.toRadians(6);
    // perigee argument
    double omega = FastMath.toRadians(180);
    // right ascention of ascending node
    double raan = FastMath.toRadians(261);
    // mean anomaly
    double lM = 0;
    Orbit initialOrbit = new KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN, inertialFrame, initialDate, mu);
    // Initial state definition
    SpacecraftState initialState = new SpacecraftState(initialOrbit);
    // Propagator
    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit);
    propagator.addAdditionalStateProvider(new SevenProvider());
    propagator.setEphemerisMode();
    propagator.propagate(initialState.getDate().shiftedBy(40000));
    BoundedPropagator ephemeris = propagator.getGeneratedEphemeris();
    Assert.assertSame(inertialFrame, ephemeris.getFrame());
    ByteArrayOutputStream bos = new ByteArrayOutputStream();
    ObjectOutputStream oos = new ObjectOutputStream(bos);
    oos.writeObject(ephemeris);
    Assert.assertTrue(bos.size() > 2250);
    Assert.assertTrue(bos.size() < 2350);
    ByteArrayInputStream bis = new ByteArrayInputStream(bos.toByteArray());
    ObjectInputStream ois = new ObjectInputStream(bis);
    BoundedPropagator deserialized = (BoundedPropagator) ois.readObject();
    Assert.assertEquals(initialOrbit.getA(), deserialized.getInitialState().getA(), 1.0e-10);
    Assert.assertEquals(initialOrbit.getEquinoctialEx(), deserialized.getInitialState().getEquinoctialEx(), 1.0e-10);
    SpacecraftState s = deserialized.propagate(initialState.getDate().shiftedBy(20000));
    Map<String, double[]> additional = s.getAdditionalStates();
    Assert.assertEquals(1, additional.size());
    Assert.assertEquals(1, additional.get("seven").length);
    Assert.assertEquals(7, additional.get("seven")[0], 1.0e-15);
}