Examples with Frame org.orekit.frames.Frame used on opensource projects

Example 91 with Frame

use of org.orekit.frames.Frame in project Orekit by CS-SI.

the class FieldKeplerianPropagatorTest method doTestWrappedAttitudeException.

private <T extends RealFieldElement<T>> void doTestWrappedAttitudeException(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    final FieldKeplerianOrbit<T> orbit = new FieldKeplerianOrbit<>(zero.add(7.8e6), zero.add(0.032), zero.add(0.4), zero.add(0.1), zero.add(0.2), zero.add(0.3), PositionAngle.TRUE, FramesFactory.getEME2000(), new FieldAbsoluteDate<>(field), 3.986004415e14);
    FieldKeplerianPropagator<T> propagator = new FieldKeplerianPropagator<>(orbit, new AttitudeProvider() {

        private static final long serialVersionUID = 1L;

        public Attitude getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) throws OrekitException {
            throw new OrekitException((Throwable) null, new DummyLocalizable("dummy error"));
        }

        public <Q extends RealFieldElement<Q>> FieldAttitude<Q> getAttitude(FieldPVCoordinatesProvider<Q> pvProv, FieldAbsoluteDate<Q> date, Frame frame) throws OrekitException {
            throw new OrekitException((Throwable) null, new DummyLocalizable("dummy error"));
        }
    });
    propagator.propagate(orbit.getDate().shiftedBy(10.09));
}

Example 92 with Frame

use of org.orekit.frames.Frame 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 93 with Frame

use of org.orekit.frames.Frame 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);
}

Example 94 with Frame

use of org.orekit.frames.Frame in project Orekit by CS-SI.

the class TabulatedEphemerisTest method testPiWraping.

@Test
public void testPiWraping() throws OrekitException {
    TimeScale utc = TimeScalesFactory.getUTC();
    Frame frame = FramesFactory.getEME2000();
    double mu = CelestialBodyFactory.getEarth().getGM();
    AbsoluteDate t0 = new AbsoluteDate(2009, 10, 29, 0, 0, 0, utc);
    AbsoluteDate t1 = new AbsoluteDate(t0, 1320.0);
    Vector3D p1 = new Vector3D(-0.17831296727974E+08, 0.67919502669856E+06, -0.16591008368477E+07);
    Vector3D v1 = new Vector3D(-0.38699705630724E+04, -0.36209408682762E+04, -0.16255053872347E+03);
    SpacecraftState s1 = new SpacecraftState(new EquinoctialOrbit(new PVCoordinates(p1, v1), frame, t1, mu));
    AbsoluteDate t2 = new AbsoluteDate(t0, 1440.0);
    Vector3D p2 = new Vector3D(-0.18286942572033E+08, 0.24442124296930E+06, -0.16777961761695E+07);
    Vector3D v2 = new Vector3D(-0.37252897467918E+04, -0.36246628128896E+04, -0.14917724596280E+03);
    SpacecraftState s2 = new SpacecraftState(new EquinoctialOrbit(new PVCoordinates(p2, v2), frame, t2, mu));
    AbsoluteDate t3 = new AbsoluteDate(t0, 1560.0);
    Vector3D p3 = new Vector3D(-0.18725635245837E+08, -0.19058407701834E+06, -0.16949352249614E+07);
    Vector3D v3 = new Vector3D(-0.35873348682393E+04, -0.36248828501784E+04, -0.13660045394149E+03);
    SpacecraftState s3 = new SpacecraftState(new EquinoctialOrbit(new PVCoordinates(p3, v3), frame, t3, mu));
    Ephemeris ephem = new Ephemeris(Arrays.asList(s1, s2, s3), 2);
    AbsoluteDate tA = new AbsoluteDate(t0, 24 * 60);
    Vector3D pA = ephem.propagate(tA).getPVCoordinates(frame).getPosition();
    Assert.assertEquals(1.766, Vector3D.distance(pA, s1.shiftedBy(tA.durationFrom(s1.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    Assert.assertEquals(0.000, Vector3D.distance(pA, s2.shiftedBy(tA.durationFrom(s2.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    Assert.assertEquals(1.556, Vector3D.distance(pA, s3.shiftedBy(tA.durationFrom(s3.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    AbsoluteDate tB = new AbsoluteDate(t0, 25 * 60);
    Vector3D pB = ephem.propagate(tB).getPVCoordinates(frame).getPosition();
    Assert.assertEquals(2.646, Vector3D.distance(pB, s1.shiftedBy(tB.durationFrom(s1.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    Assert.assertEquals(2.619, Vector3D.distance(pB, s2.shiftedBy(tB.durationFrom(s2.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    Assert.assertEquals(2.632, Vector3D.distance(pB, s3.shiftedBy(tB.durationFrom(s3.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    AbsoluteDate tC = new AbsoluteDate(t0, 26 * 60);
    Vector3D pC = ephem.propagate(tC).getPVCoordinates(frame).getPosition();
    Assert.assertEquals(6.851, Vector3D.distance(pC, s1.shiftedBy(tC.durationFrom(s1.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    Assert.assertEquals(1.605, Vector3D.distance(pC, s2.shiftedBy(tC.durationFrom(s2.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
    Assert.assertEquals(0.000, Vector3D.distance(pC, s3.shiftedBy(tC.durationFrom(s3.getDate())).getPVCoordinates(frame).getPosition()), 1.0e-3);
}

Example 95 with Frame

use of org.orekit.frames.Frame in project Orekit by CS-SI.

the class TabulatedEphemerisTest method testGetFrame.

@Test
public void testGetFrame() throws MathIllegalArgumentException, IllegalArgumentException, OrekitException {
    // setup
    Frame frame = FramesFactory.getICRF();
    AbsoluteDate date = AbsoluteDate.JULIAN_EPOCH;
    // create ephemeris with 2 arbitrary points
    SpacecraftState state = new SpacecraftState(new KeplerianOrbit(1e9, 0.01, 1, 1, 1, 1, PositionAngle.TRUE, frame, date, mu));
    Ephemeris ephem = new Ephemeris(Arrays.asList(state, state.shiftedBy(1)), 2);
    // action + verify
    Assert.assertSame(ephem.getFrame(), frame);
}