Examples with LofOffset org.orekit.attitudes.LofOffset used on opensource projects

Example 21 with LofOffset

use of org.orekit.attitudes.LofOffset in project Orekit by CS-SI.

the class EphemerisTest method testAttitudeOverride.

@Test
public void testAttitudeOverride() throws IllegalArgumentException, OrekitException {
    final double positionTolerance = 1e-6;
    final double velocityTolerance = 1e-5;
    final double attitudeTolerance = 1e-6;
    int numberOfInterals = 1440;
    double deltaT = finalDate.durationFrom(initDate) / ((double) numberOfInterals);
    propagator.setAttitudeProvider(new LofOffset(inertialFrame, LOFType.VVLH));
    List<SpacecraftState> states = new ArrayList<SpacecraftState>(numberOfInterals + 1);
    for (int j = 0; j <= numberOfInterals; j++) {
        states.add(propagator.propagate(initDate.shiftedBy((j * deltaT))));
    }
    int numInterpolationPoints = 2;
    Ephemeris ephemPropagator = new Ephemeris(states, numInterpolationPoints);
    Assert.assertEquals(0, ephemPropagator.getManagedAdditionalStates().length);
    // First test that we got position, velocity and attitude nailed
    int numberEphemTestIntervals = 2880;
    deltaT = finalDate.durationFrom(initDate) / ((double) numberEphemTestIntervals);
    for (int j = 0; j <= numberEphemTestIntervals; j++) {
        AbsoluteDate currentDate = initDate.shiftedBy(j * deltaT);
        SpacecraftState ephemState = ephemPropagator.propagate(currentDate);
        SpacecraftState keplerState = propagator.propagate(currentDate);
        double positionDelta = calculatePositionDelta(ephemState, keplerState);
        double velocityDelta = calculateVelocityDelta(ephemState, keplerState);
        double attitudeDelta = calculateAttitudeDelta(ephemState, keplerState);
        Assert.assertEquals("VVLH Unmatched Position at: " + currentDate, 0.0, positionDelta, positionTolerance);
        Assert.assertEquals("VVLH Unmatched Velocity at: " + currentDate, 0.0, velocityDelta, velocityTolerance);
        Assert.assertEquals("VVLH Unmatched Attitude at: " + currentDate, 0.0, attitudeDelta, attitudeTolerance);
    }
    // Now force an override on the attitude and check it against a Keplerian propagator
    // setup identically to the first but with a different attitude
    // If override isn't working this will fail.
    propagator = new KeplerianPropagator(propagator.getInitialState().getOrbit());
    propagator.setAttitudeProvider(new LofOffset(inertialFrame, LOFType.QSW));
    ephemPropagator.setAttitudeProvider(new LofOffset(inertialFrame, LOFType.QSW));
    for (int j = 0; j <= numberEphemTestIntervals; j++) {
        AbsoluteDate currentDate = initDate.shiftedBy(j * deltaT);
        SpacecraftState ephemState = ephemPropagator.propagate(currentDate);
        SpacecraftState keplerState = propagator.propagate(currentDate);
        double positionDelta = calculatePositionDelta(ephemState, keplerState);
        double velocityDelta = calculateVelocityDelta(ephemState, keplerState);
        double attitudeDelta = calculateAttitudeDelta(ephemState, keplerState);
        Assert.assertEquals("QSW Unmatched Position at: " + currentDate, 0.0, positionDelta, positionTolerance);
        Assert.assertEquals("QSW Unmatched Velocity at: " + currentDate, 0.0, velocityDelta, velocityTolerance);
        Assert.assertEquals("QSW Unmatched Attitude at: " + currentDate, 0.0, attitudeDelta, attitudeTolerance);
    }
}

Example 22 with LofOffset

use of org.orekit.attitudes.LofOffset in project Orekit by CS-SI.

the class EphemerisTest method testProtectedMethods.

@Test
public void testProtectedMethods() throws OrekitException, SecurityException, NoSuchMethodException, InvocationTargetException, IllegalAccessException {
    propagator.setAttitudeProvider(new LofOffset(inertialFrame, LOFType.VVLH));
    List<SpacecraftState> states = new ArrayList<SpacecraftState>();
    for (double dt = 0; dt >= -1200; dt -= 60.0) {
        final SpacecraftState original = propagator.propagate(initDate.shiftedBy(dt));
        final SpacecraftState modified = new SpacecraftState(original.getOrbit(), original.getAttitude(), original.getMass() - 0.0625 * dt);
        states.add(modified);
    }
    final Propagator ephem = new Ephemeris(states, 2);
    Method propagateOrbit = Ephemeris.class.getDeclaredMethod("propagateOrbit", AbsoluteDate.class);
    propagateOrbit.setAccessible(true);
    Method getMass = Ephemeris.class.getDeclaredMethod("getMass", AbsoluteDate.class);
    getMass.setAccessible(true);
    SpacecraftState s = ephem.propagate(initDate.shiftedBy(-270.0));
    Orbit o = (Orbit) propagateOrbit.invoke(ephem, s.getDate());
    double m = ((Double) getMass.invoke(ephem, s.getDate())).doubleValue();
    Assert.assertEquals(0.0, Vector3D.distance(s.getPVCoordinates().getPosition(), o.getPVCoordinates().getPosition()), 1.0e-15);
    Assert.assertEquals(s.getMass(), m, 1.0e-15);
}

Example 23 with LofOffset

use of org.orekit.attitudes.LofOffset in project Orekit by CS-SI.

the class FieldEcksteinHechlerPropagatorTest method doPropagatedKeplerian.

private <T extends RealFieldElement<T>> void doPropagatedKeplerian(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    FieldAbsoluteDate<T> date = new FieldAbsoluteDate<>(field);
    // Definition of initial conditions with Keplerian parameters
    // -----------------------------------------------------------
    FieldAbsoluteDate<T> initDate = date.shiftedBy(584.);
    FieldOrbit<T> initialOrbit = new FieldKeplerianOrbit<>(zero.add(7209668.0), zero.add(0.5e-4), zero.add(1.7), zero.add(2.1), zero.add(2.9), zero.add(6.2), PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, provider.getMu());
    // Extrapolator definition
    // -----------------------
    FieldEcksteinHechlerPropagator<T> extrapolator = new FieldEcksteinHechlerPropagator<>(initialOrbit, new LofOffset(initialOrbit.getFrame(), LOFType.VNC), zero.add(2000.0), provider);
    // Extrapolation at a final date different from initial date
    // ---------------------------------------------------------
    // extrapolation duration in seconds
    double delta_t = 100000.0;
    FieldAbsoluteDate<T> extrapDate = initDate.shiftedBy(delta_t);
    FieldSpacecraftState<T> finalOrbit = extrapolator.propagate(extrapDate);
    Assert.assertEquals(0.0, finalOrbit.getDate().durationFrom(extrapDate).getReal(), 1.0e-9);
    // computation of M final orbit
    T LM = finalOrbit.getLE().subtract(finalOrbit.getEquinoctialEx().multiply(finalOrbit.getLE().sin())).add(finalOrbit.getEquinoctialEy().multiply(finalOrbit.getLE().cos()));
    Assert.assertEquals(LM.getReal(), finalOrbit.getLM().getReal(), Utils.epsilonAngle);
    // test of tan((LE - Lv)/2) :
    Assert.assertEquals(FastMath.tan((finalOrbit.getLE().getReal() - finalOrbit.getLv().getReal()) / 2.), tangLEmLv(finalOrbit.getLv(), finalOrbit.getEquinoctialEx(), finalOrbit.getEquinoctialEy()).getReal(), 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
    T deltaM = finalOrbit.getLM().subtract(initialOrbit.getLM());
    T deltaE = finalOrbit.getLE().subtract(initialOrbit.getLE());
    T delta = finalOrbit.getEquinoctialEx().multiply(finalOrbit.getLE().sin()).subtract(initialOrbit.getEquinoctialEx().multiply(initialOrbit.getLE().sin())).subtract(finalOrbit.getEquinoctialEy().multiply(finalOrbit.getLE().cos())).add(initialOrbit.getEquinoctialEy().multiply(initialOrbit.getLE().cos()));
    Assert.assertEquals(deltaM.getReal(), deltaE.getReal() - delta.getReal(), Utils.epsilonAngle * FastMath.abs(deltaE.getReal() - delta.getReal()));
    // for final orbit
    T ex = finalOrbit.getEquinoctialEx();
    T ey = finalOrbit.getEquinoctialEy();
    T hx = finalOrbit.getHx();
    T hy = finalOrbit.getHy();
    T LE = finalOrbit.getLE();
    T ex2 = ex.multiply(ex);
    T ey2 = ey.multiply(ey);
    T hx2 = hx.multiply(hx);
    T hy2 = hy.multiply(hy);
    T h2p1 = hx2.add(1).add(hy2);
    T beta = ex2.negate().add(1.).subtract(ey2).sqrt().add(1).reciprocal();
    T x3 = ex.negate().add(beta.negate().multiply(ey2).add(1).multiply(LE.cos())).add(beta.multiply(ex).multiply(ey).multiply(LE.sin()));
    T y3 = ey.negate().add(beta.negate().multiply(ex2).add(1).multiply(LE.sin())).add(beta.multiply(ex).multiply(ey).multiply(LE.cos()));
    FieldVector3D<T> U = new FieldVector3D<>(hx2.subtract(hy2).add(1.).divide(h2p1), hx.multiply(2).multiply(hy).divide(h2p1), hy.multiply(-2.).divide(h2p1));
    FieldVector3D<T> V = new FieldVector3D<>(hx.multiply(2.).multiply(hy).divide(h2p1), hy2.subtract(hx2).add(1.).divide(h2p1), hx.multiply(2).divide(h2p1));
    FieldVector3D<T> r = new FieldVector3D<>(finalOrbit.getA(), new FieldVector3D<>(x3, U, y3, V));
    Assert.assertEquals(finalOrbit.getPVCoordinates().getPosition().getNorm().getReal(), r.getNorm().getReal(), Utils.epsilonTest * r.getNorm().getReal());
}

Example 24 with LofOffset

use of org.orekit.attitudes.LofOffset in project Orekit by CS-SI.

the class FieldEcksteinHechlerPropagatorTest method doPropagatedCartesian.

private <T extends RealFieldElement<T>> void doPropagatedCartesian(Field<T> field) throws OrekitException {
    T zero = field.getZero();
    FieldAbsoluteDate<T> date = new FieldAbsoluteDate<>(field);
    // Definition of initial conditions with position and velocity
    // ------------------------------------------------------------
    // with e around e = 1.4e-4 and i = 1.7 rad
    FieldVector3D<T> position = new FieldVector3D<>(zero.add(3220103.), zero.add(69623.), zero.add(6449822.));
    FieldVector3D<T> velocity = new FieldVector3D<>(zero.add(6414.7), zero.add(-2006.), zero.add(-3180.));
    FieldAbsoluteDate<T> initDate = date.shiftedBy(584.);
    FieldOrbit<T> initialOrbit = new FieldEquinoctialOrbit<>(new FieldPVCoordinates<>(position, velocity), FramesFactory.getEME2000(), initDate, provider.getMu());
    // Extrapolator definition
    // -----------------------
    FieldEcksteinHechlerPropagator<T> extrapolator = new FieldEcksteinHechlerPropagator<>(initialOrbit, new LofOffset(initialOrbit.getFrame(), LOFType.VNC, RotationOrder.XYZ, 0, 0, 0), provider);
    // Extrapolation at a final date different from initial date
    // ---------------------------------------------------------
    // extrapolation duration in seconds
    double delta_t = 100000.0;
    FieldAbsoluteDate<T> extrapDate = initDate.shiftedBy(delta_t);
    FieldSpacecraftState<T> finalOrbit = extrapolator.propagate(extrapDate);
    Assert.assertEquals(0.0, finalOrbit.getDate().durationFrom(extrapDate).getReal(), 1.0e-9);
    // computation of M final orbit
    T LM = finalOrbit.getLE().subtract(finalOrbit.getEquinoctialEx().multiply(finalOrbit.getLE().sin())).add(finalOrbit.getEquinoctialEy().multiply(finalOrbit.getLE().cos()));
    Assert.assertEquals(LM.getReal(), finalOrbit.getLM().getReal(), Utils.epsilonAngle * FastMath.abs(finalOrbit.getLM().getReal()));
    // test of tan ((LE - Lv)/2) :
    Assert.assertEquals(FastMath.tan((finalOrbit.getLE().getReal() - finalOrbit.getLv().getReal()) / 2.), tangLEmLv(finalOrbit.getLv(), finalOrbit.getEquinoctialEx(), finalOrbit.getEquinoctialEy()).getReal(), Utils.epsilonAngle);
    // test of evolution of M vs E: LM = LE - ex*sin(LE) + ey*cos(LE)
    T deltaM = finalOrbit.getLM().subtract(initialOrbit.getLM());
    T deltaE = finalOrbit.getLE().subtract(initialOrbit.getLE());
    T delta = finalOrbit.getEquinoctialEx().multiply(finalOrbit.getLE().sin()).subtract(initialOrbit.getEquinoctialEx().multiply(initialOrbit.getLE().sin())).subtract(finalOrbit.getEquinoctialEy().multiply(finalOrbit.getLE().cos())).add(initialOrbit.getEquinoctialEy().multiply(initialOrbit.getLE().cos()));
    Assert.assertEquals(deltaM.getReal(), deltaE.getReal() - delta.getReal(), Utils.epsilonAngle * FastMath.abs(deltaE.getReal() - delta.getReal()));
    // for final orbit
    T ex = finalOrbit.getEquinoctialEx();
    T ey = finalOrbit.getEquinoctialEy();
    T hx = finalOrbit.getHx();
    T hy = finalOrbit.getHy();
    T LE = finalOrbit.getLE();
    T ex2 = ex.multiply(ex);
    T ey2 = ey.multiply(ey);
    T hx2 = hx.multiply(hx);
    T hy2 = hy.multiply(hy);
    T h2p1 = hx2.add(1.).add(hy2);
    T beta = ex2.negate().add(1.).subtract(ey2).sqrt().add(1.).reciprocal();
    T x3 = ex.negate().add(ey2.multiply(beta).negate().add(1.).multiply(LE.cos())).add(beta.multiply(ex).multiply(ey).multiply(LE.sin()));
    T y3 = ey.negate().add(ex2.negate().multiply(beta).add(1).multiply(LE.sin())).add(beta.multiply(ex).multiply(ey).multiply(LE.cos()));
    FieldVector3D<T> U = new FieldVector3D<>(hx2.add(1).subtract(hy2).divide(h2p1), hx.multiply(hy).multiply(2).divide(h2p1), hy.multiply(-2).divide(h2p1));
    FieldVector3D<T> V = new FieldVector3D<>(hx.multiply(2).multiply(hy).divide(h2p1), hy2.add(1).subtract(hx2).divide(h2p1), hx.multiply(2).divide(h2p1));
    FieldVector3D<T> r = new FieldVector3D<>(finalOrbit.getA(), new FieldVector3D<>(x3, U, y3, V));
    Assert.assertEquals(finalOrbit.getPVCoordinates().getPosition().getNorm().getReal(), r.getNorm().getReal(), Utils.epsilonTest * r.getNorm().getReal());
}

Example 25 with LofOffset

use of org.orekit.attitudes.LofOffset 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());
}