SequenceOfAttitudes 4.4 : Différence entre versions

public class SequenceOfAttitudeLaws {
 
    public static void main(String[] args) throws PatriusException, IOException, URISyntaxException {
 
        // Patrius Dataset initialization (needed for example to get the UTC time
        PatriusDataset.addResourcesFromPatriusDataset() ;
 
        // Recovery of the UTC time scale using a "factory" (not to duplicate such unique object)
        final TimeScale TUC = TimeScalesFactory.getUTC();
 
        // Date of the orbit given in UTC time scale)
        final AbsoluteDate iniDate = new AbsoluteDate("2010-01-01T12:00:00.000", TUC);
 
        // Getting the frame with wich will defined the orbit parameters
        // As for time scale, we will use also a "factory".
        final Frame GCRF = FramesFactory.getGCRF();
 
        // Initial orbit
        final double sma = 7200.e+3;
        final double exc = 0.01;
        final double inc = FastMath.toRadians(98.);
        final double pa = FastMath.toRadians(0.);
        final double raan = FastMath.toRadians(90.);
        final double anm = FastMath.toRadians(0.);
        final double MU = Constants.WGS84_EARTH_MU;
 
        final KeplerianParameters par = new KeplerianParameters(sma, exc, inc, pa, raan, anm, PositionAngle.MEAN, MU);
        final Orbit iniOrbit = new KeplerianOrbit(par, GCRF, iniDate);
 
        // Using the Meeus model for the Sun.
        final CelestialBody sun = new MeeusSun();
        final double sunRadius = Constants.SUN_RADIUS;
 
        // Definition of the Earth ellipsoid for later atmospheric density computation
        final Frame ITRF = FramesFactory.getITRF();
        final double earthRadius = Constants.WGS84_EARTH_EQUATORIAL_RADIUS;
        final GeometricBodyShape earth = new ExtendedOneAxisEllipsoid(earthRadius, Constants.WGS84_EARTH_FLATTENING, ITRF, "EARTH");
 
        // Initializing attitude sequence
        final AttitudesSequence seqAtt = new AttitudesSequence();
 
        // Building a first attitude law (Sun pointing)
        final Vector3D firstAxis = new Vector3D(1., 0., 0.);
        final Vector3D secondAxis = new Vector3D(0., 1., 0.);
        final AttitudeLaw sunPointingLaw = new SunPointing(sun, firstAxis, secondAxis, sun);
 
        // Building a second attitude law (LVLH)
        final AttitudeLaw lvlhLaw = new LofOffset(LOFType.LVLH);
 
        // Events that will switch from a law to another
        final double maxCheck = 10.;
        final double threshold = 1.e-3;
        final EventDetector eventEntryEclipse = new EclipseDetector(sun, sunRadius, earth, earthRadius, 0,
                maxCheck, threshold, Action.RESET_STATE, Action.RESET_STATE);
        final EventDetector eventExitEclipse = new EclipseDetector(sun, sunRadius, earth, earthRadius, 0,
                maxCheck, threshold, Action.RESET_STATE, Action.RESET_STATE);
 
        //Adding switches
        seqAtt.addSwitchingCondition(lvlhLaw, eventEntryEclipse, true, false, sunPointingLaw);
        seqAtt.addSwitchingCondition(sunPointingLaw, eventExitEclipse, false, true, lvlhLaw);
 
        testByPropagation(iniOrbit, seqAtt, sun);
 
    }
 
    public static void testByPropagation ( final Orbit iniOrbit, final AttitudesSequence seqAtt, final CelestialBody sun ) throws PatriusException {
 
        // We create a spacecratftstate
        final SpacecraftState iniState = new SpacecraftState(iniOrbit);
 
        // Initialization of the Runge Kutta integrator with a 2 s step
        final double pasRk = 2.;
        final FirstOrderIntegrator integrator = new ClassicalRungeKuttaIntegrator(pasRk);
 
        // Initialization of the propagator
        final NumericalPropagator propagator = new NumericalPropagator(integrator);
        propagator.resetInitialState(iniState);
 
        // Forcing integration using cartesian equations
        propagator.setOrbitType(OrbitType.CARTESIAN);
 
        // Adding the attitude sequence
        propagator.setAttitudeProvider(seqAtt);
        seqAtt.registerSwitchEvents(propagator);
 
        // Loop every 10 mn ...
        final double step = 600.;
        final double epsilon = 1.e-12;
        for (int i = 1; i <= 20; i++) {
 
            AbsoluteDate date = iniOrbit.getDate().shiftedBy(i*step);
 
            final SpacecraftState state = propagator.propagate(date);
 
            // Attitude in LVLH
            final Attitude attLVLH = state.getAttitude(LOFType.LVLH);
            final double psiLVLH  = attLVLH.getRotation().getAngles(RotationOrder.ZYX)[0];
            final double tetaLVLH = attLVLH.getRotation().getAngles(RotationOrder.ZYX)[1];
 
            // Attitude in GCRF
            final Attitude attGCRF = state.getAttitude();
            final double psiGCRF  = attGCRF.getRotation().getAngles(RotationOrder.ZYX)[0];
            final double tetaGCRF = attGCRF.getRotation().getAngles(RotationOrder.ZYX)[1];
 
            // Direction of the Sun from the cdg of the satellite
            final Vector3D sunPos = sun.getPVCoordinates(date, FramesFactory.getGCRF()).getPosition();
            final Vector3D satPos = state.getPVCoordinates().getPosition();
            final Rotation sunDir = new Rotation(Vector3D.PLUS_I, sunPos.subtract(satPos));
 
            // Sun direction
            final double psiSun  = sunDir.getAngles(RotationOrder.ZYX)[0];
            final double tetaSun = sunDir.getAngles(RotationOrder.ZYX)[1];
 
            if ( (FastMath.abs(psiLVLH) < epsilon) || (FastMath.abs(tetaLVLH) < epsilon) ) {
                System.out.println(date+" => LVLH mode");
            } else {
                System.out.println(date+" => Sun pointing mode");
                System.out.println("    Delta Psi  = "+FastMath.toDegrees(psiSun-psiGCRF)+" deg");
                System.out.println("    Delta Teta = "+FastMath.toDegrees(tetaSun-tetaGCRF)+" deg");
            }
 
        }
 
    }
 
}