NumericalPropagationWithCustomEvent 4.4

NumericalPropagationWithCustomEvent 4.4

public class NumericalPropagationWithCustomEvent {
 
    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 date = 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 = 0.02;
        final double per = sma*(1.-exc);
        final double apo = sma*(1.+exc);
        final double inc = FastMath.toRadians(98.);
        final double pa = FastMath.toRadians(0.);
        final double raan = FastMath.toRadians(0.);
        final double anm = FastMath.toRadians(180.);
        final double MU = Constants.WGS84_EARTH_MU;
 
        final ApsisRadiusParameters par = new ApsisRadiusParameters(per, apo, inc, pa, raan, anm, PositionAngle.MEAN, MU);
        final Orbit iniOrbit = new ApsisOrbit(par, GCRF, date);
 
        // 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);
 
//SPECIFIC
        // Definition of the custom event
        EventDetector event = new EventDetector() {
 
            private static final long serialVersionUID = 1L;
            public double g(SpacecraftState s) throws PatriusException {
                // We want to raise the event when Lv = 45 deg
                final double delta = s.getLv() - FastMath.toRadians(45.);
                return delta;
            }
 
            public Action eventOccurred(SpacecraftState s, boolean increasing,
                    boolean forward) throws PatriusException {
                System.out.println("Event occured at date : "+s.getDate().toString(TUC)+" (LM = "+FastMath.toDegrees(s.getLv())+")");
                return Action.CONTINUE;
            }
 
            public boolean shouldBeRemoved() {
                return false;
            }
            public SpacecraftState resetState(SpacecraftState oldState)
                    throws PatriusException {
                return null;
            }
            public void init(SpacecraftState s0, AbsoluteDate t) {
            }
            public double getThreshold() {
                return AbstractDetector.DEFAULT_THRESHOLD;
            }
            public int getSlopeSelection() {
                return 0;
            }
            public int getMaxIterationCount() {
                return 20;
            }
            public double getMaxCheckInterval() {
                return AbstractDetector.DEFAULT_MAXCHECK;
            }
 
            @Override
            public EventDetector copy() {
                return null;
            }
 
        };
 
        // Adding the event to the propagator
        propagator.addEventDetector(event);
//SPECIFIC
 
        // Propagating on several orbits
        final double dt = 5.*iniOrbit.getKeplerianPeriod();
        final AbsoluteDate finalDate = date.shiftedBy(dt);
        propagator.propagate(finalDate);
 
    }
 
}