NumericalPropagationWithCustomEvent 4.5.1 : Différence entre versions
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Version actuelle en date du 17 août 2020 à 09:05
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); } }