NumericalPropagationWithOrbitalIncrementManeuvers 4.4 : Différence entre versions

public class NumericalPropagationWithOrbitalIncrementManeuvers {
 
    public static void main(String[] args) throws PatriusException {
 
        Locale.setDefault(Locale.US);
 
        // 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 date0 = 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 ecc = 0.;
        final double inc = FastMath.toRadians(98.);
        final double pa = FastMath.toRadians(0.);
        final double raan = FastMath.toRadians(0.);
        final double anm = FastMath.toRadians(0.);
        final double MU = Constants.WGS84_EARTH_MU;
 
        final KeplerianParameters par = new KeplerianParameters(sma, ecc, inc, pa, raan, anm, PositionAngle.MEAN, MU);
        final KeplerianOrbit iniOrbit = new KeplerianOrbit(par, GCRF, date0);
        final double period0 = iniOrbit.getKeplerianPeriod();
        final double sma0 = iniOrbit.getA();
 
        // Creating a mass model (see also specific example)
        final AssemblyBuilder builder = new AssemblyBuilder();
 
        // Main part
        final double iniMass = 900.;
        builder.addMainPart("MAIN");
        builder.addProperty(new MassProperty(iniMass), "MAIN");
 
        // Tank part (ergols mass)
        final double ergolsMass = 100.;
        final TankProperty tank = new TankProperty(ergolsMass);
        builder.addPart("TANK", "MAIN", Transform.IDENTITY);
        builder.addProperty(tank, "TANK");
 
        // Engine part
        final double isp = 300.;
        final double thrust = 400.;
        final PropulsiveProperty prop = new PropulsiveProperty(thrust, isp); // au lieu de new PropulsiveProperty("PROP", thrust, isp);
 
        builder.addPart("PROP", "MAIN", Transform.IDENTITY);
        builder.addProperty(prop, "PROP");
 
        final Assembly assembly = builder.returnAssembly();
        final MassProvider mm = new MassModel(assembly);
 
        // We create a spacecratftstate
        final SpacecraftState iniState = new SpacecraftState(iniOrbit, mm);
 
        // 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);
 
        final ArrayList<DateDetector> listOfEvents = new ArrayList<DateDetector>();
        final ArrayList<ImpulseManeuver> listOfMan = new ArrayList<ImpulseManeuver>();
        final ArrayList<String> listOfManLabels = new ArrayList<String>();
        final boolean error = false;
 
        // Event corresponding to the criteria to trigger the Da impulsive maneuver
        double da = 10.e3;
        final DateDetector eventDa = new DateDetector(date0);
        listOfEvents.add(eventDa);
        final ImpulseManeuver impDa = new ImpulseDaManeuver(eventDa, da, prop, mm, tank);
        listOfMan.add(impDa);
        listOfManLabels.add("Da impulsive maneuver");
        final double newa = sma0 + da;
        final double newPeriod = 2.*FastMath.PI*FastMath.sqrt(newa*newa*newa/MU);
 
        // Event corresponding to the criteria to trigger the Da/De impulsive maneuver ... but an impossible one !
        da = -10000.;
        double de = -0.001386962552;
        final DateDetector eventDaDe1 = new DateDetector(date0.shiftedBy(0.5*newPeriod));
        listOfEvents.add(eventDaDe1);
        final ImpulseManeuver impDaDe1 = new ImpulseDeManeuver(eventDaDe1, de, da, prop, mm, tank, error);
        listOfMan.add(impDaDe1);
        listOfManLabels.add("DaDe impossible impulsive maneuver");
 
        // Event corresponding to the criteria to trigger the Da/De impulsive maneuver
        da = -10000.;
        de = -0.001386962552;
        final DateDetector eventDaDe2 = new DateDetector(date0.shiftedBy(newPeriod));
        listOfEvents.add(eventDaDe2);
        final ImpulseManeuver impDaDe2 = new ImpulseDeManeuver(eventDaDe2, de, da, prop, mm, tank, error);
        listOfMan.add(impDaDe2);
        listOfManLabels.add("DaDe impulsive maneuver");
 
        // Event corresponding to the criteria to trigger the Di impulsive maneuver
        final DateDetector eventDi = new DateDetector(date0.shiftedBy(newPeriod+period0));
        listOfEvents.add(eventDi);
        double di = FastMath.toRadians(0.1);
        final ImpulseManeuver impDi = new ImpulseDiManeuver(eventDi, di, prop, mm, tank, error);
        //final ImpulseManeuver impDi = new ImpulseDiManeuver(eventDi, di, 0., prop, mm, tank, error);
        listOfMan.add(impDi);
        listOfManLabels.add("Di impulsive maneuver");
 
        // Event corresponding to the standard criteria
        final DateDetector eventStd = new DateDetector(date0.shiftedBy(newPeriod+2.*period0));
        listOfEvents.add(eventStd);
        final Vector3D deltaV = new Vector3D(10., 0., 0.);
        final ImpulseManeuver impStd = new ImpulseManeuver(eventStd, deltaV, prop, mm, tank, LOFType.TNW);
        listOfMan.add(impStd);
        listOfManLabels.add("Std impulsive maneuver");
 
        // Creation of the sequence of maneuver
        ManeuversSequence seq = new ManeuversSequence(0., 0.);
        for (ImpulseManeuver impulseManeuver : listOfMan) {
            seq.add(impulseManeuver);
        }
 
        // Adding the maneuver sequence to the propagator
        seq.applyTo(propagator);
        // Adding additional state
        propagator.setMassProviderEquation(mm);
 
        printResults(iniOrbit, listOfMan, "Initial conditions", false);
 
        // For propagating just after maneuvers
        final double dt = 1.e-6;
 
        for (int i = 0; i < listOfMan.size(); i++) {
            final SpacecraftState finalState = propagator.propagate(listOfEvents.get(i).getDate().shiftedBy(dt));
            KeplerianOrbit kep = new KeplerianOrbit(finalState.getOrbit());
            printResults(kep, listOfMan, listOfManLabels.get(i), true);
 
        }
 
    }
 
    private static void printResults (final KeplerianOrbit kep, final ArrayList<ImpulseManeuver> listOfMan,
            final String manLabel, final boolean isUsedDeltaV ) throws PatriusException {
 
        String str = "";
        if ( !isUsedDeltaV ) {
            str = "Initial ";
        }
 
        System.out.println();
        System.out.println(manLabel);
        System.out.println("Initial date = "+kep.getDate().toString(TimeScalesFactory.getUTC()));
        System.out.println(String.format("%sSemi major axis = %8.3f km", str, kep.getA()/1000.));
        System.out.println(String.format("%seccentricity    =    %8.6f", str, kep.getE()));
        System.out.println(String.format("%sinclination     = %8.3f deg", str, FastMath.toDegrees(kep.getI())));
        System.out.println(String.format("%sAOL             = %8.3f deg", str, FastMath.toDegrees(kep.getPerigeeArgument()+kep.getTrueAnomaly())));
        for (int i = 0; i < listOfMan.size(); i++) {
            if ( isUsedDeltaV ) {
                System.out.println("Maneuver #"+i+" = "+listOfMan.get(i).getUsedDV() + " m/s");
            } else {
                System.out.println("Maneuver #"+i+" = "+listOfMan.get(i).getDeltaVSat() + " m/s");
            }
        }
 
    }
 
}