NumericalPropagationWithMass

public class NumericalPropagationWithMass {

    public static void main(String[] args) throws PatriusException, IOException, ParseException {
        
        // 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 = 0.01;
        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(0.);
        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);
        
//SPECIFIC
        // Mass model using an Assembly
        final AssemblyBuilder builder = new AssemblyBuilder();
        
        // Initial mass
        final double dryMass = 1000.;
        builder.addMainPart("MAIN");
        builder.addProperty(new MassProperty(dryMass), "MAIN");
        
        final Assembly assembly = builder.returnAssembly();
        final MassProvider mm = new MassModel(assembly);

        // We create a spacecratftstate
        final SpacecraftState iniState = new SpacecraftState(iniOrbit, mm);
//SPECIFIC
        
        // 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);

//SPECIFIC
        // Adding additional state (change name add to set for V3.3)
        propagator.setMassProviderEquation(mm);
//SPECIFIC
        
        // Forcing integration using cartesian equations
        propagator.setOrbitType(OrbitType.CARTESIAN);

        // Propagating 1000s
        final double dt = 1000.;
        final AbsoluteDate finalDate = date.shiftedBy(dt);
        final SpacecraftState finalState = propagator.propagate(finalDate);
        final Orbit finalOrbit = finalState.getOrbit();
        
        // Printing new date semi major axis
        System.out.println();
        System.out.println("Initial semi major axis = "+iniOrbit.getA()/1000.+" km");
        System.out.println("New date = "+finalOrbit.getDate().toString(TUC)+" deg");
        System.out.println("Final semi major axis = "+finalOrbit.getA()/1000.+" km");
        // Printing mass
        System.out.println();
        System.out.println("Mass = "+finalState.getMass("MAIN")+" kg");

    }
    
}