NumericalPropagationWithLiftAndDrag

NumericalPropagationWithLiftAndDrag

public class NumericalPropagationWithLiftAndDrag {
 
    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 = 6600.e+3;
        final double exc = 0.;
        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);
 
        // Mass model using an Assembly
 
        final AssemblyBuilder builder = new AssemblyBuilder();
 
        // Initial mass (mandatory to take into account mass for atmospheric force computation)
        final double dryMass = 100.;
        builder.addMainPart("MAIN");
        builder.addProperty(new MassProperty(dryMass), "MAIN");
 
//SPECIFIC
        // Adding the AeroGSphere property for drag only
        final double cd = 2.0;
        final double cl = 0.2;
        final double sref = 10.;        
        //builder.addProperty(new AeroGlobalProperty(cd, cl, new ConstantFunction(sref)), "MAIN");
        builder.addProperty(new AeroGlobalProperty(cd, cl, new Sphere(Vector3D.ZERO, FastMath.sqrt(sref/FastMath.PI))), "MAIN");
 
        final UpdatableFrame mainFrame = new UpdatableFrame(GCRF, Transform.IDENTITY, "mainPartFrame"); 
        builder.initMainPartFrame(mainFrame);
//SPECIFIC
 
        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);
 
        // Adding additional state (change name add to set for V3.3)
        propagator.setMassProviderEquation(mm);
 
        // Forcing integration using cartesian equations
        propagator.setOrbitType(OrbitType.CARTESIAN);
 
//SPECIFIC
        // Adding an attitude law
        final AttitudeLaw attitudeLaw = new LofOffset(LOFType.LVLH, RotationOrder.ZYX, 0., 0., 0.);
        propagator.setAttitudeProvider(attitudeLaw);
 
        // Definition of the Earth ellipsoid for later atmospheric density computation
        final Frame ITRF = FramesFactory.getITRF();
        final double AE = Constants.WGS84_EARTH_EQUATORIAL_RADIUS;
        final BodyShape EARTH = new OneAxisEllipsoid(AE, Constants.WGS84_EARTH_FLATTENING, ITRF, "EARTH");
 
        // Adding atmospheric forces using US76 model
        final Atmosphere atmosphere = new US76(EARTH);
        final DragLiftModel dragLiftModel = new DragLiftModel(assembly);
        final ForceModel atm = new DragForce(atmosphere, dragLiftModel);
        propagator.addForceModel(atm);
//SPECIFIC
 
        // Propagating 5 periods
        final double dt = 5.*iniOrbit.getKeplerianPeriod();
        final AbsoluteDate finalDate = date.shiftedBy(dt);
        final SpacecraftState finalState = propagator.propagate(finalDate);
        final Orbit finalOrbit = finalState.getOrbit();
 
        // Printing new date and 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");
 
    }
 
}