NumericalPropagationWithOrbitalIncrementManeuvers 4.4
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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"); } } } }
