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	// Adding additional state		// Adding additional state
	propagator.setMassProviderEquation(mm);		propagator.setMassProviderEquation(mm);

	~~//SPECIFIC~~

	printResults(iniOrbit, listOfMan, "Initial conditions", false);		printResults(iniOrbit, listOfMan, "Initial conditions", false);

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2019-10-03T13:05:33Z

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<syntaxhighlight lang="java">
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);

//SPECIFIC

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");
}
}

}

}
</syntaxhighlight>

NumericalPropagationWithOrbitalIncrementManeuvers 4.4 - Historique des versions

Admin le 3 octobre 2019 à 13:06

Admin : Page créée avec « public class NumericalPropagationWithOrbitalIncrementManeuvers { public static void main(String[] args) throws PatriusException {... »

Admin : Page créée avec « public class NumericalPropagationWithOrbitalIncrementManeuvers { public static void main(String[] args) throws PatriusException {... »