NumericalPropagationWithUsedDV 4.4 - Historique des versions

Admin le 4 octobre 2019 à 14:32

2019-10-04T14:32:07Z

← Version précédente		Version du 4 octobre 2019 à 14:32
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	// Initial orbit		// Initial orbit
	final double sma = 7200.e+3;		final double sma = 7200.e+3;
	final double ecc = 0.01;		final double ecc = 0.;
	final double inc = FastMath.toRadians(98.);		final double inc = FastMath.toRadians(98.);
	final double pa = FastMath.toRadians(0.);		final double pa = FastMath.toRadians(0.);
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	final KeplerianParameters par = new KeplerianParameters(sma, ecc, inc, pa, raan, anm, PositionAngle.MEAN, MU);		final KeplerianParameters par = new KeplerianParameters(sma, ecc, inc, pa, raan, anm, PositionAngle.MEAN, MU);
	final KeplerianOrbit iniOrbit = new KeplerianOrbit(par, GCRF, date0);		final KeplerianOrbit iniOrbit = new KeplerianOrbit(par, GCRF, date0);

	// Creating a mass model (see also specific example)		// Creating a mass model (see also specific example)
	final AssemblyBuilder builder = new AssemblyBuilder();		final AssemblyBuilder builder = new AssemblyBuilder();
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	// Forcing integration using cartesian equations		// Forcing integration using cartesian equations
	propagator.setOrbitType(OrbitType.CARTESIAN);		propagator.setOrbitType(OrbitType.CARTESIAN);

	final ArrayList<DateDetector> listOfEvents = new ArrayList<DateDetector>();		final ArrayList<DateDetector> listOfEvents = new ArrayList<DateDetector>();

Ligne 76 :		Ligne 76 :
	listOfEvents.add(eventImp);		listOfEvents.add(eventImp);
	// Creation of the impulsive maneuver		// Creation of the impulsive maneuver
	final Vector3D deltaV = new Vector3D(~~20.~~, 0., 0.);		final double dv = 20.;
			final Vector3D deltaV = new Vector3D(dv, 0., 0.);
	final ImpulseManeuver imp = new ImpulseManeuver(eventImp, deltaV, prop, mm, tank, LOFType.TNW);		final ImpulseManeuver imp = new ImpulseManeuver(eventImp, deltaV, prop, mm, tank, LOFType.TNW);

	// Duration of the maneuver to reach the initial semi major axis		// Duration of the maneuver to reach the initial semi major axis
	final double duration = 51.~~03781404091~~;		final double duration = 49.4933;
			System.out.println(duration);
	// Creation of the continuous thrust maneuver		// Creation of the continuous thrust maneuver
	final AbsoluteDate startDate = date0.shiftedBy(0.5*(iniOrbit.getKeplerianPeriod()-duration));		final AbsoluteDate startDate = date0.shiftedBy(iniOrbit.getKeplerianPeriod()-0.5*duration);
	final DateDetector eventStart = new DateDetector(startDate);		final DateDetector eventStart = new DateDetector(startDate);
	final DateDetector eventEnd = new DateDetector(startDate.shiftedBy(duration));		final DateDetector eventEnd = new DateDetector(startDate.shiftedBy(duration));
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	final AttitudeLaw attitudeLaw = new LofOffset(LOFType.TNW, RotationOrder.ZYX, 0., 0., 0.);		final AttitudeLaw attitudeLaw = new LofOffset(LOFType.TNW, RotationOrder.ZYX, 0., 0., 0.);
	propagator.setAttitudeProvider(attitudeLaw);		propagator.setAttitudeProvider(attitudeLaw);

	// Dt to get information just before/after an event		// Dt to get information just before/after an event
	final double dt = 1.e-6;		final double dt = 1.e-6;
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	}		}

	~~System.out.println("\nTHE END ...");~~
	~~final double dtf = iniOrbit.getKeplerianPeriod();~~
	~~final AbsoluteDate finalDate = date0.shiftedBy(dtf);~~
	~~final SpacecraftState finalState = propagator.propagate(finalDate);~~
	~~printResults(finalDate.toString(TUC), finalState, imp, man);~~

	}		}

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	System.out.println(" Ergols Mass = "+sc.getMass("TANK")+" kg");		System.out.println(" Ergols Mass = "+sc.getMass("TANK")+" kg");
	System.out.println(" Semi major axis = "+sc.getA()/1000.+" km");		System.out.println(" Semi major axis = "+sc.getA()/1000.+" km");
			System.out.println(" Eccentricity = "+sc.getE());
	}		}

	}		}
	</syntaxhighlight>		</syntaxhighlight>

Admin le 3 octobre 2019 à 13:09

2019-10-03T13:09:49Z

← Version précédente		Version du 3 octobre 2019 à 13:09
Ligne 30 :		Ligne 30 :
	final KeplerianParameters par = new KeplerianParameters(sma, ecc, inc, pa, raan, anm, PositionAngle.MEAN, MU);		final KeplerianParameters par = new KeplerianParameters(sma, ecc, inc, pa, raan, anm, PositionAngle.MEAN, MU);
	final KeplerianOrbit iniOrbit = new KeplerianOrbit(par, GCRF, date0);		final KeplerianOrbit iniOrbit = new KeplerianOrbit(par, GCRF, date0);

	~~//SPECIFIC~~
	// Creating a mass model (see also specific example)		// Creating a mass model (see also specific example)
	final AssemblyBuilder builder = new AssemblyBuilder();		final AssemblyBuilder builder = new AssemblyBuilder();
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	// We create a spacecratftstate		// We create a spacecratftstate
	final SpacecraftState iniState = new SpacecraftState(iniOrbit, mm);		final SpacecraftState iniState = new SpacecraftState(iniOrbit, mm);
	~~//SPECIFIC~~

	// Initialization of the Runge Kutta integrator with a 2 s step		// Initialization of the Runge Kutta integrator with a 2 s step
Ligne 71 :		Ligne 69 :
	// Forcing integration using cartesian equations		// Forcing integration using cartesian equations
	propagator.setOrbitType(OrbitType.CARTESIAN);		propagator.setOrbitType(OrbitType.CARTESIAN);

	~~//SPECIFIC~~

	final ArrayList<DateDetector> listOfEvents = new ArrayList<DateDetector>();		final ArrayList<DateDetector> listOfEvents = new ArrayList<DateDetector>();

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	final AttitudeLaw attitudeLaw = new LofOffset(LOFType.TNW, RotationOrder.ZYX, 0., 0., 0.);		final AttitudeLaw attitudeLaw = new LofOffset(LOFType.TNW, RotationOrder.ZYX, 0., 0., 0.);
	propagator.setAttitudeProvider(attitudeLaw);		propagator.setAttitudeProvider(attitudeLaw);

	~~//SPECIFIC~~

	// Dt to get information just before/after an event		// Dt to get information just before/after an event

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

2019-10-03T13:09:08Z

Page créée avec « <syntaxhighlight lang="java"> public class NumericalPropagationWithUsedDV { public static void main(String[] args) throws PatriusException { Locale.s... »

Nouvelle page

<syntaxhighlight lang="java">
public class NumericalPropagationWithUsedDV {

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 = 0.01;
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);

//SPECIFIC
// 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);
//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);

// Forcing integration using cartesian equations
propagator.setOrbitType(OrbitType.CARTESIAN);

//SPECIFIC

final ArrayList<DateDetector> listOfEvents = new ArrayList<DateDetector>();

// Event corresponding to the criteria to trigger the impulsive maneuver
final DateDetector eventImp = new DateDetector(date0.shiftedBy(10.));
listOfEvents.add(eventImp);
// Creation of the impulsive maneuver
final Vector3D deltaV = new Vector3D(20., 0., 0.);
final ImpulseManeuver imp = new ImpulseManeuver(eventImp, deltaV, prop, mm, tank, LOFType.TNW);

// Duration of the maneuver to reach the initial semi major axis
final double duration = 51.03781404091;
// Creation of the continuous thrust maneuver
final AbsoluteDate startDate = date0.shiftedBy(0.5*(iniOrbit.getKeplerianPeriod()-duration));
final DateDetector eventStart = new DateDetector(startDate);
final DateDetector eventEnd = new DateDetector(startDate.shiftedBy(duration));
listOfEvents.add(eventStart);
listOfEvents.add(eventEnd);
final Vector3D direction = new Vector3D(-1., 0., 0.);
final ContinuousThrustManeuver man = new ContinuousThrustManeuver(eventStart, eventEnd, prop, direction, mm, tank);

// Creation of the sequence of maneuver
ManeuversSequence seq = new ManeuversSequence(0., 0.);
seq.add(imp);
seq.add(man);

// Adding the maneuver sequence to the propagator
seq.applyTo(propagator);
// Adding additional state
propagator.setMassProviderEquation(mm);

// Adding an attitude law (or attitude sequence : mandatory)
final AttitudeLaw attitudeLaw = new LofOffset(LOFType.TNW, RotationOrder.ZYX, 0., 0., 0.);
propagator.setAttitudeProvider(attitudeLaw);

//SPECIFIC

// Dt to get information just before/after an event
final double dt = 1.e-6;

for (int i = 0; i < listOfEvents.size(); i++) {

System.out.println("\nEVENT #"+i);

System.out.println("Before ...");
final AbsoluteDate dateBefore = listOfEvents.get(i).getDate().shiftedBy(-dt);
final SpacecraftState finalStateBefore = propagator.propagate(dateBefore);
printResults(dateBefore.toString(TUC), finalStateBefore, imp, man);

System.out.println("After ...");
final AbsoluteDate dateAfter = listOfEvents.get(i).getDate().shiftedBy(dt);
final SpacecraftState finalStateAfter = propagator.propagate(dateAfter);
printResults(dateAfter.toString(TUC), finalStateAfter, imp, man);
}

System.out.println("\nTHE END ...");
final double dtf = iniOrbit.getKeplerianPeriod();
final AbsoluteDate finalDate = date0.shiftedBy(dtf);
final SpacecraftState finalState = propagator.propagate(finalDate);
printResults(finalDate.toString(TUC), finalState, imp, man);

}

private static void printResults ( final String sdate, final SpacecraftState sc,
final ImpulseManeuver imp, final ContinuousThrustManeuver man ) throws PatriusException {
System.out.println(" Date = "+sdate);
System.out.println(" Impulsive Maneuver = "+imp.getUsedDV()+" m/s");
System.out.println(" Continuous Maneuver = "+man.getUsedDV()+" m/s");
System.out.println(" Ergols Mass = "+sc.getMass("TANK")+" kg");
System.out.println(" Semi major axis = "+sc.getA()/1000.+" km");
}

}
</syntaxhighlight>

NumericalPropagationWithUsedDV 4.4 - Historique des versions

Admin le 4 octobre 2019 à 14:32

Admin le 3 octobre 2019 à 13:09

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

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