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	<syntaxhighlight lang="java">		<syntaxhighlight lang="java">
	/**
	* Example allowing to numerically propagate an orbit with a SRP model force.
	* @author goesterjf
	*
	*/
	public class NumericalPropagationWithSRP {		public class NumericalPropagationWithSRP {

Admin le 21 décembre 2018 à 09:54

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			<syntaxhighlight lang="java">
	/**		/**
	* Example allowing to numerically propagate an orbit with a SRP model force.		* Example allowing to numerically propagate an orbit with a SRP model force.
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	}		}
			</syntaxhighlight>

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/**
* Example allowing to numerically propagate an orbit with a SRP model force.
* @author goesterjf
*
*/
public class NumericalPropagationWithSRP {

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 = 7000.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
final double ka = 1.0;
final double ks = 0.0;
final double kd = 0.0;
builder.addProperty(new RadiativeProperty(ka, ks, kd), "MAIN");
builder.addProperty(new RadiativeIRProperty(ka, ks, kd), "MAIN");
final double radius = 10.;
builder.addProperty(new RadiativeSphereProperty(radius), "MAIN");
//SPECIFIC

final Assembly assembly = builder.returnAssembly();

//SPECIFIC
// Sun ephemeris
CelestialBody sun = new MeeusSun();

// Definition of the Earth ellipsoid for later SRP computation
final Frame ITRF = FramesFactory.getITRF();
final double AE = Constants.WGS84_EARTH_EQUATORIAL_RADIUS;
final GeometricBodyShape EARTH = new ExtendedOneAxisEllipsoid(AE, Constants.WGS84_EARTH_FLATTENING, ITRF, "EARTH");

// Direct SRP data
final double dRef = 1.4959787E11;
final double pRef = 4.5605E-6;
DirectRadiativeModel rm = new DirectRadiativeModel(assembly);
SolarRadiationPressureEllipsoid radPres = new SolarRadiationPressureEllipsoid(dRef, pRef, sun, EARTH, rm);

// Rediffused DRP data
final int inCorona = 1;
final int inMeridian = 10;
final IEmissivityModel inEmissivityModel = new KnockeRiesModel();
final boolean inAlbedo = true;
final boolean inIr = true;
final double coefAlbedo = 1.;
final double coefIr = 1.;
RediffusedRadiativeModel rdm = new RediffusedRadiativeModel(inAlbedo, inIr, coefAlbedo, coefIr, assembly);
RediffusedRadiationPressure reDiff = new RediffusedRadiationPressure(sun, GCRF, inCorona, inMeridian, inEmissivityModel, rdm);
//SPECIFIC

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 SRP forces
propagator.addForceModel(radPres);
propagator.addForceModel(reDiff);
//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");

}

}

NumericalPropagationWithSRP - Historique des versions

Admin le 21 décembre 2018 à 09:54

Admin le 21 décembre 2018 à 09:54

Admin : Page créée avec « /** * Example allowing to numerically propagate an orbit with a SRP model force. * @author goesterjf * */ public class NumericalPropagationWithSRP { public static... »

Admin : Page créée avec « /** * Example allowing to numerically propagate an orbit with a SRP model force. * @author goesterjf * */ public class NumericalPropagationWithSRP { public static... »