Examples with PropagatorConverter org.orekit.propagation.conversion.PropagatorConverter used on opensource projects

Example 1 with PropagatorConverter

use of org.orekit.propagation.conversion.PropagatorConverter in project Orekit by CS-SI.

the class EcksteinHechlerPropagatorTest method testInitializationCorrectness.

@Test
public void testInitializationCorrectness() throws OrekitException, IOException {
    // Definition of initial conditions
    AbsoluteDate date = AbsoluteDate.J2000_EPOCH.shiftedBy(154.);
    Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
    Frame eme2000 = FramesFactory.getEME2000();
    Vector3D pole = itrf.getTransformTo(eme2000, date).transformVector(Vector3D.PLUS_K);
    Frame poleAligned = new Frame(FramesFactory.getEME2000(), new Transform(date, new Rotation(pole, Vector3D.PLUS_K)), "pole aligned", true);
    CircularOrbit initial = new CircularOrbit(7208669.8179538045, 1.3740461966386876E-4, -3.2364250248363356E-5, FastMath.toRadians(97.40236024565775), FastMath.toRadians(166.15873160992115), FastMath.toRadians(90.1282370098961), PositionAngle.MEAN, poleAligned, date, provider.getMu());
    // find the default Eckstein-Hechler propagator initialized from the initial orbit
    EcksteinHechlerPropagator defaultEH = new EcksteinHechlerPropagator(initial, provider);
    // the osculating parameters recomputed by the default Eckstein-Hechler propagator are quite different
    // from initial orbit
    CircularOrbit defaultOrbit = (CircularOrbit) OrbitType.CIRCULAR.convertType(defaultEH.propagateOrbit(initial.getDate()));
    Assert.assertEquals(267.4, defaultOrbit.getA() - initial.getA(), 0.1);
    // the position on the other hand match perfectly
    Assert.assertEquals(0.0, Vector3D.distance(defaultOrbit.getPVCoordinates().getPosition(), initial.getPVCoordinates().getPosition()), 1.0e-8);
    // set up a reference numerical propagator starting for the specified start orbit
    // using the same force models (i.e. the first few zonal terms)
    double[][] tol = NumericalPropagator.tolerances(0.1, initial, OrbitType.CIRCULAR);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator.setInitialStepSize(60);
    NumericalPropagator num = new NumericalPropagator(integrator);
    num.addForceModel(new HolmesFeatherstoneAttractionModel(itrf, GravityFieldFactory.getNormalizedProvider(provider)));
    num.setInitialState(new SpacecraftState(initial));
    num.setOrbitType(OrbitType.CIRCULAR);
    // find the best Eckstein-Hechler propagator that match the orbit evolution
    PropagatorConverter converter = new FiniteDifferencePropagatorConverter(new EcksteinHechlerPropagatorBuilder(initial, provider, PositionAngle.TRUE, 1.0), 1.0e-6, 100);
    EcksteinHechlerPropagator fittedEH = (EcksteinHechlerPropagator) converter.convert(num, 3 * initial.getKeplerianPeriod(), 300);
    // the default Eckstein-Hechler propagator did however quite a good job, as it found
    // an orbit close to the best fitting
    CircularOrbit fittedOrbit = (CircularOrbit) OrbitType.CIRCULAR.convertType(fittedEH.propagateOrbit(initial.getDate()));
    Assert.assertEquals(0.623, defaultOrbit.getA() - fittedOrbit.getA(), 0.1);
    // the position on the other hand are slightly different
    // because the fitted orbit minimizes the residuals over a complete time span,
    // not on a single point
    Assert.assertEquals(58.0, Vector3D.distance(defaultOrbit.getPVCoordinates().getPosition(), fittedOrbit.getPVCoordinates().getPosition()), 0.1);
}

Example 2 with PropagatorConverter

use of org.orekit.propagation.conversion.PropagatorConverter in project Orekit by CS-SI.

the class PropagatorConversion method main.

/**
 * Program entry point.
 * @param args program arguments (unused here)
 */
public static void main(String[] args) {
    try {
        // configure Orekit
        File home = new File(System.getProperty("user.home"));
        File orekitData = new File(home, "orekit-data");
        if (!orekitData.exists()) {
            System.err.format(Locale.US, "Failed to find %s folder%n", orekitData.getAbsolutePath());
            System.err.format(Locale.US, "You need to download %s from the %s page and unzip it in %s for this tutorial to work%n", "orekit-data.zip", "https://www.orekit.org/forge/projects/orekit/files", home.getAbsolutePath());
            System.exit(1);
        }
        DataProvidersManager manager = DataProvidersManager.getInstance();
        manager.addProvider(new DirectoryCrawler(orekitData));
        // gravity field
        NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(2, 0);
        double mu = provider.getMu();
        // inertial frame
        Frame inertialFrame = FramesFactory.getEME2000();
        // Initial date
        AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, TimeScalesFactory.getUTC());
        // Initial orbit (GTO)
        // semi major axis in meters
        final double a = 24396159;
        // eccentricity
        final double e = 0.72831215;
        // inclination
        final double i = FastMath.toRadians(7);
        // perigee argument
        final double omega = FastMath.toRadians(180);
        // right ascention of ascending node
        final double raan = FastMath.toRadians(261);
        // mean anomaly
        final double lM = 0;
        Orbit initialOrbit = new KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN, inertialFrame, initialDate, mu);
        final double period = initialOrbit.getKeplerianPeriod();
        // Initial state definition
        final SpacecraftState initialState = new SpacecraftState(initialOrbit);
        // Adaptive step integrator with a minimum step of 0.001 and a maximum step of 1000
        final double minStep = 0.001;
        final double maxStep = 1000.;
        final double dP = 1.e-2;
        final OrbitType orbType = OrbitType.CARTESIAN;
        final double[][] tol = NumericalPropagator.tolerances(dP, initialOrbit, orbType);
        final AbstractIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tol[0], tol[1]);
        // Propagator
        NumericalPropagator numProp = new NumericalPropagator(integrator);
        numProp.setInitialState(initialState);
        numProp.setOrbitType(orbType);
        // Force Models:
        // 1 - Perturbing gravity field (only J2 is considered here)
        ForceModel gravity = new HolmesFeatherstoneAttractionModel(FramesFactory.getITRF(IERSConventions.IERS_2010, true), provider);
        // Add force models to the propagator
        numProp.addForceModel(gravity);
        // Propagator factory
        PropagatorBuilder builder = new KeplerianPropagatorBuilder(initialOrbit, PositionAngle.TRUE, dP);
        // Propagator converter
        PropagatorConverter fitter = new FiniteDifferencePropagatorConverter(builder, 1.e-6, 5000);
        // Resulting propagator
        KeplerianPropagator kepProp = (KeplerianPropagator) fitter.convert(numProp, 2 * period, 251);
        // Step handlers
        StatesHandler numStepHandler = new StatesHandler();
        StatesHandler kepStepHandler = new StatesHandler();
        // Set up operating mode for the propagator as master mode
        // with fixed step and specialized step handler
        numProp.setMasterMode(60., numStepHandler);
        kepProp.setMasterMode(60., kepStepHandler);
        // Extrapolate from the initial to the final date
        numProp.propagate(initialDate.shiftedBy(10. * period));
        kepProp.propagate(initialDate.shiftedBy(10. * period));
        // retrieve the states
        List<SpacecraftState> numStates = numStepHandler.getStates();
        List<SpacecraftState> kepStates = kepStepHandler.getStates();
        // Print the results on the output file
        File output = new File(new File(System.getProperty("user.home")), "elements.dat");
        try (final PrintStream stream = new PrintStream(output, "UTF-8")) {
            stream.println("# date Anum Akep Enum Ekep Inum Ikep LMnum LMkep");
            for (SpacecraftState numState : numStates) {
                for (SpacecraftState kepState : kepStates) {
                    if (numState.getDate().compareTo(kepState.getDate()) == 0) {
                        stream.println(numState.getDate() + " " + numState.getA() + " " + kepState.getA() + " " + numState.getE() + " " + kepState.getE() + " " + FastMath.toDegrees(numState.getI()) + " " + FastMath.toDegrees(kepState.getI()) + " " + FastMath.toDegrees(MathUtils.normalizeAngle(numState.getLM(), FastMath.PI)) + " " + FastMath.toDegrees(MathUtils.normalizeAngle(kepState.getLM(), FastMath.PI)));
                        break;
                    }
                }
            }
        }
        System.out.println("Results saved as file " + output);
        File output1 = new File(new File(System.getProperty("user.home")), "elts_pv.dat");
        try (final PrintStream stream = new PrintStream(output1, "UTF-8")) {
            stream.println("# date pxn pyn pzn vxn vyn vzn pxk pyk pzk vxk vyk vzk");
            for (SpacecraftState numState : numStates) {
                for (SpacecraftState kepState : kepStates) {
                    if (numState.getDate().compareTo(kepState.getDate()) == 0) {
                        final double pxn = numState.getPVCoordinates().getPosition().getX();
                        final double pyn = numState.getPVCoordinates().getPosition().getY();
                        final double pzn = numState.getPVCoordinates().getPosition().getZ();
                        final double vxn = numState.getPVCoordinates().getVelocity().getX();
                        final double vyn = numState.getPVCoordinates().getVelocity().getY();
                        final double vzn = numState.getPVCoordinates().getVelocity().getZ();
                        final double pxk = kepState.getPVCoordinates().getPosition().getX();
                        final double pyk = kepState.getPVCoordinates().getPosition().getY();
                        final double pzk = kepState.getPVCoordinates().getPosition().getZ();
                        final double vxk = kepState.getPVCoordinates().getVelocity().getX();
                        final double vyk = kepState.getPVCoordinates().getVelocity().getY();
                        final double vzk = kepState.getPVCoordinates().getVelocity().getZ();
                        stream.println(numState.getDate() + " " + pxn + " " + pyn + " " + pzn + " " + vxn + " " + vyn + " " + vzn + " " + pxk + " " + pyk + " " + pzk + " " + vxk + " " + vyk + " " + vzk);
                        break;
                    }
                }
            }
        }
        System.out.println("Results saved as file " + output1);
    } catch (OrekitException oe) {
        System.err.println(oe.getLocalizedMessage());
        System.exit(1);
    } catch (IOException ioe) {
        System.err.println(ioe.getLocalizedMessage());
        System.exit(1);
    }
}