Examples with OrekitFixedStepHandler org.orekit.propagation.sampling.OrekitFixedStepHandler used on opensource projects

Example 16 with OrekitFixedStepHandler

use of org.orekit.propagation.sampling.OrekitFixedStepHandler in project Orekit by CS-SI.

the class Frames3 method main.

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));
        // Initial state definition :
        // ==========================
        // Date
        // ****
        AbsoluteDate initialDate = new AbsoluteDate(2003, 4, 7, 10, 55, 21.575, TimeScalesFactory.getUTC());
        // Orbit
        // *****
        // The Sun is in the orbital plane for raan ~ 202
        // gravitation coefficient
        double mu = 3.986004415e+14;
        // inertial frame
        Frame eme2000 = FramesFactory.getEME2000();
        Orbit orbit = new CircularOrbit(7178000.0, 0.5e-4, -0.5e-4, FastMath.toRadians(50.), FastMath.toRadians(220.), FastMath.toRadians(5.300), PositionAngle.MEAN, eme2000, initialDate, mu);
        // Attitude laws
        // *************
        // Earth
        Frame earthFrame = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
        BodyShape earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, earthFrame);
        // Target pointing attitude provider over satellite nadir at date, without yaw compensation
        NadirPointing nadirLaw = new NadirPointing(eme2000, earth);
        // Target pointing attitude provider with yaw compensation
        final PVCoordinatesProvider sun = CelestialBodyFactory.getSun();
        YawSteering yawSteeringLaw = new YawSteering(eme2000, nadirLaw, sun, Vector3D.MINUS_I);
        // Propagator : Eckstein-Hechler analytic propagator
        Propagator propagator = new EcksteinHechlerPropagator(orbit, yawSteeringLaw, Constants.EIGEN5C_EARTH_EQUATORIAL_RADIUS, Constants.EIGEN5C_EARTH_MU, Constants.EIGEN5C_EARTH_C20, Constants.EIGEN5C_EARTH_C30, Constants.EIGEN5C_EARTH_C40, Constants.EIGEN5C_EARTH_C50, Constants.EIGEN5C_EARTH_C60);
        // Let's write the results in a file in order to draw some plots.
        propagator.setMasterMode(10, new OrekitFixedStepHandler() {

            PrintStream out = null;

            public void init(SpacecraftState s0, AbsoluteDate t, double step) throws OrekitException {
                try {
                    File file = new File(System.getProperty("user.home"), "XYZ.dat");
                    System.out.println("Results written to file: " + file.getAbsolutePath());
                    out = new PrintStream(file, "UTF-8");
                    out.println("#time X Y Z Wx Wy Wz");
                } catch (IOException ioe) {
                    throw new OrekitException(ioe, LocalizedCoreFormats.SIMPLE_MESSAGE, ioe.getLocalizedMessage());
                }
            }

            public void handleStep(SpacecraftState currentState, boolean isLast) throws OrekitException {
                // get the transform from orbit/attitude reference frame to spacecraft frame
                Transform inertToSpacecraft = currentState.toTransform();
                // get the position of the Sun in orbit/attitude reference frame
                Vector3D sunInert = sun.getPVCoordinates(currentState.getDate(), currentState.getFrame()).getPosition();
                // convert Sun position to spacecraft frame
                Vector3D sunSat = inertToSpacecraft.transformPosition(sunInert);
                // and the spacecraft rotational rate also
                Vector3D spin = inertToSpacecraft.getRotationRate();
                // Lets calculate the reduced coordinates
                double sunX = sunSat.getX() / sunSat.getNorm();
                double sunY = sunSat.getY() / sunSat.getNorm();
                double sunZ = sunSat.getZ() / sunSat.getNorm();
                out.format(Locale.US, "%s %12.3f %12.3f %12.3f %12.7f %12.7f %12.7f%n", currentState.getDate(), sunX, sunY, sunZ, spin.getX(), spin.getY(), spin.getZ());
                if (isLast) {
                    out.close();
                }
            }
        });
        System.out.println("Running...");
        propagator.propagate(initialDate.shiftedBy(6000));
    } catch (OrekitException oe) {
        System.err.println(oe.getMessage());
    }
}