Examples with DataProvidersManager org.orekit.data.DataProvidersManager used on opensource projects

Example 16 with DataProvidersManager

use of org.orekit.data.DataProvidersManager in project Orekit by CS-SI.

the class Frames1 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, orbit
        TimeScale utc = TimeScalesFactory.getUTC();
        AbsoluteDate initialDate = new AbsoluteDate(2008, 10, 01, 0, 0, 00.000, utc);
        // gravitation coefficient
        double mu = 3.986004415e+14;
        // inertial frame for orbit definition
        Frame inertialFrame = FramesFactory.getEME2000();
        Vector3D posisat = new Vector3D(-6142438.668, 3492467.560, -25767.25680);
        Vector3D velosat = new Vector3D(505.8479685, 942.7809215, 7435.922231);
        PVCoordinates pvsat = new PVCoordinates(posisat, velosat);
        Orbit initialOrbit = new CartesianOrbit(pvsat, inertialFrame, initialDate, mu);
        // Propagator : consider a simple Keplerian motion
        Propagator kepler = new KeplerianPropagator(initialOrbit);
        // Earth and frame
        Frame earthFrame = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
        BodyShape earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, earthFrame);
        // Station
        final double longitude = FastMath.toRadians(45.);
        final double latitude = FastMath.toRadians(25.);
        final double altitude = 0.;
        final GeodeticPoint station = new GeodeticPoint(latitude, longitude, altitude);
        final TopocentricFrame staF = new TopocentricFrame(earth, station, "station");
        System.out.println("          time           doppler (m/s)");
        // Stop date
        final AbsoluteDate finalDate = new AbsoluteDate(initialDate, 6000, utc);
        // Loop
        AbsoluteDate extrapDate = initialDate;
        while (extrapDate.compareTo(finalDate) <= 0) {
            // We can simply get the position and velocity of spacecraft in station frame at any time
            PVCoordinates pvInert = kepler.propagate(extrapDate).getPVCoordinates();
            PVCoordinates pvStation = inertialFrame.getTransformTo(staF, extrapDate).transformPVCoordinates(pvInert);
            // And then calculate the doppler signal
            double doppler = Vector3D.dotProduct(pvStation.getPosition(), pvStation.getVelocity()) / pvStation.getPosition().getNorm();
            System.out.format(Locale.US, "%s   %9.3f%n", extrapDate, doppler);
            extrapDate = extrapDate.shiftedBy(600);
        }
    } catch (OrekitException oe) {
        System.err.println(oe.getMessage());
    }
}

Example 17 with DataProvidersManager

use of org.orekit.data.DataProvidersManager 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());
    }
}

Example 18 with DataProvidersManager

use of org.orekit.data.DataProvidersManager in project Orekit by CS-SI.

the class SlaveMode 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));
        // Initial orbit parameters
        // semi major axis in meters
        double a = 24396159;
        // eccentricity
        double e = 0.72831215;
        // inclination
        double i = FastMath.toRadians(7);
        // perigee argument
        double omega = FastMath.toRadians(180);
        // right ascension of ascending node
        double raan = FastMath.toRadians(261);
        // mean anomaly
        double lM = 0;
        // Inertial frame
        Frame inertialFrame = FramesFactory.getEME2000();
        // Initial date in UTC time scale
        TimeScale utc = TimeScalesFactory.getUTC();
        AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, utc);
        // gravitation coefficient
        double mu = 3.986004415e+14;
        // Orbit construction as Keplerian
        Orbit initialOrbit = new KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN, inertialFrame, initialDate, mu);
        // Simple extrapolation with Keplerian motion
        KeplerianPropagator kepler = new KeplerianPropagator(initialOrbit);
        // Set the propagator to slave mode (could be omitted as it is the default mode)
        kepler.setSlaveMode();
        // Overall duration in seconds for extrapolation
        double duration = 600.;
        // Stop date
        final AbsoluteDate finalDate = initialDate.shiftedBy(duration);
        // Step duration in seconds
        double stepT = 60.;
        // Extrapolation loop
        int cpt = 1;
        for (AbsoluteDate extrapDate = initialDate; extrapDate.compareTo(finalDate) <= 0; extrapDate = extrapDate.shiftedBy(stepT)) {
            SpacecraftState currentState = kepler.propagate(extrapDate);
            System.out.println("step " + cpt++);
            System.out.println(" time : " + currentState.getDate());
            System.out.println(" " + currentState.getOrbit());
        }
    } catch (OrekitException oe) {
        System.err.println(oe.getMessage());
    }
}

Example 19 with DataProvidersManager

use of org.orekit.data.DataProvidersManager in project Orekit by CS-SI.

the class VisibilityCheck 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));
        // Initial state definition : date, orbit
        AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, TimeScalesFactory.getUTC());
        // gravitation coefficient
        double mu = 3.986004415e+14;
        // inertial frame for orbit definition
        Frame inertialFrame = FramesFactory.getEME2000();
        Vector3D position = new Vector3D(-6142438.668, 3492467.560, -25767.25680);
        Vector3D velocity = new Vector3D(505.8479685, 942.7809215, 7435.922231);
        PVCoordinates pvCoordinates = new PVCoordinates(position, velocity);
        Orbit initialOrbit = new KeplerianOrbit(pvCoordinates, inertialFrame, initialDate, mu);
        // Propagator : consider a simple Keplerian motion (could be more elaborate)
        Propagator kepler = new KeplerianPropagator(initialOrbit);
        // Earth and frame
        Frame earthFrame = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
        BodyShape earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, earthFrame);
        // Station
        final double longitude = FastMath.toRadians(45.);
        final double latitude = FastMath.toRadians(25.);
        final double altitude = 0.;
        final GeodeticPoint station1 = new GeodeticPoint(latitude, longitude, altitude);
        final TopocentricFrame sta1Frame = new TopocentricFrame(earth, station1, "station1");
        // Event definition
        final double maxcheck = 60.0;
        final double threshold = 0.001;
        final double elevation = FastMath.toRadians(5.0);
        final EventDetector sta1Visi = new ElevationDetector(maxcheck, threshold, sta1Frame).withConstantElevation(elevation).withHandler(new VisibilityHandler());
        // Add event to be detected
        kepler.addEventDetector(sta1Visi);
        // Propagate from the initial date to the first raising or for the fixed duration
        SpacecraftState finalState = kepler.propagate(initialDate.shiftedBy(1500.));
        System.out.println(" Final state : " + finalState.getDate().durationFrom(initialDate));
    } catch (OrekitException oe) {
        System.err.println(oe.getMessage());
    }
}

Example 20 with DataProvidersManager

use of org.orekit.data.DataProvidersManager in project Orekit by CS-SI.

the class Frames2 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));
        // Considering the following Computing/Measurement date in UTC time scale
        TimeScale utc = TimeScalesFactory.getUTC();
        AbsoluteDate date = new AbsoluteDate(2008, 10, 01, 12, 00, 00.000, utc);
        // The Center of Gravity frame has its origin at the satellite center of gravity (CoG)
        // and its axes parallel to EME2000. It is derived from EME2000 frame at any moment
        // by an unknown transform which depends on the current position and the velocity.
        // Let's initialize this transform by the identity transform.
        UpdatableFrame cogFrame = new UpdatableFrame(FramesFactory.getEME2000(), Transform.IDENTITY, "LOF", false);
        // The satellite frame, with origin also at the CoG, depends on attitude.
        // For the sake of this tutorial, we consider a simple inertial attitude here
        Transform cogToSat = new Transform(date, new Rotation(0.6, 0.48, 0, 0.64, false));
        Frame satFrame = new Frame(cogFrame, cogToSat, "sat", false);
        // Finally, the GPS antenna frame can be defined from the satellite frame by 2 transforms:
        // a translation and a rotation
        Transform translateGPS = new Transform(date, new Vector3D(0, 0, 1));
        Transform rotateGPS = new Transform(date, new Rotation(new Vector3D(0, 1, 3), FastMath.toRadians(10), RotationConvention.VECTOR_OPERATOR));
        Frame gpsFrame = new Frame(satFrame, new Transform(date, translateGPS, rotateGPS), "GPS", false);
        // Let's get the satellite position and velocity in ITRF as measured by GPS antenna at this moment:
        final Vector3D position = new Vector3D(-6142438.668, 3492467.560, -25767.25680);
        final Vector3D velocity = new Vector3D(505.8479685, 942.7809215, 7435.922231);
        System.out.format(Locale.US, "GPS antenna position in ITRF:    %12.3f %12.3f %12.3f%n", position.getX(), position.getY(), position.getZ());
        System.out.format(Locale.US, "GPS antenna velocity in ITRF:    %12.7f %12.7f %12.7f%n", velocity.getX(), velocity.getY(), velocity.getZ());
        // The transform from GPS frame to ITRF frame at this moment is defined by
        // a translation and a rotation. The translation is directly provided by the
        // GPS measurement above. The rotation is extracted from the existing tree, where
        // we know all rotations are already up to date, even if one translation is still
        // unknown. We combine the extracted rotation and the measured translation by
        // applying the rotation first because the position/velocity vector are given in
        // ITRF frame not in GPS antenna frame:
        Transform measuredTranslation = new Transform(date, position, velocity);
        Transform formerTransform = gpsFrame.getTransformTo(FramesFactory.getITRF(IERSConventions.IERS_2010, true), date);
        Transform preservedRotation = new Transform(date, formerTransform.getRotation(), formerTransform.getRotationRate());
        Transform gpsToItrf = new Transform(date, preservedRotation, measuredTranslation);
        // So we can update the transform from EME2000 to CoG frame
        cogFrame.updateTransform(gpsFrame, FramesFactory.getITRF(IERSConventions.IERS_2010, true), gpsToItrf, date);
        // And we can get the position and velocity of satellite CoG in EME2000 frame
        PVCoordinates origin = PVCoordinates.ZERO;
        Transform cogToItrf = cogFrame.getTransformTo(FramesFactory.getITRF(IERSConventions.IERS_2010, true), date);
        PVCoordinates satItrf = cogToItrf.transformPVCoordinates(origin);
        System.out.format(Locale.US, "Satellite   position in ITRF:    %12.3f %12.3f %12.3f%n", satItrf.getPosition().getX(), satItrf.getPosition().getY(), satItrf.getPosition().getZ());
        System.out.format(Locale.US, "Satellite   velocity in ITRF:    %12.7f %12.7f %12.7f%n", satItrf.getVelocity().getX(), satItrf.getVelocity().getY(), satItrf.getVelocity().getZ());
        Transform cogToEme2000 = cogFrame.getTransformTo(FramesFactory.getEME2000(), date);
        PVCoordinates satEME2000 = cogToEme2000.transformPVCoordinates(origin);
        System.out.format(Locale.US, "Satellite   position in EME2000: %12.3f %12.3f %12.3f%n", satEME2000.getPosition().getX(), satEME2000.getPosition().getY(), satEME2000.getPosition().getZ());
        System.out.format(Locale.US, "Satellite   velocity in EME2000: %12.7f %12.7f %12.7f%n", satEME2000.getVelocity().getX(), satEME2000.getVelocity().getY(), satEME2000.getVelocity().getZ());
    } catch (OrekitException oe) {
        System.err.println(oe.getMessage());
    }
}