Examples with KeplerianPropagator org.orekit.propagation.analytical.KeplerianPropagator used on opensource projects

Example 56 with KeplerianPropagator

use of org.orekit.propagation.analytical.KeplerianPropagator 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);
    }
}

Example 57 with KeplerianPropagator

use of org.orekit.propagation.analytical.KeplerianPropagator 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 58 with KeplerianPropagator

use of org.orekit.propagation.analytical.KeplerianPropagator 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 59 with KeplerianPropagator

use of org.orekit.propagation.analytical.KeplerianPropagator 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());
    }
}