Examples with NumericalPropagator org.orekit.propagation.numerical.NumericalPropagator used on opensource projects

Example 91 with NumericalPropagator

use of org.orekit.propagation.numerical.NumericalPropagator in project Orekit by CS-SI.

the class DSSTPropagation method createNumProp.

/**
 * Create the numerical propagator
 *
 *  @param orbit initial orbit
 *  @param mass S/C mass (kg)
 *  @throws OrekitException
 */
private NumericalPropagator createNumProp(final Orbit orbit, final double mass) throws OrekitException {
    final double[][] tol = NumericalPropagator.tolerances(1.0, orbit, orbit.getType());
    final double minStep = 1.e-3;
    final double maxStep = 1.e+3;
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tol[0], tol[1]);
    integrator.setInitialStepSize(100.);
    NumericalPropagator numProp = new NumericalPropagator(integrator);
    numProp.setInitialState(new SpacecraftState(orbit, mass));
    return numProp;
}

Example 92 with NumericalPropagator

use of org.orekit.propagation.numerical.NumericalPropagator in project Orekit by CS-SI.

the class SecularAndHarmonicTest method fitGMST.

private double[] fitGMST(CircularOrbit orbit, int nbOrbits, double mst) throws OrekitException {
    double period = orbit.getKeplerianPeriod();
    double duration = nbOrbits * period;
    NumericalPropagator propagator = createPropagator(orbit);
    SecularAndHarmonic mstModel = new SecularAndHarmonic(2, 2.0 * FastMath.PI / Constants.JULIAN_YEAR, 4.0 * FastMath.PI / Constants.JULIAN_YEAR, 2.0 * FastMath.PI / Constants.JULIAN_DAY, 4.0 * FastMath.PI / Constants.JULIAN_DAY);
    mstModel.resetFitting(orbit.getDate(), new double[] { mst, -1.0e-10, -1.0e-17, 1.0e-3, 1.0e-3, 1.0e-5, 1.0e-5, 1.0e-5, 1.0e-5, 1.0e-5, 1.0e-5 });
    // first descending node
    SpacecraftState crossing = findFirstCrossing(0.0, false, orbit.getDate(), orbit.getDate().shiftedBy(2 * period), 0.01 * period, propagator);
    while (crossing != null && crossing.getDate().durationFrom(orbit.getDate()) < (nbOrbits * period)) {
        final AbsoluteDate previousDate = crossing.getDate();
        crossing = findLatitudeCrossing(0.0, previousDate.shiftedBy(period), previousDate.shiftedBy(2 * period), 0.01 * period, period / 8, propagator);
        if (crossing != null) {
            // store current point
            mstModel.addPoint(crossing.getDate(), meanSolarTime(crossing.getOrbit()));
            // use the same time separation to pinpoint next crossing
            period = crossing.getDate().durationFrom(previousDate);
        }
    }
    // fit the mean solar time to a parabolic model, taking care the main
    // periods are properly removed
    mstModel.fit();
    return mstModel.approximateAsPolynomialOnly(1, orbit.getDate(), 2, 2, orbit.getDate(), orbit.getDate().shiftedBy(duration), 0.01 * period);
}

Example 93 with NumericalPropagator

use of org.orekit.propagation.numerical.NumericalPropagator in project Orekit by CS-SI.

the class EphemerisMode 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);
        // Initialize state
        SpacecraftState initialState = new SpacecraftState(initialOrbit);
        // Numerical propagation with no perturbation (only Keplerian movement)
        // Using a very simple integrator with a fixed step: classical Runge-Kutta
        // the step is ten seconds
        double stepSize = 10;
        AbstractIntegrator integrator = new ClassicalRungeKuttaIntegrator(stepSize);
        NumericalPropagator propagator = new NumericalPropagator(integrator);
        // Set the propagator to ephemeris mode
        propagator.setEphemerisMode();
        // Initialize propagation
        propagator.setInitialState(initialState);
        // Propagation with storage of the results in an integrated ephemeris
        SpacecraftState finalState = propagator.propagate(initialDate.shiftedBy(6000));
        System.out.println(" Numerical propagation :");
        System.out.println("  Final date : " + finalState.getDate());
        System.out.println("  " + finalState.getOrbit());
        // Getting the integrated ephemeris
        BoundedPropagator ephemeris = propagator.getGeneratedEphemeris();
        System.out.println(" Ephemeris defined from " + ephemeris.getMinDate() + " to " + ephemeris.getMaxDate());
        System.out.println(" Ephemeris propagation :");
        AbsoluteDate intermediateDate = initialDate.shiftedBy(3000);
        SpacecraftState intermediateState = ephemeris.propagate(intermediateDate);
        System.out.println("  date :  " + intermediateState.getDate());
        System.out.println("  " + intermediateState.getOrbit());
        intermediateDate = finalState.getDate();
        intermediateState = ephemeris.propagate(intermediateDate);
        System.out.println("  date :  " + intermediateState.getDate());
        System.out.println("  " + intermediateState.getOrbit());
        intermediateDate = initialDate.shiftedBy(-1000);
        System.out.println();
        System.out.println("Attempting to propagate to date " + intermediateDate + " which is OUT OF RANGE");
        System.out.println("This propagation attempt should fail, " + "so an error message shoud appear below, " + "this is expected and shows that errors are handled correctly");
        intermediateState = ephemeris.propagate(intermediateDate);
        // these two print should never happen as en exception should have been triggered
        System.out.println("  date :  " + intermediateState.getDate());
        System.out.println("  " + intermediateState.getOrbit());
    } catch (OrekitException oe) {
        System.out.println(oe.getMessage());
    }
}

Example 94 with NumericalPropagator

use of org.orekit.propagation.numerical.NumericalPropagator in project Orekit by CS-SI.

the class MasterMode 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));
        // gravitation coefficient
        double mu = 3.986004415e+14;
        // inertial frame
        Frame inertialFrame = FramesFactory.getEME2000();
        // Initial date
        AbsoluteDate initialDate = new AbsoluteDate(2004, 01, 01, 23, 30, 00.000, TimeScalesFactory.getUTC());
        // Initial orbit
        // 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 ascention of ascending node
        double raan = FastMath.toRadians(261);
        // mean anomaly
        double lM = 0;
        Orbit initialOrbit = new KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN, inertialFrame, initialDate, mu);
        // Initial state definition
        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 = 1000.0;
        final double positionTolerance = 10.0;
        final OrbitType propagationType = OrbitType.KEPLERIAN;
        final double[][] tolerances = NumericalPropagator.tolerances(positionTolerance, initialOrbit, propagationType);
        AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxstep, tolerances[0], tolerances[1]);
        // Propagator
        NumericalPropagator propagator = new NumericalPropagator(integrator);
        propagator.setOrbitType(propagationType);
        // Force Model (reduced to perturbing gravity field)
        final NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(10, 10);
        ForceModel holmesFeatherstone = new HolmesFeatherstoneAttractionModel(FramesFactory.getITRF(IERSConventions.IERS_2010, true), provider);
        // Add force model to the propagator
        propagator.addForceModel(holmesFeatherstone);
        // Set up initial state in the propagator
        propagator.setInitialState(initialState);
        // Set up operating mode for the propagator as master mode
        // with fixed step and specialized step handler
        propagator.setMasterMode(60., new TutorialStepHandler());
        // Extrapolate from the initial to the final date
        SpacecraftState finalState = propagator.propagate(initialDate.shiftedBy(630.));
        KeplerianOrbit o = (KeplerianOrbit) OrbitType.KEPLERIAN.convertType(finalState.getOrbit());
        System.out.format(Locale.US, "Final state:%n%s %12.3f %10.8f %10.6f %10.6f %10.6f %10.6f%n", finalState.getDate(), o.getA(), o.getE(), FastMath.toDegrees(o.getI()), FastMath.toDegrees(o.getPerigeeArgument()), FastMath.toDegrees(o.getRightAscensionOfAscendingNode()), FastMath.toDegrees(o.getTrueAnomaly()));
    } catch (OrekitException oe) {
        System.err.println(oe.getMessage());
    }
}

Example 95 with NumericalPropagator

use of org.orekit.propagation.numerical.NumericalPropagator in project SpriteOrbits by ProjectPersephone.

the class SpritePropOrig method createPropagator.

/**
 * Create a numerical propagator for a state.
 * @param state state to propagate
 * @param attitudeProvider provider for the attitude
 * @param crossSection cross section of the object
 * @param dragCoeff drag coefficient
 */
private Propagator createPropagator(final SpacecraftState state, final AttitudeProvider attitudeProvider, final double crossSection, final double dragCoeff) throws OrekitException {
    // see https://www.orekit.org/static/architecture/propagation.html
    // steps limits
    final double minStep = 0.001;
    final double maxStep = 1000;
    final double initStep = 60;
    // error control parameters (absolute and relative)
    final double positionError = 10.0;
    // we will propagate in Cartesian coordinates
    final OrbitType orbitType = OrbitType.CARTESIAN;
    final double[][] tolerances = NumericalPropagator.tolerances(positionError, state.getOrbit(), orbitType);
    // set up mathematical integrator
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tolerances[0], tolerances[1]);
    integrator.setInitialStepSize(initStep);
    // set up space dynamics propagator
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(orbitType);
    // add gravity field force model
    final NormalizedSphericalHarmonicsProvider gravityProvider = GravityFieldFactory.getNormalizedProvider(8, 8);
    propagator.addForceModel(new HolmesFeatherstoneAttractionModel(earth.getBodyFrame(), gravityProvider));
    // add atmospheric drag force model
    propagator.addForceModel(new DragForce(new HarrisPriester(sun, earth), new SphericalSpacecraft(crossSection, dragCoeff, 0.0, 0.0)));
    // set attitude mode
    propagator.setAttitudeProvider(attitudeProvider);
    propagator.setInitialState(state);
    return propagator;
}