Examples with CartesianOrbit org.orekit.orbits.CartesianOrbit used on opensource projects

Example 21 with CartesianOrbit

use of org.orekit.orbits.CartesianOrbit in project SpriteOrbits by ProjectPersephone.

the class SpritePropOrig method main.

/**
 * @param args
 */
public static void main(String[] args) {
    PrintStream out = null;
    try {
        // set up orekit data
        // for this to work, the orekit-data.zip file avaiable at the following URL
        // must be put in the current working directory
        // URL:     https://www.orekit.org/forge/projects/orekit/files
        File userDir = new File(System.getProperty("user.dir"));
        File orekitZip = new File(userDir, "orekit-data.zip");
        DataProvidersManager.getInstance().addProvider(new ZipJarCrawler(orekitZip));
        // reference models
        // this is the frame labeled J2K in the NASA page
        final Frame eme2000 = FramesFactory.getEME2000();
        final Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
        final TimeScale utc = TimeScalesFactory.getUTC();
        // central attraction coefficient MU
        final double mu = Constants.EIGEN5C_EARTH_MU;
        // set up some data, ideally, this should be provided as input to the program
        // ISS orbit is from http://spaceflight.nasa.gov/realdata/sightings/SSapplications/Post/JavaSSOP/orbit/ISS/SVPOST.html
        // 
        // Coasting Arc #15 (Orbit 3386)
        // ---------------------------------------
        // 
        // Vector Time (GMT): 2014/055/16:27:06.921
        // Vector Time (MET): N/A
        // Weight (LBS)     : 911651.1
        // 
        // ...
        // 
        // J2K Cartesian
        // --------------------------------
        // X    =         2998767.75
        // Y    =        -6097451.56  meter
        // ...              Z    =         -141448.92
        // XDOT =        4323.077242
        // YDOT =        1994.291706  meter/sec
        // ZDOT =        6000.774574
        // we start small, we can increase this later
        final int numberOfSprites = 12;
        // (m/s)
        final double relativeReleaseVelocity = 1.0;
        final AbsoluteDate releaseDate = new // year, month, day
        AbsoluteDate(// year, month, day
        2014, // year, month, day
        3, // year, month, day
        1, // hours, minutes, seconds
        12, // hours, minutes, seconds
        0, // hours, minutes, seconds
        0.0, utc);
        final CartesianOrbit kickSatOrbit = new CartesianOrbit(new // position (m)
        PVCoordinates(// position (m)
        new Vector3D(2998767.75, -6097451.56, -141448.92), // velocity (m/s)
        new Vector3D(4323.077242, 1994.291706, 6000.774574)), eme2000, new // year, day in year as NASA page above
        AbsoluteDate(// year, day in year as NASA page above
        new DateComponents(2014, 55), // hour in day
        new TimeComponents(16, 27, 6.921), utc), mu);
        // kg
        final double kickSatMass = 10.0;
        // m^2
        final double kickSatCrossSection = 0.03;
        // no units
        final double kickSatDragCoeff = 2.2;
        // kg
        final double spriteMass = 0.01;
        // m^2
        final double spriteCrossSection = 2.5e-3;
        // no units
        final double spriteDragCoeff = 2.2;
        SpritePropOrig spriteProp = new SpritePropOrig(numberOfSprites, kickSatOrbit, kickSatMass, kickSatCrossSection, kickSatDragCoeff, spriteMass, spriteCrossSection, spriteDragCoeff, relativeReleaseVelocity, releaseDate, itrf);
        // days after release
        final double propagationDuration = 0.2;
        // seconds
        final double step = 60.0;
        out = new PrintStream(new File(userDir, "sprites-prop.txt"));
        spriteProp.run(out, propagationDuration, step, utc);
    } catch (IOException e) {
        e.printStackTrace();
    } catch (IllegalArgumentException e) {
        e.printStackTrace();
    } catch (OrekitException e) {
        e.printStackTrace();
    } finally {
        if (out != null) {
            out.close();
        }
    }
}

Example 22 with CartesianOrbit

use of org.orekit.orbits.CartesianOrbit in project Orekit by CS-SI.

the class KalmanOrbitDeterminationTest method runReference.

/**
 * Use the physical models in the input file
 * Incorporate the initial reference values
 * And run the propagation until the last measurement to get the reference orbit at the same date
 * as the Kalman filter
 * @param input Input configuration file
 * @param orbitType Orbit type to use (calculation and display)
 * @param refPosition Initial reference position
 * @param refVelocity Initial reference velocity
 * @param refPropagationParameters Reference propagation parameters
 * @param kalmanFinalDate The final date of the Kalman filter
 * @return The reference orbit at the same date as the Kalman filter
 * @throws IOException Input file cannot be opened
 * @throws IllegalArgumentException Issue in key/value reading of input file
 * @throws OrekitException An Orekit exception... should be explicit
 * @throws ParseException Parsing of the input file or measurement file failed
 */
private Orbit runReference(final File input, final OrbitType orbitType, final Vector3D refPosition, final Vector3D refVelocity, final ParameterDriversList refPropagationParameters, final AbsoluteDate kalmanFinalDate) throws IOException, IllegalArgumentException, OrekitException, ParseException {
    // Read input parameters
    KeyValueFileParser<ParameterKey> parser = new KeyValueFileParser<ParameterKey>(ParameterKey.class);
    parser.parseInput(input.getAbsolutePath(), new FileInputStream(input));
    // Gravity field
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("eigen-5c.gfc", true));
    final NormalizedSphericalHarmonicsProvider gravityField = createGravityField(parser);
    // Orbit initial guess
    Orbit initialRefOrbit = new CartesianOrbit(new PVCoordinates(refPosition, refVelocity), parser.getInertialFrame(ParameterKey.INERTIAL_FRAME), parser.getDate(ParameterKey.ORBIT_DATE, TimeScalesFactory.getUTC()), gravityField.getMu());
    // Convert to desired orbit type
    initialRefOrbit = orbitType.convertType(initialRefOrbit);
    // IERS conventions
    final IERSConventions conventions;
    if (!parser.containsKey(ParameterKey.IERS_CONVENTIONS)) {
        conventions = IERSConventions.IERS_2010;
    } else {
        conventions = IERSConventions.valueOf("IERS_" + parser.getInt(ParameterKey.IERS_CONVENTIONS));
    }
    // Central body
    final OneAxisEllipsoid body = createBody(parser);
    // Propagator builder
    final NumericalPropagatorBuilder propagatorBuilder = createPropagatorBuilder(parser, conventions, gravityField, body, initialRefOrbit);
    // Force the selected propagation parameters to their reference values
    if (refPropagationParameters != null) {
        for (DelegatingDriver refDriver : refPropagationParameters.getDrivers()) {
            for (DelegatingDriver driver : propagatorBuilder.getPropagationParametersDrivers().getDrivers()) {
                if (driver.getName().equals(refDriver.getName())) {
                    driver.setValue(refDriver.getValue());
                }
            }
        }
    }
    // Build the reference propagator
    final NumericalPropagator propagator = propagatorBuilder.buildPropagator(propagatorBuilder.getSelectedNormalizedParameters());
    // Propagate until last date and return the orbit
    return propagator.propagate(kalmanFinalDate).getOrbit();
}

Example 23 with CartesianOrbit

use of org.orekit.orbits.CartesianOrbit in project Orekit by CS-SI.

the class OEMParserTest method testParseOEM1.

@Test
public void testParseOEM1() throws OrekitException, IOException {
    // 
    final String ex = "/ccsds/OEMExample.txt";
    final InputStream inEntry = getClass().getResourceAsStream(ex);
    final OEMParser parser = new OEMParser().withMu(CelestialBodyFactory.getEarth().getGM());
    final OEMFile file = parser.parse(inEntry, "OEMExample.txt");
    Assert.assertEquals(CcsdsTimeScale.UTC, file.getEphemeridesBlocks().get(0).getMetaData().getTimeSystem());
    Assert.assertEquals("MARS GLOBAL SURVEYOR", file.getEphemeridesBlocks().get(0).getMetaData().getObjectName());
    Assert.assertEquals("1996-062A", file.getEphemeridesBlocks().get(0).getMetaData().getObjectID());
    Assert.assertEquals("MARS BARYCENTER", file.getEphemeridesBlocks().get(0).getMetaData().getCenterName());
    Assert.assertEquals(1996, file.getEphemeridesBlocks().get(0).getMetaData().getLaunchYear());
    Assert.assertEquals(62, file.getEphemeridesBlocks().get(0).getMetaData().getLaunchNumber());
    Assert.assertEquals("A", file.getEphemeridesBlocks().get(0).getMetaData().getLaunchPiece());
    Assert.assertFalse(file.getEphemeridesBlocks().get(0).getMetaData().getHasCreatableBody());
    Assert.assertNull(file.getEphemeridesBlocks().get(0).getMetaData().getCenterBody());
    Assert.assertEquals(new AbsoluteDate(1996, 12, 18, 12, 00, 0.331, TimeScalesFactory.getUTC()), file.getEphemeridesBlocks().get(0).getStartTime());
    Assert.assertEquals(new AbsoluteDate(1996, 12, 28, 21, 28, 0.331, TimeScalesFactory.getUTC()), file.getEphemeridesBlocks().get(0).getStopTime());
    Assert.assertEquals(new AbsoluteDate(1996, 12, 18, 12, 10, 0.331, TimeScalesFactory.getUTC()), file.getEphemeridesBlocks().get(0).getUseableStartTime());
    Assert.assertEquals(new AbsoluteDate(1996, 12, 28, 21, 23, 0.331, TimeScalesFactory.getUTC()), file.getEphemeridesBlocks().get(0).getUseableStopTime());
    Assert.assertEquals("HERMITE", file.getEphemeridesBlocks().get(0).getInterpolationMethod());
    Assert.assertEquals(7, file.getEphemeridesBlocks().get(0).getInterpolationDegree());
    ArrayList<String> ephemeridesDataLinesComment = new ArrayList<String>();
    ephemeridesDataLinesComment.add("This file was produced by M.R. Somebody, MSOO NAV/JPL, 1996NOV 04. It is");
    ephemeridesDataLinesComment.add("to be used for DSN scheduling purposes only.");
    Assert.assertEquals(ephemeridesDataLinesComment, file.getEphemeridesBlocks().get(0).getEphemeridesDataLinesComment());
    CartesianOrbit orbit = new CartesianOrbit(new PVCoordinates(new Vector3D(2789.619 * 1000, -280.045 * 1000, -1746.755 * 1000), new Vector3D(4.73372 * 1000, -2.49586 * 1000, -1.04195 * 1000)), FramesFactory.getEME2000(), new AbsoluteDate("1996-12-18T12:00:00.331", TimeScalesFactory.getUTC()), CelestialBodyFactory.getEarth().getGM());
    Assert.assertArrayEquals(orbit.getPVCoordinates().getPosition().toArray(), file.getEphemeridesBlocks().get(0).getEphemeridesDataLines().get(0).getPosition().toArray(), 1e-10);
    Assert.assertArrayEquals(orbit.getPVCoordinates().getVelocity().toArray(), file.getEphemeridesBlocks().get(0).getEphemeridesDataLines().get(0).getVelocity().toArray(), 1e-10);
    Assert.assertArrayEquals((new Vector3D(1, 1, 1)).toArray(), file.getEphemeridesBlocks().get(1).getEphemeridesDataLines().get(0).getAcceleration().toArray(), 1e-10);
    Assert.assertEquals(Vector3D.ZERO, file.getEphemeridesBlocks().get(1).getEphemeridesDataLines().get(1).getAcceleration());
    final Array2DRowRealMatrix covMatrix = new Array2DRowRealMatrix(6, 6);
    final double[] column1 = { 3.331349476038534e-04, 4.618927349220216e-04, -3.070007847730449e-04, -3.349365033922630e-07, -2.211832501084875e-07, -3.041346050686871e-07 };
    final double[] column2 = { 4.618927349220216e-04, 6.782421679971363e-04, -4.221234189514228e-04, -4.686084221046758e-07, -2.864186892102733e-07, -4.989496988610662e-07 };
    final double[] column3 = { -3.070007847730449e-04, -4.221234189514228e-04, 3.231931992380369e-04, 2.484949578400095e-07, 1.798098699846038e-07, 3.540310904497689e-07 };
    final double[] column4 = { -3.349365033922630e-07, -4.686084221046758e-07, 2.484949578400095e-07, 4.296022805587290e-10, 2.608899201686016e-10, 1.869263192954590e-10 };
    final double[] column5 = { -2.211832501084875e-07, -2.864186892102733e-07, 1.798098699846038e-07, 2.608899201686016e-10, 1.767514756338532e-10, 1.008862586240695e-10 };
    final double[] column6 = { -3.041346050686871e-07, -4.989496988610662e-07, 3.540310904497689e-07, 1.869263192954590e-10, 1.008862586240695e-10, 6.224444338635500e-10 };
    covMatrix.setColumn(0, column1);
    covMatrix.setColumn(1, column2);
    covMatrix.setColumn(2, column3);
    covMatrix.setColumn(3, column4);
    covMatrix.setColumn(4, column5);
    covMatrix.setColumn(5, column6);
    for (int i = 0; i < 6; i++) {
        for (int j = 0; j < 6; j++) {
            Assert.assertEquals(covMatrix.getEntry(i, j), file.getEphemeridesBlocks().get(2).getCovarianceMatrices().get(0).getMatrix().getEntry(i, j), 1e-10);
        }
    }
    Assert.assertEquals(new AbsoluteDate("1996-12-28T21:29:07.267", TimeScalesFactory.getUTC()), file.getEphemeridesBlocks().get(2).getCovarianceMatrices().get(0).getEpoch());
    Assert.assertEquals(LOFType.QSW, file.getEphemeridesBlocks().get(2).getCovarianceMatrices().get(0).getLofType());
    Assert.assertNull(file.getEphemeridesBlocks().get(2).getCovarianceMatrices().get(0).getFrame());
    Assert.assertEquals(FramesFactory.getEME2000(), file.getEphemeridesBlocks().get(2).getCovarianceMatrices().get(1).getFrame());
    Assert.assertNull(file.getEphemeridesBlocks().get(2).getCovarianceMatrices().get(1).getLofType());
}

Example 24 with CartesianOrbit

use of org.orekit.orbits.CartesianOrbit in project Orekit by CS-SI.

the class ConstantThrustManeuverTest method testParameterDerivative.

@Test
public void testParameterDerivative() throws OrekitException {
    // pos-vel (from a ZOOM ephemeris reference)
    final Vector3D pos = new Vector3D(6.46885878304673824e+06, -1.88050918456274318e+06, -1.32931592294715829e+04);
    final Vector3D vel = new Vector3D(2.14718074509906819e+03, 7.38239351251748485e+03, -1.14097953925384523e+01);
    final SpacecraftState state = new SpacecraftState(new CartesianOrbit(new PVCoordinates(pos, vel), FramesFactory.getGCRF(), new AbsoluteDate(2005, 3, 5, 0, 24, 0.0, TimeScalesFactory.getTAI()), Constants.EIGEN5C_EARTH_MU));
    final ConstantThrustManeuver maneuver = new ConstantThrustManeuver(state.getDate().shiftedBy(-10), 100.0, 20.0, 350.0, Vector3D.PLUS_I, "along-X-");
    maneuver.init(state, state.getDate().shiftedBy(3600.0));
    checkParameterDerivative(state, maneuver, "along-X-thrust", 1.0e-3, 3.0e-14);
    checkParameterDerivative(state, maneuver, "along-X-flow rate", 1.0e-3, 1.0e-15);
}

Example 25 with CartesianOrbit

use of org.orekit.orbits.CartesianOrbit in project Orekit by CS-SI.

the class ImpulseManeuverTest method testAdditionalStateNumerical.

@Test
public void testAdditionalStateNumerical() throws OrekitException {
    final double mu = CelestialBodyFactory.getEarth().getGM();
    final double initialX = 7100e3;
    final double initialY = 0.0;
    final double initialZ = 1300e3;
    final double initialVx = 0;
    final double initialVy = 8000;
    final double initialVz = 1000;
    final Vector3D position = new Vector3D(initialX, initialY, initialZ);
    final Vector3D velocity = new Vector3D(initialVx, initialVy, initialVz);
    final AbsoluteDate epoch = new AbsoluteDate(2010, 1, 1, 0, 0, 0, TimeScalesFactory.getUTC());
    final TimeStampedPVCoordinates pv = new TimeStampedPVCoordinates(epoch, position, velocity, Vector3D.ZERO);
    final Orbit initialOrbit = new CartesianOrbit(pv, FramesFactory.getEME2000(), mu);
    final double totalPropagationTime = 10.0;
    final double deltaX = 0.01;
    final double deltaY = 0.02;
    final double deltaZ = 0.03;
    final double isp = 300;
    final Vector3D deltaV = new Vector3D(deltaX, deltaY, deltaZ);
    final AttitudeProvider attitudeProvider = new LofOffset(initialOrbit.getFrame(), LOFType.VNC);
    final Attitude initialAttitude = attitudeProvider.getAttitude(initialOrbit, initialOrbit.getDate(), initialOrbit.getFrame());
    double[][] tolerances = NumericalPropagator.tolerances(10.0, initialOrbit, initialOrbit.getType());
    DormandPrince853Integrator integrator = new DormandPrince853Integrator(1.0e-3, 60, tolerances[0], tolerances[1]);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(initialOrbit.getType());
    PartialDerivativesEquations pde = new PartialDerivativesEquations("derivatives", propagator);
    final SpacecraftState initialState = pde.setInitialJacobians(new SpacecraftState(initialOrbit, initialAttitude));
    propagator.resetInitialState(initialState);
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(0.5 * totalPropagationTime));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, RotationConvention.VECTOR_OPERATOR, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector, attitudeOverride, deltaV, isp).withThreshold(1.0e-3);
    propagator.addEventDetector(burnAtEpoch);
    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));
    Assert.assertEquals(1, finalState.getAdditionalStates().size());
    Assert.assertEquals(36, finalState.getAdditionalState("derivatives").length);
    double[][] stateTransitionMatrix = new double[6][6];
    pde.getMapper().getStateJacobian(finalState, stateTransitionMatrix);
    for (int i = 0; i < 6; ++i) {
        for (int j = 0; j < 6; ++j) {
            double sIJ = stateTransitionMatrix[i][j];
            if (j == i) {
                // dPi/dPj and dVi/dVj are roughly 1 for small propagation times
                Assert.assertEquals(1.0, sIJ, 2.0e-4);
            } else if (j == i + 3) {
                // dVi/dPi is roughly the propagation time for small propagation times
                Assert.assertEquals(totalPropagationTime, sIJ, 4.0e-5 * totalPropagationTime);
            } else {
                // other derivatives are almost zero for small propagation times
                Assert.assertEquals(0, sIJ, 1.0e-4);
            }
        }
    }
}