Examples with DSSTSolarRadiationPressure org.orekit.propagation.semianalytical.dsst.forces.DSSTSolarRadiationPressure used on opensource projects

Example 6 with DSSTSolarRadiationPressure

use of org.orekit.propagation.semianalytical.dsst.forces.DSSTSolarRadiationPressure in project Orekit by CS-SI.

the class DSSTPropagatorTest method testShortPeriodCoefficients.

@Test
public void testShortPeriodCoefficients() throws OrekitException {
    Utils.setDataRoot("regular-data:potential/icgem-format");
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("^eigen-6s-truncated$", false));
    UnnormalizedSphericalHarmonicsProvider nshp = GravityFieldFactory.getUnnormalizedProvider(4, 4);
    Orbit orbit = new KeplerianOrbit(13378000, 0.05, 0, 0, FastMath.PI, 0, PositionAngle.MEAN, FramesFactory.getTOD(false), new AbsoluteDate(2003, 5, 6, TimeScalesFactory.getUTC()), nshp.getMu());
    double period = orbit.getKeplerianPeriod();
    double[][] tolerance = DSSTPropagator.tolerances(1.0, orbit);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(period / 100, period * 100, tolerance[0], tolerance[1]);
    integrator.setInitialStepSize(10 * period);
    DSSTPropagator propagator = new DSSTPropagator(integrator, false);
    OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getGTOD(false));
    CelestialBody sun = CelestialBodyFactory.getSun();
    CelestialBody moon = CelestialBodyFactory.getMoon();
    propagator.addForceModel(new DSSTZonal(nshp, 4, 3, 9));
    propagator.addForceModel(new DSSTTesseral(earth.getBodyFrame(), Constants.WGS84_EARTH_ANGULAR_VELOCITY, nshp, 4, 4, 4, 8, 4, 4, 2));
    propagator.addForceModel(new DSSTThirdBody(sun));
    propagator.addForceModel(new DSSTThirdBody(moon));
    propagator.addForceModel(new DSSTAtmosphericDrag(new HarrisPriester(sun, earth), 2.1, 180));
    propagator.addForceModel(new DSSTSolarRadiationPressure(1.2, 180, sun, earth.getEquatorialRadius()));
    final AbsoluteDate finalDate = orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY);
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateNoConfig = propagator.propagate(finalDate);
    Assert.assertEquals(0, stateNoConfig.getAdditionalStates().size());
    propagator.setSelectedCoefficients(new HashSet<String>());
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateConfigEmpty = propagator.propagate(finalDate);
    Assert.assertEquals(234, stateConfigEmpty.getAdditionalStates().size());
    final Set<String> selected = new HashSet<String>();
    selected.add("DSST-3rd-body-Moon-s[7]");
    selected.add("DSST-central-body-tesseral-c[-2][3]");
    propagator.setSelectedCoefficients(selected);
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateConfigeSelected = propagator.propagate(finalDate);
    Assert.assertEquals(selected.size(), stateConfigeSelected.getAdditionalStates().size());
    propagator.setSelectedCoefficients(null);
    propagator.resetInitialState(new SpacecraftState(orbit, 45.0));
    final SpacecraftState stateConfigNull = propagator.propagate(finalDate);
    Assert.assertEquals(0, stateConfigNull.getAdditionalStates().size());
}

Example 7 with DSSTSolarRadiationPressure

use of org.orekit.propagation.semianalytical.dsst.forces.DSSTSolarRadiationPressure in project Orekit by CS-SI.

the class IntegratedEphemerisTest method doTestSerializationDSST.

private void doTestSerializationDSST(boolean meanOnly, int expectedSize) throws OrekitException, IOException, ClassNotFoundException {
    AbsoluteDate finalDate = initialOrbit.getDate().shiftedBy(Constants.JULIAN_DAY);
    final double[][] tol = DSSTPropagator.tolerances(1.0, initialOrbit);
    AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(10, Constants.JULIAN_DAY, tol[0], tol[1]);
    DSSTPropagator dsstProp = new DSSTPropagator(integrator, meanOnly);
    dsstProp.setInitialState(new SpacecraftState(initialOrbit), false);
    dsstProp.setEphemerisMode();
    final Frame itrf = FramesFactory.getITRF(IERSConventions.IERS_2010, true);
    final UnnormalizedSphericalHarmonicsProvider gravity = GravityFieldFactory.getUnnormalizedProvider(8, 8);
    final CelestialBody sun = CelestialBodyFactory.getSun();
    final CelestialBody moon = CelestialBodyFactory.getMoon();
    final RadiationSensitive spacecraft = new IsotropicRadiationSingleCoefficient(20.0, 2.0);
    dsstProp.addForceModel(new DSSTZonal(gravity, 8, 7, 17));
    dsstProp.addForceModel(new DSSTTesseral(itrf, Constants.WGS84_EARTH_ANGULAR_VELOCITY, gravity, 8, 8, 4, 12, 8, 8, 4));
    dsstProp.addForceModel(new DSSTThirdBody(sun));
    dsstProp.addForceModel(new DSSTThirdBody(moon));
    dsstProp.addForceModel(new DSSTSolarRadiationPressure(sun, Constants.WGS84_EARTH_EQUATORIAL_RADIUS, spacecraft));
    dsstProp.propagate(finalDate);
    IntegratedEphemeris ephemeris = (IntegratedEphemeris) dsstProp.getGeneratedEphemeris();
    ByteArrayOutputStream bos = new ByteArrayOutputStream();
    ObjectOutputStream oos = new ObjectOutputStream(bos);
    oos.writeObject(ephemeris);
    Assert.assertTrue("size = " + bos.size(), bos.size() > 9 * expectedSize / 10);
    Assert.assertTrue("size = " + bos.size(), bos.size() < 11 * expectedSize / 10);
    Assert.assertNotNull(ephemeris.getFrame());
    Assert.assertSame(ephemeris.getFrame(), dsstProp.getFrame());
    ByteArrayInputStream bis = new ByteArrayInputStream(bos.toByteArray());
    ObjectInputStream ois = new ObjectInputStream(bis);
    IntegratedEphemeris deserialized = (IntegratedEphemeris) ois.readObject();
    Assert.assertEquals(deserialized.getMinDate(), deserialized.getMinDate());
    Assert.assertEquals(deserialized.getMaxDate(), deserialized.getMaxDate());
}