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

Example 1 with DSSTAtmosphericDrag

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

the class DSSTPropagatorTest method testIssue339WithAccelerations.

@Test
public void testIssue339WithAccelerations() throws OrekitException {
    final SpacecraftState osculatingState = getLEOStatePropagatedBy30Minutes();
    final CelestialBody sun = CelestialBodyFactory.getSun();
    final OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
    final BoxAndSolarArraySpacecraft boxAndWing = new BoxAndSolarArraySpacecraft(5.0, 2.0, 2.0, sun, 50.0, Vector3D.PLUS_J, 2.0, 0.1, 0.2, 0.6);
    final Atmosphere atmosphere = new HarrisPriester(CelestialBodyFactory.getSun(), earth, 6);
    final AttitudeProvider attitudeProvider = new LofOffset(osculatingState.getFrame(), LOFType.VVLH, RotationOrder.XYZ, 0.0, 0.0, 0.0);
    // Surface force models that require an attitude provider
    final Collection<DSSTForceModel> forces = new ArrayList<DSSTForceModel>();
    forces.add(new DSSTAtmosphericDrag(atmosphere, boxAndWing));
    final SpacecraftState meanState = DSSTPropagator.computeMeanState(osculatingState, attitudeProvider, forces);
    final SpacecraftState computedOsculatingState = DSSTPropagator.computeOsculatingState(meanState, attitudeProvider, forces);
    Assert.assertEquals(0.0, Vector3D.distance(osculatingState.getPVCoordinates().getPosition(), computedOsculatingState.getPVCoordinates().getPosition()), 5.0e-6);
}

Example 2 with DSSTAtmosphericDrag

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

the class DSSTPropagatorTest method testIssue339.

@Test
public void testIssue339() throws OrekitException {
    final SpacecraftState osculatingState = getLEOState();
    final CelestialBody sun = CelestialBodyFactory.getSun();
    final OneAxisEllipsoid earth = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true));
    final BoxAndSolarArraySpacecraft boxAndWing = new BoxAndSolarArraySpacecraft(5.0, 2.0, 2.0, sun, 50.0, Vector3D.PLUS_J, 2.0, 0.1, 0.2, 0.6);
    final Atmosphere atmosphere = new HarrisPriester(CelestialBodyFactory.getSun(), earth, 6);
    final AttitudeProvider attitudeProvider = new LofOffset(osculatingState.getFrame(), LOFType.VVLH, RotationOrder.XYZ, 0.0, 0.0, 0.0);
    // Surface force models that require an attitude provider
    final Collection<DSSTForceModel> forces = new ArrayList<DSSTForceModel>();
    forces.add(new DSSTSolarRadiationPressure(sun, Constants.WGS84_EARTH_EQUATORIAL_RADIUS, boxAndWing));
    forces.add(new DSSTAtmosphericDrag(atmosphere, boxAndWing));
    final SpacecraftState meanState = DSSTPropagator.computeMeanState(osculatingState, attitudeProvider, forces);
    Assert.assertEquals(0.522, Vector3D.distance(osculatingState.getPVCoordinates().getPosition(), meanState.getPVCoordinates().getPosition()), 0.001);
    final SpacecraftState computedOsculatingState = DSSTPropagator.computeOsculatingState(meanState, attitudeProvider, forces);
    Assert.assertEquals(0.0, Vector3D.distance(osculatingState.getPVCoordinates().getPosition(), computedOsculatingState.getPVCoordinates().getPosition()), 5.0e-6);
}

Example 3 with DSSTAtmosphericDrag

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

the class DSSTPropagatorTest method testEphemerisGeneration.

@Test
public void testEphemerisGeneration() throws OrekitException {
    Utils.setDataRoot("regular-data:potential/icgem-format");
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("^eigen-6s-truncated$", false));
    UnnormalizedSphericalHarmonicsProvider nshp = GravityFieldFactory.getUnnormalizedProvider(8, 8);
    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, 8, 7, 17));
    propagator.addForceModel(new DSSTTesseral(earth.getBodyFrame(), Constants.WGS84_EARTH_ANGULAR_VELOCITY, nshp, 8, 8, 4, 12, 8, 8, 4));
    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()));
    propagator.setInterpolationGridToMaxTimeGap(0.5 * Constants.JULIAN_DAY);
    // direct generation of states
    propagator.setInitialState(new SpacecraftState(orbit, 45.0), false);
    final List<SpacecraftState> states = new ArrayList<SpacecraftState>();
    propagator.setMasterMode(600, (currentState, isLast) -> states.add(currentState));
    propagator.propagate(orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY));
    // ephemeris generation
    propagator.setInitialState(new SpacecraftState(orbit, 45.0), false);
    propagator.setEphemerisMode();
    propagator.propagate(orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY));
    BoundedPropagator ephemeris = propagator.getGeneratedEphemeris();
    double maxError = 0;
    for (final SpacecraftState state : states) {
        final SpacecraftState fromEphemeris = ephemeris.propagate(state.getDate());
        final double error = Vector3D.distance(state.getPVCoordinates().getPosition(), fromEphemeris.getPVCoordinates().getPosition());
        maxError = FastMath.max(maxError, error);
    }
    Assert.assertEquals(0.0, maxError, 1.0e-10);
}

Example 4 with DSSTAtmosphericDrag

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

the class DSSTPropagation method setForceModel.

/**
 * Set DSST propagator force models
 *
 *  @param parser input file parser
 *  @param unnormalized spherical harmonics provider
 *  @param earthFrame Earth rotating frame
 *  @param rotationRate central body rotation rate (rad/s)
 *  @param dsstProp DSST propagator
 *  @throws IOException
 *  @throws OrekitException
 */
private void setForceModel(final KeyValueFileParser<ParameterKey> parser, final UnnormalizedSphericalHarmonicsProvider unnormalized, final Frame earthFrame, final double rotationRate, final DSSTPropagator dsstProp) throws IOException, OrekitException {
    final double ae = unnormalized.getAe();
    final int degree = parser.getInt(ParameterKey.CENTRAL_BODY_DEGREE);
    final int order = parser.getInt(ParameterKey.CENTRAL_BODY_ORDER);
    if (order > degree) {
        throw new IOException("Potential order cannot be higher than potential degree");
    }
    // Central Body Force Model with un-normalized coefficients
    dsstProp.addForceModel(new DSSTZonal(unnormalized, parser.getInt(ParameterKey.MAX_DEGREE_ZONAL_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_ECCENTRICITY_POWER_ZONAL_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_FREQUENCY_TRUE_LONGITUDE_ZONAL_SHORT_PERIODS)));
    dsstProp.addForceModel(new DSSTTesseral(earthFrame, rotationRate, unnormalized, parser.getInt(ParameterKey.MAX_DEGREE_TESSERAL_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_ORDER_TESSERAL_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_ECCENTRICITY_POWER_TESSERAL_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_FREQUENCY_MEAN_LONGITUDE_TESSERAL_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_DEGREE_TESSERAL_M_DAILIES_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_ORDER_TESSERAL_M_DAILIES_SHORT_PERIODS), parser.getInt(ParameterKey.MAX_ECCENTRICITY_POWER_TESSERAL_M_DAILIES_SHORT_PERIODS)));
    // 3rd body (SUN)
    if (parser.containsKey(ParameterKey.THIRD_BODY_SUN) && parser.getBoolean(ParameterKey.THIRD_BODY_SUN)) {
        dsstProp.addForceModel(new DSSTThirdBody(CelestialBodyFactory.getSun()));
    }
    // 3rd body (MOON)
    if (parser.containsKey(ParameterKey.THIRD_BODY_MOON) && parser.getBoolean(ParameterKey.THIRD_BODY_MOON)) {
        dsstProp.addForceModel(new DSSTThirdBody(CelestialBodyFactory.getMoon()));
    }
    // Drag
    if (parser.containsKey(ParameterKey.DRAG) && parser.getBoolean(ParameterKey.DRAG)) {
        final OneAxisEllipsoid earth = new OneAxisEllipsoid(ae, Constants.WGS84_EARTH_FLATTENING, earthFrame);
        final Atmosphere atm = new HarrisPriester(CelestialBodyFactory.getSun(), earth, 6);
        dsstProp.addForceModel(new DSSTAtmosphericDrag(atm, parser.getDouble(ParameterKey.DRAG_CD), parser.getDouble(ParameterKey.DRAG_SF)));
    }
    // Solar Radiation Pressure
    if (parser.containsKey(ParameterKey.SOLAR_RADIATION_PRESSURE) && parser.getBoolean(ParameterKey.SOLAR_RADIATION_PRESSURE)) {
        dsstProp.addForceModel(new DSSTSolarRadiationPressure(parser.getDouble(ParameterKey.SOLAR_RADIATION_PRESSURE_CR), parser.getDouble(ParameterKey.SOLAR_RADIATION_PRESSURE_SF), CelestialBodyFactory.getSun(), ae));
    }
}

Example 5 with DSSTAtmosphericDrag

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

the class DSSTPropagatorTest method testIssue157.

@Test
public void testIssue157() throws OrekitException {
    Utils.setDataRoot("regular-data:potential/icgem-format");
    GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("^eigen-6s-truncated$", false));
    UnnormalizedSphericalHarmonicsProvider nshp = GravityFieldFactory.getUnnormalizedProvider(8, 8);
    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, true);
    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, 8, 7, 17));
    propagator.addForceModel(new DSSTTesseral(earth.getBodyFrame(), Constants.WGS84_EARTH_ANGULAR_VELOCITY, nshp, 8, 8, 4, 12, 8, 8, 4));
    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()));
    propagator.setInitialState(new SpacecraftState(orbit, 45.0), true);
    SpacecraftState finalState = propagator.propagate(orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY));
    // the following comparison is in fact meaningless
    // the initial orbit is osculating the final orbit is a mean orbit
    // and they are not considered at the same epoch
    // we keep it only as is was an historical test
    Assert.assertEquals(2189.4, orbit.getA() - finalState.getA(), 1.0);
    propagator.setInitialState(new SpacecraftState(orbit, 45.0), false);
    finalState = propagator.propagate(orbit.getDate().shiftedBy(30 * Constants.JULIAN_DAY));
    // the following comparison is realistic
    // both the initial orbit and final orbit are mean orbits
    Assert.assertEquals(1478.05, orbit.getA() - finalState.getA(), 1.0);
}