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

Example 31 with KeplerianOrbit

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

the class DSSTPropagatorTest method testPropagationWithThirdBody.

@Test
public void testPropagationWithThirdBody() throws OrekitException, IOException {
    // Central Body geopotential 2x0
    final UnnormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getUnnormalizedProvider(2, 0);
    final Frame earthFrame = CelestialBodyFactory.getEarth().getBodyOrientedFrame();
    DSSTForceModel zonal = new DSSTZonal(provider, 2, 1, 5);
    DSSTForceModel tesseral = new DSSTTesseral(earthFrame, Constants.WGS84_EARTH_ANGULAR_VELOCITY, provider, 2, 0, 0, 2, 2, 0, 0);
    // Third Bodies Force Model (Moon + Sun)
    DSSTForceModel moon = new DSSTThirdBody(CelestialBodyFactory.getMoon());
    DSSTForceModel sun = new DSSTThirdBody(CelestialBodyFactory.getSun());
    // SIRIUS Orbit
    final AbsoluteDate initDate = new AbsoluteDate(2003, 7, 1, 0, 0, 00.000, TimeScalesFactory.getUTC());
    final Orbit orbit = new KeplerianOrbit(42163393., 0.2684, FastMath.toRadians(63.435), FastMath.toRadians(270.0), FastMath.toRadians(285.0), FastMath.toRadians(344.0), PositionAngle.MEAN, FramesFactory.getEME2000(), initDate, provider.getMu());
    // Set propagator with state and force model
    setDSSTProp(new SpacecraftState(orbit));
    dsstProp.addForceModel(zonal);
    dsstProp.addForceModel(tesseral);
    dsstProp.addForceModel(moon);
    dsstProp.addForceModel(sun);
    // 5 days propagation
    final SpacecraftState state = dsstProp.propagate(initDate.shiftedBy(5. * 86400.));
    // Ref Standalone_DSST:
    // a    = 42163393.0 m
    // h/ey =  -0.06893353670734315
    // k/ex =  -0.2592789733084587
    // p/hy =  -0.5968524904937771
    // q/hx =   0.1595005111738418
    // lM   = 183°9386620425922
    Assert.assertEquals(42163393.0, state.getA(), 1.e-1);
    Assert.assertEquals(-0.2592789733084587, state.getEquinoctialEx(), 5.e-7);
    Assert.assertEquals(-0.06893353670734315, state.getEquinoctialEy(), 2.e-7);
    Assert.assertEquals(0.1595005111738418, state.getHx(), 2.e-7);
    Assert.assertEquals(-0.5968524904937771, state.getHy(), 5.e-8);
    Assert.assertEquals(183.9386620425922, FastMath.toDegrees(MathUtils.normalizeAngle(state.getLM(), FastMath.PI)), 3.e-2);
}

Example 32 with KeplerianOrbit

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

the class DSSTPropagatorTest method testPropagationWithCentralBody.

@Test
public void testPropagationWithCentralBody() throws Exception {
    // Central Body geopotential 4x4
    final UnnormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getUnnormalizedProvider(4, 4);
    final Frame earthFrame = CelestialBodyFactory.getEarth().getBodyOrientedFrame();
    // GPS Orbit
    final AbsoluteDate initDate = new AbsoluteDate(2007, 4, 16, 0, 46, 42.400, TimeScalesFactory.getUTC());
    final Orbit orbit = new KeplerianOrbit(26559890., 0.0041632, FastMath.toRadians(55.2), FastMath.toRadians(315.4985), FastMath.toRadians(130.7562), FastMath.toRadians(44.2377), PositionAngle.MEAN, FramesFactory.getEME2000(), initDate, provider.getMu());
    // Set propagator with state and force model
    setDSSTProp(new SpacecraftState(orbit));
    dsstProp.addForceModel(new DSSTZonal(provider, 4, 3, 9));
    dsstProp.addForceModel(new DSSTTesseral(earthFrame, Constants.WGS84_EARTH_ANGULAR_VELOCITY, provider, 4, 4, 4, 8, 4, 4, 2));
    // 5 days propagation
    final SpacecraftState state = dsstProp.propagate(initDate.shiftedBy(5. * 86400.));
    // Ref GTDS_DSST:
    // a    = 26559.92081 km
    // h/ey =   0.2731622444E-03
    // k/ex =   0.4164167597E-02
    // p/hy =  -0.3399607878
    // q/hx =   0.3971568634
    // lM   = 140.6375352°
    Assert.assertEquals(26559920.81, state.getA(), 1.e-1);
    Assert.assertEquals(0.2731622444E-03, state.getEquinoctialEx(), 2.e-8);
    Assert.assertEquals(0.4164167597E-02, state.getEquinoctialEy(), 2.e-8);
    Assert.assertEquals(-0.3399607878, state.getHx(), 5.e-8);
    Assert.assertEquals(0.3971568634, state.getHy(), 2.e-6);
    Assert.assertEquals(140.6375352, FastMath.toDegrees(MathUtils.normalizeAngle(state.getLM(), FastMath.PI)), 5.e-3);
}

Example 33 with KeplerianOrbit

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

the class DSSTPropagatorTest method testIssueMeanInclination.

@Test
public void testIssueMeanInclination() throws OrekitException {
    final double earthAe = 6378137.0;
    final double earthMu = 3.9860044E14;
    final double earthJ2 = 0.0010826;
    // Initialize the DSST propagator with only J2 perturbation
    Orbit orb = new KeplerianOrbit(new TimeStampedPVCoordinates(new AbsoluteDate("1992-10-08T15:20:38.821", TimeScalesFactory.getUTC()), new Vector3D(5392808.809823, -4187618.3357927715, -44206.638015847195), new Vector3D(2337.4472786270794, 2474.0146611860464, 6778.507766114648)), FramesFactory.getTOD(false), earthMu);
    final SpacecraftState ss = new SpacecraftState(orb);
    final UnnormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getUnnormalizedProvider(earthAe, earthMu, TideSystem.UNKNOWN, new double[][] { { 0.0 }, { 0.0 }, { -earthJ2 } }, new double[][] { { 0.0 }, { 0.0 }, { 0.0 } });
    final Frame earthFrame = CelestialBodyFactory.getEarth().getBodyOrientedFrame();
    DSSTForceModel zonal = new DSSTZonal(provider, 2, 1, 5);
    DSSTForceModel tesseral = new DSSTTesseral(earthFrame, Constants.WGS84_EARTH_ANGULAR_VELOCITY, provider, 2, 0, 0, 2, 2, 0, 0);
    final Collection<DSSTForceModel> forces = new ArrayList<DSSTForceModel>();
    forces.add(zonal);
    forces.add(tesseral);
    // Computes J2 mean elements using the DSST osculating to mean converter
    final Orbit meanOrb = DSSTPropagator.computeMeanState(ss, null, forces).getOrbit();
    Assert.assertEquals(0.0164196, FastMath.toDegrees(orb.getI() - meanOrb.getI()), 1.0e-7);
}

Example 34 with KeplerianOrbit

use of org.orekit.orbits.KeplerianOrbit 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 35 with KeplerianOrbit

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

the class DSSTPropagatorTest method testPropagationWithSolarRadiationPressure.

@Test
public void testPropagationWithSolarRadiationPressure() throws OrekitException {
    // Central Body geopotential 2x0
    final UnnormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getUnnormalizedProvider(2, 0);
    DSSTForceModel zonal = new DSSTZonal(provider, 2, 1, 5);
    DSSTForceModel tesseral = new DSSTTesseral(CelestialBodyFactory.getEarth().getBodyOrientedFrame(), Constants.WGS84_EARTH_ANGULAR_VELOCITY, provider, 2, 0, 0, 2, 2, 0, 0);
    // SRP Force Model
    DSSTForceModel srp = new DSSTSolarRadiationPressure(1.2, 100., CelestialBodyFactory.getSun(), Constants.WGS84_EARTH_EQUATORIAL_RADIUS);
    // GEO Orbit
    final AbsoluteDate initDate = new AbsoluteDate(2003, 9, 16, 0, 0, 00.000, TimeScalesFactory.getUTC());
    final Orbit orbit = new KeplerianOrbit(42166258., 0.0001, FastMath.toRadians(0.001), FastMath.toRadians(315.4985), FastMath.toRadians(130.7562), FastMath.toRadians(44.2377), PositionAngle.MEAN, FramesFactory.getGCRF(), initDate, provider.getMu());
    // Set propagator with state and force model
    dsstProp = new DSSTPropagator(new ClassicalRungeKuttaIntegrator(86400.));
    dsstProp.setInitialState(new SpacecraftState(orbit), false);
    dsstProp.addForceModel(zonal);
    dsstProp.addForceModel(tesseral);
    dsstProp.addForceModel(srp);
    // 10 days propagation
    final SpacecraftState state = dsstProp.propagate(initDate.shiftedBy(10. * 86400.));
    // Ref Standalone_DSST:
    // a    = 42166257.99807995 m
    // h/ey = -0.1191876027555493D-03
    // k/ex = -0.1781865038201885D-05
    // p/hy =  0.6618387121369373D-05
    // q/hx = -0.5624363171289686D-05
    // lM   = 140°3496229467104
    Assert.assertEquals(42166257.99807995, state.getA(), 0.8);
    Assert.assertEquals(-0.1781865038201885e-05, state.getEquinoctialEx(), 3.e-7);
    Assert.assertEquals(-0.1191876027555493e-03, state.getEquinoctialEy(), 4.e-6);
    Assert.assertEquals(-0.5624363171289686e-05, state.getHx(), 4.e-9);
    Assert.assertEquals(0.6618387121369373e-05, state.getHy(), 3.e-10);
    Assert.assertEquals(140.3496229467104, FastMath.toDegrees(MathUtils.normalizeAngle(state.getLM(), FastMath.PI)), 2.e-4);
}