use of org.orekit.time.DateComponents in project Orekit by CS-SI.
the class HarmonicParametricAccelerationTest method testCoefficientsDetermination.
@Test
public void testCoefficientsDetermination() throws OrekitException {
final double mass = 2500;
final Orbit orbit = new CircularOrbit(7500000.0, 1.0e-4, 1.0e-3, 1.7, 0.3, 0.5, PositionAngle.TRUE, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 2, 3), TimeComponents.H00, TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU);
final double period = orbit.getKeplerianPeriod();
AttitudeProvider maneuverLaw = new LofOffset(orbit.getFrame(), LOFType.VNC);
SpacecraftState initialState = new SpacecraftState(orbit, maneuverLaw.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass);
double dP = 10.0;
double minStep = 0.001;
double maxStep = 100;
double[][] tolerance = NumericalPropagator.tolerances(dP, orbit, orbit.getType());
// generate PV measurements corresponding to a tangential maneuver
AdaptiveStepsizeIntegrator integrator0 = new DormandPrince853Integrator(minStep, maxStep, tolerance[0], tolerance[1]);
integrator0.setInitialStepSize(60);
final NumericalPropagator propagator0 = new NumericalPropagator(integrator0);
propagator0.setInitialState(initialState);
propagator0.setAttitudeProvider(maneuverLaw);
ForceModel hpaRefX1 = new HarmonicParametricAcceleration(Vector3D.PLUS_I, true, "refX1", null, period, 1);
ForceModel hpaRefY1 = new HarmonicParametricAcceleration(Vector3D.PLUS_J, true, "refY1", null, period, 1);
ForceModel hpaRefZ2 = new HarmonicParametricAcceleration(Vector3D.PLUS_K, true, "refZ2", null, period, 2);
hpaRefX1.getParametersDrivers()[0].setValue(2.4e-2);
hpaRefX1.getParametersDrivers()[1].setValue(3.1);
hpaRefY1.getParametersDrivers()[0].setValue(4.0e-2);
hpaRefY1.getParametersDrivers()[1].setValue(0.3);
hpaRefZ2.getParametersDrivers()[0].setValue(1.0e-2);
hpaRefZ2.getParametersDrivers()[1].setValue(1.8);
propagator0.addForceModel(hpaRefX1);
propagator0.addForceModel(hpaRefY1);
propagator0.addForceModel(hpaRefZ2);
final List<ObservedMeasurement<?>> measurements = new ArrayList<>();
propagator0.setMasterMode(10.0, (state, isLast) -> measurements.add(new PV(state.getDate(), state.getPVCoordinates().getPosition(), state.getPVCoordinates().getVelocity(), 1.0e-3, 1.0e-6, 1.0)));
propagator0.propagate(orbit.getDate().shiftedBy(900));
// set up an estimator to retrieve the maneuver as several harmonic accelerations in inertial frame
final NumericalPropagatorBuilder propagatorBuilder = new NumericalPropagatorBuilder(orbit, new DormandPrince853IntegratorBuilder(minStep, maxStep, dP), PositionAngle.TRUE, dP);
propagatorBuilder.addForceModel(new HarmonicParametricAcceleration(Vector3D.PLUS_I, true, "X1", null, period, 1));
propagatorBuilder.addForceModel(new HarmonicParametricAcceleration(Vector3D.PLUS_J, true, "Y1", null, period, 1));
propagatorBuilder.addForceModel(new HarmonicParametricAcceleration(Vector3D.PLUS_K, true, "Z2", null, period, 2));
final BatchLSEstimator estimator = new BatchLSEstimator(new LevenbergMarquardtOptimizer(), propagatorBuilder);
estimator.setParametersConvergenceThreshold(1.0e-2);
estimator.setMaxIterations(20);
estimator.setMaxEvaluations(100);
for (final ObservedMeasurement<?> measurement : measurements) {
estimator.addMeasurement(measurement);
}
// we will estimate only the force model parameters, not the orbit
for (final ParameterDriver d : estimator.getOrbitalParametersDrivers(false).getDrivers()) {
d.setSelected(false);
}
setParameter(estimator, "X1 γ", 1.0e-2);
setParameter(estimator, "X1 φ", 4.0);
setParameter(estimator, "Y1 γ", 1.0e-2);
setParameter(estimator, "Y1 φ", 0.0);
setParameter(estimator, "Z2 γ", 1.0e-2);
setParameter(estimator, "Z2 φ", 1.0);
estimator.estimate();
Assert.assertTrue(estimator.getIterationsCount() < 15);
Assert.assertTrue(estimator.getEvaluationsCount() < 15);
Assert.assertEquals(0.0, estimator.getOptimum().getRMS(), 1.0e-5);
Assert.assertEquals(hpaRefX1.getParametersDrivers()[0].getValue(), getParameter(estimator, "X1 γ"), 1.e-12);
Assert.assertEquals(hpaRefX1.getParametersDrivers()[1].getValue(), getParameter(estimator, "X1 φ"), 1.e-12);
Assert.assertEquals(hpaRefY1.getParametersDrivers()[0].getValue(), getParameter(estimator, "Y1 γ"), 1.e-12);
Assert.assertEquals(hpaRefY1.getParametersDrivers()[1].getValue(), getParameter(estimator, "Y1 φ"), 1.e-12);
Assert.assertEquals(hpaRefZ2.getParametersDrivers()[0].getValue(), getParameter(estimator, "Z2 γ"), 1.e-12);
Assert.assertEquals(hpaRefZ2.getParametersDrivers()[1].getValue(), getParameter(estimator, "Z2 φ"), 1.e-12);
}
use of org.orekit.time.DateComponents in project Orekit by CS-SI.
the class PolynomialParametricAccelerationTest method setUp.
@Before
public void setUp() {
try {
Utils.setDataRoot("regular-data");
final double a = 24396159;
final double e = 0.72831215;
final double i = FastMath.toRadians(7);
final double omega = FastMath.toRadians(180);
final double OMEGA = FastMath.toRadians(261);
final double lv = 0;
final AbsoluteDate initDate = new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC());
initialOrbit = new KeplerianOrbit(a, e, i, omega, OMEGA, lv, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, Constants.EIGEN5C_EARTH_MU);
} catch (OrekitException oe) {
Assert.fail(oe.getLocalizedMessage());
}
}
use of org.orekit.time.DateComponents in project Orekit by CS-SI.
the class DragForceTest method testStateJacobianSphere.
@Test
public void testStateJacobianSphere() throws OrekitException {
// initialization
AbsoluteDate date = new AbsoluteDate(new DateComponents(2003, 03, 01), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
double i = FastMath.toRadians(98.7);
double omega = FastMath.toRadians(93.0);
double OMEGA = FastMath.toRadians(15.0 * 22.5);
Orbit orbit = new KeplerianOrbit(7201009.7124401, 1e-3, i, omega, OMEGA, 0, PositionAngle.MEAN, FramesFactory.getEME2000(), date, Constants.EIGEN5C_EARTH_MU);
OrbitType integrationType = OrbitType.CARTESIAN;
double[][] tolerances = NumericalPropagator.tolerances(0.01, orbit, integrationType);
NumericalPropagator propagator = new NumericalPropagator(new DormandPrince853Integrator(1.0e-3, 120, tolerances[0], tolerances[1]));
propagator.setOrbitType(integrationType);
final DragForce forceModel = new DragForce(new HarrisPriester(CelestialBodyFactory.getSun(), new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true))), new IsotropicDrag(2.5, 1.2));
propagator.addForceModel(forceModel);
SpacecraftState state0 = new SpacecraftState(orbit);
checkStateJacobian(propagator, state0, date.shiftedBy(3.5 * 3600.0), 1e3, tolerances[0], 2.0e-8);
}
use of org.orekit.time.DateComponents in project Orekit by CS-SI.
the class DragForceTest method testJacobianBoxVsFiniteDifferences.
@Test
public void testJacobianBoxVsFiniteDifferences() throws OrekitException {
// initialization
AbsoluteDate date = new AbsoluteDate(new DateComponents(2003, 03, 01), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
double i = FastMath.toRadians(98.7);
double omega = FastMath.toRadians(93.0);
double OMEGA = FastMath.toRadians(15.0 * 22.5);
Orbit orbit = new KeplerianOrbit(7201009.7124401, 1e-3, i, omega, OMEGA, 0, PositionAngle.MEAN, FramesFactory.getEME2000(), date, Constants.EIGEN5C_EARTH_MU);
final DragForce forceModel = new DragForce(new HarrisPriester(CelestialBodyFactory.getSun(), new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true))), new BoxAndSolarArraySpacecraft(1.5, 2.0, 1.8, CelestialBodyFactory.getSun(), 20.0, Vector3D.PLUS_J, 1.2, 0.7, 0.2));
SpacecraftState state = new SpacecraftState(orbit, Propagator.DEFAULT_LAW.getAttitude(orbit, orbit.getDate(), orbit.getFrame()));
checkStateJacobianVsFiniteDifferences(state, forceModel, Propagator.DEFAULT_LAW, 1.0, 5.0e-6, false);
}
use of org.orekit.time.DateComponents in project Orekit by CS-SI.
the class DragForceTest method testGlobalStateJacobianBox.
@Test
public void testGlobalStateJacobianBox() throws OrekitException {
// initialization
AbsoluteDate date = new AbsoluteDate(new DateComponents(2003, 03, 01), new TimeComponents(13, 59, 27.816), TimeScalesFactory.getUTC());
double i = FastMath.toRadians(98.7);
double omega = FastMath.toRadians(93.0);
double OMEGA = FastMath.toRadians(15.0 * 22.5);
Orbit orbit = new KeplerianOrbit(7201009.7124401, 1e-3, i, omega, OMEGA, 0, PositionAngle.MEAN, FramesFactory.getEME2000(), date, Constants.EIGEN5C_EARTH_MU);
OrbitType integrationType = OrbitType.CARTESIAN;
double[][] tolerances = NumericalPropagator.tolerances(0.01, orbit, integrationType);
NumericalPropagator propagator = new NumericalPropagator(new DormandPrince853Integrator(1.0e-3, 120, tolerances[0], tolerances[1]));
propagator.setOrbitType(integrationType);
final DragForce forceModel = new DragForce(new HarrisPriester(CelestialBodyFactory.getSun(), new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS, Constants.WGS84_EARTH_FLATTENING, FramesFactory.getITRF(IERSConventions.IERS_2010, true))), new BoxAndSolarArraySpacecraft(1.5, 2.0, 1.8, CelestialBodyFactory.getSun(), 20.0, Vector3D.PLUS_J, 1.2, 0.7, 0.2));
propagator.addForceModel(forceModel);
SpacecraftState state0 = new SpacecraftState(orbit);
checkStateJacobian(propagator, state0, date.shiftedBy(3.5 * 3600.0), 1e3, tolerances[0], 3.0e-8);
}
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