@Test
public void testEventDetectionBug() throws OrekitException, IOException, ParseException {
TimeScale utc = TimeScalesFactory.getUTC();
AbsoluteDate initialDate = new AbsoluteDate(2005, 1, 1, 0, 0, 0.0, utc);
double duration = 100000.;
AbsoluteDate endDate = new AbsoluteDate(initialDate, duration);
// Initialization of the frame EME2000
Frame EME2000 = FramesFactory.getEME2000();
// Initial orbit
double a = 35786000. + 6378137.0;
double e = 0.70;
double rApogee = a * (1 + e);
double vApogee = FastMath.sqrt(mu * (1 - e) / (a * (1 + e)));
Orbit geo =
new CartesianOrbit(
new PVCoordinates(new Vector3D(rApogee, 0., 0.), new Vector3D(0., vApogee, 0.)),
EME2000,
initialDate,
mu);
duration = geo.getKeplerianPeriod();
endDate = new AbsoluteDate(initialDate, duration);
// Numerical Integration
final double minStep = 0.001;
final double maxStep = 1000;
final double initStep = 60;
final double[] absTolerance = {0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001};
final double[] relTolerance = {1.0e-7, 1.0e-4, 1.0e-4, 1.0e-7, 1.0e-7, 1.0e-7, 1.0e-7};
AdaptiveStepsizeIntegrator integrator =
new DormandPrince853Integrator(minStep, maxStep, absTolerance, relTolerance);
integrator.setInitialStepSize(initStep);
// Numerical propagator based on the integrator
propagator = new NumericalPropagator(integrator);
double mass = 1000.;
SpacecraftState initialState = new SpacecraftState(geo, mass);
propagator.setInitialState(initialState);
propagator.setOrbitType(OrbitType.CARTESIAN);
// Set the events Detectors
ApsideDetector event1 = new ApsideDetector(geo);
propagator.addEventDetector(event1);
// Set the propagation mode
propagator.setSlaveMode();
// Propagate
SpacecraftState finalState = propagator.propagate(endDate);
// we should stop long before endDate
Assert.assertTrue(endDate.durationFrom(finalState.getDate()) > 40000.0);
}
@Before
public void setUp() throws OrekitException {
Utils.setDataRoot("regular-data:potential/shm-format");
GravityFieldFactory.addPotentialCoefficientsReader(
new SHMFormatReader("^eigen_cg03c_coef$", false));
mu = GravityFieldFactory.getUnnormalizedProvider(0, 0).getMu();
final Vector3D position = new Vector3D(7.0e6, 1.0e6, 4.0e6);
final Vector3D velocity = new Vector3D(-500.0, 8000.0, 1000.0);
initDate = AbsoluteDate.J2000_EPOCH;
final Orbit orbit =
new EquinoctialOrbit(
new PVCoordinates(position, velocity), FramesFactory.getEME2000(), initDate, mu);
initialState = new SpacecraftState(orbit);
double[][] tolerance = NumericalPropagator.tolerances(0.001, orbit, OrbitType.EQUINOCTIAL);
AdaptiveStepsizeIntegrator integrator =
new DormandPrince853Integrator(0.001, 200, tolerance[0], tolerance[1]);
integrator.setInitialStepSize(60);
propagator = new NumericalPropagator(integrator);
propagator.setInitialState(initialState);
}
@Test
public void testJacobianIssue18() throws OrekitException {
// Body mu
final double mu = 3.9860047e14;
final double isp = 318;
final double mass = 2500;
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 double duration = 3653.99;
final double f = 420;
final double delta = FastMath.toRadians(-7.4978);
final double alpha = FastMath.toRadians(351);
final AttitudeProvider law =
new InertialProvider(new Rotation(new Vector3D(alpha, delta), Vector3D.PLUS_I));
final AbsoluteDate initDate =
new AbsoluteDate(
new DateComponents(2004, 01, 01),
new TimeComponents(23, 30, 00.000),
TimeScalesFactory.getUTC());
final Orbit orbit =
new KeplerianOrbit(
a,
e,
i,
omega,
OMEGA,
lv,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
initDate,
mu);
final SpacecraftState initialState =
new SpacecraftState(orbit, law.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass);
final AbsoluteDate fireDate =
new AbsoluteDate(
new DateComponents(2004, 01, 02),
new TimeComponents(04, 15, 34.080),
TimeScalesFactory.getUTC());
final ConstantThrustManeuver maneuver =
new ConstantThrustManeuver(fireDate, duration, f, isp, Vector3D.PLUS_I);
double[] absTolerance = {0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001};
double[] relTolerance = {1.0e-7, 1.0e-4, 1.0e-4, 1.0e-7, 1.0e-7, 1.0e-7, 1.0e-7};
AdaptiveStepsizeIntegrator integrator =
new DormandPrince853Integrator(0.001, 1000, absTolerance, relTolerance);
integrator.setInitialStepSize(60);
final NumericalPropagator propagator = new NumericalPropagator(integrator);
propagator.setAttitudeProvider(law);
propagator.addForceModel(maneuver);
propagator.setOrbitType(OrbitType.CARTESIAN);
PartialDerivativesEquations PDE = new PartialDerivativesEquations("derivatives", propagator);
PDE.selectParamAndStep("thrust", Double.NaN);
Assert.assertEquals(3, PDE.getAvailableParameters().size());
Assert.assertEquals("central attraction coefficient", PDE.getAvailableParameters().get(0));
Assert.assertEquals("thrust", PDE.getAvailableParameters().get(1));
Assert.assertEquals("flow rate", PDE.getAvailableParameters().get(2));
propagator.setInitialState(PDE.setInitialJacobians(initialState, 7, 1));
final AbsoluteDate finalDate = fireDate.shiftedBy(3800);
final SpacecraftState finalorb = propagator.propagate(finalDate);
Assert.assertEquals(0, finalDate.durationFrom(finalorb.getDate()), 1.0e-11);
}