/**
* Conversion from osculating to mean, orbit.
*
* <p>Compute osculating state <b>in a DSST sense</b>, corresponding to the mean SpacecraftState
* in input, and according to the Force models taken into account.
*
* <p>Since the osculating state is obtained with the computation of short-periodic variation of
* each force model, the resulting output will depend on the force models parametrized in input.
*
* <p>The computing is done through a fixed-point iteration process.
*
* @param osculating Osculating state to convert
* @param forces Forces to take into account
* @return mean state in a DSST sense
* @throws OrekitException if computation of short periodics fails or iteration algorithm does not
* converge
*/
public static SpacecraftState computeMeanState(
final SpacecraftState osculating, final Collection<DSSTForceModel> forces)
throws OrekitException {
// Creation of a DSSTPropagator instance
final AbstractIntegrator integrator = new ClassicalRungeKuttaIntegrator(43200.);
final DSSTPropagator dsst = new DSSTPropagator(integrator, false);
// Create the auxiliary object
final AuxiliaryElements aux = new AuxiliaryElements(osculating.getOrbit(), I);
// Set the force models
for (final DSSTForceModel force : forces) {
force.initialize(aux, false);
dsst.addForceModel(force);
}
dsst.setInitialState(osculating, true);
final Orbit meanOrbit = dsst.mapper.computeMeanOrbit(osculating);
return new SpacecraftState(
meanOrbit,
osculating.getAttitude(),
osculating.getMass(),
osculating.getAdditionalStates());
}
@Test
public void testEphemerisAdditionalState() throws OrekitException, IOException {
// Propagation of the initial at t + dt
final double dt = -3200;
final double rate = 2.0;
propagator.addAdditionalStateProvider(
new AdditionalStateProvider() {
public String getName() {
return "squaredA";
}
public double[] getAdditionalState(SpacecraftState state) {
return new double[] {state.getA() * state.getA()};
}
});
propagator.addAdditionalEquations(
new AdditionalEquations() {
public String getName() {
return "extra";
}
public double[] computeDerivatives(SpacecraftState s, double[] pDot) {
pDot[0] = rate;
return null;
}
});
propagator.setInitialState(propagator.getInitialState().addAdditionalState("extra", 1.5));
propagator.setOrbitType(OrbitType.CARTESIAN);
propagator.setEphemerisMode();
propagator.propagate(initDate.shiftedBy(dt));
final BoundedPropagator ephemeris1 = propagator.getGeneratedEphemeris();
Assert.assertEquals(initDate.shiftedBy(dt), ephemeris1.getMinDate());
Assert.assertEquals(initDate, ephemeris1.getMaxDate());
try {
ephemeris1.propagate(ephemeris1.getMinDate().shiftedBy(-10.0));
Assert.fail("an exception should have been thrown");
} catch (PropagationException pe) {
Assert.assertEquals(OrekitMessages.OUT_OF_RANGE_EPHEMERIDES_DATE, pe.getSpecifier());
}
try {
ephemeris1.propagate(ephemeris1.getMaxDate().shiftedBy(+10.0));
Assert.fail("an exception should have been thrown");
} catch (PropagationException pe) {
Assert.assertEquals(OrekitMessages.OUT_OF_RANGE_EPHEMERIDES_DATE, pe.getSpecifier());
}
double shift = -60;
SpacecraftState s = ephemeris1.propagate(initDate.shiftedBy(shift));
Assert.assertEquals(2, s.getAdditionalStates().size());
Assert.assertTrue(s.hasAdditionalState("squaredA"));
Assert.assertTrue(s.hasAdditionalState("extra"));
Assert.assertEquals(s.getA() * s.getA(), s.getAdditionalState("squaredA")[0], 1.0e-10);
Assert.assertEquals(1.5 + shift * rate, s.getAdditionalState("extra")[0], 1.0e-10);
try {
ephemeris1.resetInitialState(s);
Assert.fail("an exception should have been thrown");
} catch (OrekitException oe) {
Assert.assertEquals(OrekitMessages.NON_RESETABLE_STATE, oe.getSpecifier());
}
}