private double[] stateToArray(
SpacecraftState state, OrbitType orbitType, PositionAngle angleType, boolean withMass) {
double[] array = new double[withMass ? 7 : 6];
switch (orbitType) {
case CARTESIAN:
{
CartesianOrbit cart = (CartesianOrbit) orbitType.convertType(state.getOrbit());
array[0] = cart.getPVCoordinates().getPosition().getX();
array[1] = cart.getPVCoordinates().getPosition().getY();
array[2] = cart.getPVCoordinates().getPosition().getZ();
array[3] = cart.getPVCoordinates().getVelocity().getX();
array[4] = cart.getPVCoordinates().getVelocity().getY();
array[5] = cart.getPVCoordinates().getVelocity().getZ();
}
break;
case CIRCULAR:
{
CircularOrbit circ = (CircularOrbit) orbitType.convertType(state.getOrbit());
array[0] = circ.getA();
array[1] = circ.getCircularEx();
array[2] = circ.getCircularEy();
array[3] = circ.getI();
array[4] = circ.getRightAscensionOfAscendingNode();
array[5] = circ.getAlpha(angleType);
}
break;
case EQUINOCTIAL:
{
EquinoctialOrbit equ = (EquinoctialOrbit) orbitType.convertType(state.getOrbit());
array[0] = equ.getA();
array[1] = equ.getEquinoctialEx();
array[2] = equ.getEquinoctialEy();
array[3] = equ.getHx();
array[4] = equ.getHy();
array[5] = equ.getL(angleType);
}
break;
case KEPLERIAN:
{
KeplerianOrbit kep = (KeplerianOrbit) orbitType.convertType(state.getOrbit());
array[0] = kep.getA();
array[1] = kep.getE();
array[2] = kep.getI();
array[3] = kep.getPerigeeArgument();
array[4] = kep.getRightAscensionOfAscendingNode();
array[5] = kep.getAnomaly(angleType);
}
break;
}
if (withMass) {
array[6] = state.getMass();
}
return array;
}
@Test
public void testPropagationTypesElliptical() throws OrekitException, ParseException, IOException {
NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(5, 5);
ForceModel gravityField =
new HolmesFeatherstoneAttractionModel(
FramesFactory.getITRF(IERSConventions.IERS_2010, true), provider);
Orbit initialOrbit =
new KeplerianOrbit(
8000000.0,
0.01,
0.1,
0.7,
0,
1.2,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
AbsoluteDate.J2000_EPOCH,
provider.getMu());
double dt = 900;
double dP = 0.001;
for (OrbitType orbitType : OrbitType.values()) {
for (PositionAngle angleType : PositionAngle.values()) {
// compute state Jacobian using PartialDerivatives
NumericalPropagator propagator =
setUpPropagator(initialOrbit, dP, orbitType, angleType, gravityField);
PartialDerivativesEquations partials =
new PartialDerivativesEquations("partials", propagator);
final SpacecraftState initialState =
partials.setInitialJacobians(new SpacecraftState(initialOrbit), 6, 0);
propagator.setInitialState(initialState);
final JacobiansMapper mapper = partials.getMapper();
PickUpHandler pickUp = new PickUpHandler(mapper, null);
propagator.setMasterMode(pickUp);
propagator.propagate(initialState.getDate().shiftedBy(dt));
double[][] dYdY0 = pickUp.getdYdY0();
// compute reference state Jacobian using finite differences
double[][] dYdY0Ref = new double[6][6];
AbstractIntegratedPropagator propagator2 =
setUpPropagator(initialOrbit, dP, orbitType, angleType, gravityField);
double[] steps = NumericalPropagator.tolerances(1000000 * dP, initialOrbit, orbitType)[0];
for (int i = 0; i < 6; ++i) {
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -4 * steps[i], i));
SpacecraftState sM4h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -3 * steps[i], i));
SpacecraftState sM3h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -2 * steps[i], i));
SpacecraftState sM2h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -1 * steps[i], i));
SpacecraftState sM1h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 1 * steps[i], i));
SpacecraftState sP1h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 2 * steps[i], i));
SpacecraftState sP2h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 3 * steps[i], i));
SpacecraftState sP3h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 4 * steps[i], i));
SpacecraftState sP4h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
fillJacobianColumn(
dYdY0Ref, i, orbitType, angleType, steps[i], sM4h, sM3h, sM2h, sM1h, sP1h, sP2h, sP3h,
sP4h);
}
for (int i = 0; i < 6; ++i) {
for (int j = 0; j < 6; ++j) {
double error = FastMath.abs((dYdY0[i][j] - dYdY0Ref[i][j]) / dYdY0Ref[i][j]);
Assert.assertEquals(0, error, 6.0e-2);
}
}
}
}
}
