@Test
public void testUnsupportedParameter() throws OrekitException {
Orbit initialOrbit =
new KeplerianOrbit(
8000000.0,
0.01,
0.1,
0.7,
0,
1.2,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
AbsoluteDate.J2000_EPOCH,
Constants.EIGEN5C_EARTH_MU);
double dP = 0.001;
NumericalPropagator propagator =
setUpPropagator(
initialOrbit,
dP,
OrbitType.EQUINOCTIAL,
PositionAngle.TRUE,
new ThirdBodyAttraction(CelestialBodyFactory.getSun()),
new ThirdBodyAttraction(CelestialBodyFactory.getMoon()));
PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
partials.selectParamAndStep("non-existent", 1.0);
try {
partials.computeDerivatives(new SpacecraftState(initialOrbit), new double[6]);
Assert.fail("an exception should have been thrown");
} catch (OrekitException oe) {
Assert.assertEquals(OrekitMessages.UNSUPPORTED_PARAMETER_NAME, oe.getSpecifier());
}
}
@Test(expected = OrekitException.class)
public void testMismatchedDimensions() throws OrekitException {
Orbit initialOrbit =
new KeplerianOrbit(
8000000.0,
0.01,
0.1,
0.7,
0,
1.2,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
AbsoluteDate.J2000_EPOCH,
Constants.EIGEN5C_EARTH_MU);
double dP = 0.001;
NumericalPropagator propagator =
setUpPropagator(initialOrbit, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE);
PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
partials.setInitialJacobians(
new SpacecraftState(initialOrbit), new double[6][6], new double[7][2]);
}
@Test
public void testWrongParametersDimension() throws OrekitException {
Orbit initialOrbit =
new KeplerianOrbit(
8000000.0,
0.01,
0.1,
0.7,
0,
1.2,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
AbsoluteDate.J2000_EPOCH,
Constants.EIGEN5C_EARTH_MU);
double dP = 0.001;
NumericalPropagator propagator =
setUpPropagator(
initialOrbit,
dP,
OrbitType.EQUINOCTIAL,
PositionAngle.TRUE,
new ThirdBodyAttraction(CelestialBodyFactory.getSun()),
new ThirdBodyAttraction(CelestialBodyFactory.getMoon()));
PartialDerivativesEquations partials = new PartialDerivativesEquations("partials", propagator);
partials.setInitialJacobians(
new SpacecraftState(initialOrbit), new double[6][6], new double[6][3]);
partials.selectParameters("Sun attraction coefficient");
try {
partials.computeDerivatives(new SpacecraftState(initialOrbit), new double[6]);
Assert.fail("an exception should have been thrown");
} catch (OrekitException oe) {
Assert.assertEquals(
OrekitMessages.INITIAL_MATRIX_AND_PARAMETERS_NUMBER_MISMATCH, oe.getSpecifier());
}
}
@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);
}
}
}
}
}
@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);
}