@Test
public void testSpin() throws OrekitException {
AbsoluteDate date =
new AbsoluteDate(
new DateComponents(1970, 01, 01),
new TimeComponents(3, 25, 45.6789),
TimeScalesFactory.getUTC());
KeplerianOrbit orbit =
new KeplerianOrbit(
7178000.0,
1.e-4,
FastMath.toRadians(50.),
FastMath.toRadians(10.),
FastMath.toRadians(20.),
FastMath.toRadians(30.),
PositionAngle.MEAN,
FramesFactory.getEME2000(),
date,
3.986004415e14);
final AttitudeProvider law =
new LofOffsetPointing(
earthSpheric,
new LofOffset(orbit.getFrame(), LOFType.VVLH, RotationOrder.XYX, 0.1, 0.2, 0.3),
Vector3D.PLUS_K);
Propagator propagator = new KeplerianPropagator(orbit, law);
double h = 0.01;
SpacecraftState sMinus = propagator.propagate(date.shiftedBy(-h));
SpacecraftState s0 = propagator.propagate(date);
SpacecraftState sPlus = propagator.propagate(date.shiftedBy(h));
// check spin is consistent with attitude evolution
double errorAngleMinus =
Rotation.distance(
sMinus.shiftedBy(h).getAttitude().getRotation(), s0.getAttitude().getRotation());
double evolutionAngleMinus =
Rotation.distance(sMinus.getAttitude().getRotation(), s0.getAttitude().getRotation());
Assert.assertEquals(0.0, errorAngleMinus, 1.0e-6 * evolutionAngleMinus);
double errorAnglePlus =
Rotation.distance(
s0.getAttitude().getRotation(), sPlus.shiftedBy(-h).getAttitude().getRotation());
double evolutionAnglePlus =
Rotation.distance(s0.getAttitude().getRotation(), sPlus.getAttitude().getRotation());
Assert.assertEquals(0.0, errorAnglePlus, 1.0e-6 * evolutionAnglePlus);
Vector3D spin0 = s0.getAttitude().getSpin();
Vector3D reference =
AngularCoordinates.estimateRate(
sMinus.getAttitude().getRotation(), sPlus.getAttitude().getRotation(), 2 * h);
Assert.assertTrue(spin0.getNorm() > 1.0e-3);
Assert.assertEquals(0.0, spin0.subtract(reference).getNorm(), 1.0e-10);
}
@Test
public void testCartesian() throws OrekitException {
// Propagation of the initial at t + dt
final double dt = 3200;
propagator.setOrbitType(OrbitType.CARTESIAN);
final PVCoordinates finalState =
propagator.propagate(initDate.shiftedBy(dt)).getPVCoordinates();
final Vector3D pFin = finalState.getPosition();
final Vector3D vFin = finalState.getVelocity();
// Check results
final PVCoordinates reference = initialState.shiftedBy(dt).getPVCoordinates();
final Vector3D pRef = reference.getPosition();
final Vector3D vRef = reference.getVelocity();
Assert.assertEquals(0, pRef.subtract(pFin).getNorm(), 2e-4);
Assert.assertEquals(0, vRef.subtract(vFin).getNorm(), 7e-8);
try {
propagator.getGeneratedEphemeris();
Assert.fail("an exception should have been thrown");
} catch (IllegalStateException ise) {
// expected
}
}
/**
* Simple constructor.
*
* <p>The {@code applyBefore} parameter is mainly used when the differential effect is associated
* with a maneuver. In this case, the parameter must be set to {@code false}.
*
* @param original original state at reference date
* @param directEffect direct effect changing the orbit
* @param applyBefore if true, effect is applied both before and after reference date, if false it
* is only applied after reference date
* @param referenceRadius reference radius of the Earth for the potential model (m)
* @param mu central attraction coefficient (m³/s²)
* @param j2 un-normalized zonal coefficient (about +1.08e-3 for Earth)
* @exception OrekitException if direct effect cannot be applied
*/
public J2DifferentialEffect(
final SpacecraftState original,
final AdapterPropagator.DifferentialEffect directEffect,
final boolean applyBefore,
final double referenceRadius,
final double mu,
final double j2)
throws OrekitException {
this(
original.getOrbit(),
directEffect.apply(original.shiftedBy(0.001)).getOrbit().shiftedBy(-0.001),
applyBefore,
referenceRadius,
mu,
j2);
}
