@Test
public void testPropagationTypesHyperbolic() throws OrekitException, ParseException, IOException {
SpacecraftState state =
new SpacecraftState(
new KeplerianOrbit(
-10000000.0,
2.5,
0.3,
0,
0,
0.0,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
initDate,
mu));
ForceModel gravityField =
new HolmesFeatherstoneAttractionModel(
FramesFactory.getITRF(IERSConventions.IERS_2010, true),
GravityFieldFactory.getNormalizedProvider(5, 5));
propagator.addForceModel(gravityField);
// Propagation of the initial at t + dt
final PVCoordinates pv = state.getPVCoordinates();
final double dP = 0.001;
final double dV =
state.getMu() * dP / (pv.getPosition().getNormSq() * pv.getVelocity().getNorm());
final PVCoordinates pvcM = propagateInType(state, dP, OrbitType.CARTESIAN, PositionAngle.MEAN);
final PVCoordinates pvkM = propagateInType(state, dP, OrbitType.KEPLERIAN, PositionAngle.MEAN);
final PVCoordinates pvcE =
propagateInType(state, dP, OrbitType.CARTESIAN, PositionAngle.ECCENTRIC);
final PVCoordinates pvkE =
propagateInType(state, dP, OrbitType.KEPLERIAN, PositionAngle.ECCENTRIC);
final PVCoordinates pvcT = propagateInType(state, dP, OrbitType.CARTESIAN, PositionAngle.TRUE);
final PVCoordinates pvkT = propagateInType(state, dP, OrbitType.KEPLERIAN, PositionAngle.TRUE);
Assert.assertEquals(0, pvcM.getPosition().subtract(pvkT.getPosition()).getNorm() / dP, 0.3);
Assert.assertEquals(0, pvcM.getVelocity().subtract(pvkT.getVelocity()).getNorm() / dV, 0.4);
Assert.assertEquals(0, pvkM.getPosition().subtract(pvkT.getPosition()).getNorm() / dP, 0.2);
Assert.assertEquals(0, pvkM.getVelocity().subtract(pvkT.getVelocity()).getNorm() / dV, 0.3);
Assert.assertEquals(0, pvcE.getPosition().subtract(pvkT.getPosition()).getNorm() / dP, 0.3);
Assert.assertEquals(0, pvcE.getVelocity().subtract(pvkT.getVelocity()).getNorm() / dV, 0.4);
Assert.assertEquals(0, pvkE.getPosition().subtract(pvkT.getPosition()).getNorm() / dP, 0.009);
Assert.assertEquals(0, pvkE.getVelocity().subtract(pvkT.getVelocity()).getNorm() / dV, 0.006);
Assert.assertEquals(0, pvcT.getPosition().subtract(pvkT.getPosition()).getNorm() / dP, 0.3);
Assert.assertEquals(0, pvcT.getVelocity().subtract(pvkT.getVelocity()).getNorm() / dV, 0.4);
}
@Test
public void testNoExtrapolation() throws OrekitException {
// Propagate of the initial at the initial date
final SpacecraftState finalState = propagator.propagate(initDate);
// Initial orbit definition
final Vector3D initialPosition = initialState.getPVCoordinates().getPosition();
final Vector3D initialVelocity = initialState.getPVCoordinates().getVelocity();
// Final orbit definition
final Vector3D finalPosition = finalState.getPVCoordinates().getPosition();
final Vector3D finalVelocity = finalState.getPVCoordinates().getVelocity();
// Check results
Assert.assertEquals(initialPosition.getX(), finalPosition.getX(), 1.0e-10);
Assert.assertEquals(initialPosition.getY(), finalPosition.getY(), 1.0e-10);
Assert.assertEquals(initialPosition.getZ(), finalPosition.getZ(), 1.0e-10);
Assert.assertEquals(initialVelocity.getX(), finalVelocity.getX(), 1.0e-10);
Assert.assertEquals(initialVelocity.getY(), finalVelocity.getY(), 1.0e-10);
Assert.assertEquals(initialVelocity.getZ(), finalVelocity.getZ(), 1.0e-10);
}
/** {@inheritDoc} */
public FieldVector3D<DerivativeStructure> accelerationDerivatives(
final SpacecraftState s, final String paramName) throws OrekitException {
complainIfNotSupported(paramName);
final AbsoluteDate date = s.getDate();
final Frame frame = s.getFrame();
final Vector3D position = s.getPVCoordinates().getPosition();
final Vector3D sunSatVector =
position.subtract(sun.getPVCoordinates(date, frame).getPosition());
final double r2 = sunSatVector.getNormSq();
// compute flux
final double rawP = kRef * getLightningRatio(position, frame, date) / r2;
final Vector3D flux = new Vector3D(rawP / FastMath.sqrt(r2), sunSatVector);
return spacecraft.radiationPressureAcceleration(
date, frame, position, s.getAttitude().getRotation(), s.getMass(), flux, paramName);
}
/** {@inheritDoc} */
public void addContribution(final SpacecraftState s, final TimeDerivativesEquations adder)
throws OrekitException {
// compute bodies separation vectors and squared norm
final Vector3D centralToBody = body.getPVCoordinates(s.getDate(), s.getFrame()).getPosition();
final double r2Central = centralToBody.getNormSq();
final Vector3D satToBody = centralToBody.subtract(s.getPVCoordinates().getPosition());
final double r2Sat = satToBody.getNormSq();
// compute relative acceleration
final Vector3D gamma =
new Vector3D(
gm / (r2Sat * FastMath.sqrt(r2Sat)),
satToBody,
-gm / (r2Central * FastMath.sqrt(r2Central)),
centralToBody);
// add contribution to the ODE second member
adder.addXYZAcceleration(gamma.getX(), gamma.getY(), gamma.getZ());
}
/** {@inheritDoc} */
public FieldVector3D<DerivativeStructure> accelerationDerivatives(
final SpacecraftState s, final String paramName) throws OrekitException {
complainIfNotSupported(paramName);
// compute bodies separation vectors and squared norm
final Vector3D centralToBody = body.getPVCoordinates(s.getDate(), s.getFrame()).getPosition();
final double r2Central = centralToBody.getNormSq();
final Vector3D satToBody = centralToBody.subtract(s.getPVCoordinates().getPosition());
final double r2Sat = satToBody.getNormSq();
final DerivativeStructure gmds = new DerivativeStructure(1, 1, 0, gm);
// compute relative acceleration
return new FieldVector3D<DerivativeStructure>(
gmds.divide(r2Sat * FastMath.sqrt(r2Sat)),
satToBody,
gmds.divide(-r2Central * FastMath.sqrt(r2Central)),
centralToBody);
}
/** {@inheritDoc} */
public void addContribution(final SpacecraftState s, final TimeDerivativesEquations adder)
throws OrekitException {
final AbsoluteDate date = s.getDate();
final Frame frame = s.getFrame();
final Vector3D position = s.getPVCoordinates().getPosition();
final Vector3D sunSatVector =
position.subtract(sun.getPVCoordinates(date, frame).getPosition());
final double r2 = sunSatVector.getNormSq();
// compute flux
final double rawP = kRef * getLightningRatio(position, frame, date) / r2;
final Vector3D flux = new Vector3D(rawP / FastMath.sqrt(r2), sunSatVector);
final Vector3D acceleration =
spacecraft.radiationPressureAcceleration(
date, frame, position, s.getAttitude().getRotation(), s.getMass(), flux);
// provide the perturbing acceleration to the derivatives adder
adder.addAcceleration(acceleration, s.getFrame());
}
/**
* The G-function is the difference between the Sat-Sun-Sat-Earth angle and the sum of the
* Earth's and Sun's apparent radius.
*
* @param s the current state information : date, kinematics, attitude
* @return value of the g function
* @exception OrekitException if sun or spacecraft position cannot be computed
*/
public double g(final SpacecraftState s) throws OrekitException {
final double[] angle =
getEclipseAngles(s.getPVCoordinates().getPosition(), s.getFrame(), s.getDate());
return angle[0] - angle[1] - angle[2];
}
@Test
public void testPropagationTypesElliptical() throws OrekitException, ParseException, IOException {
ForceModel gravityField =
new HolmesFeatherstoneAttractionModel(
FramesFactory.getITRF(IERSConventions.IERS_2010, true),
GravityFieldFactory.getNormalizedProvider(5, 5));
propagator.addForceModel(gravityField);
// Propagation of the initial at t + dt
final PVCoordinates pv = initialState.getPVCoordinates();
final double dP = 0.001;
final double dV =
initialState.getMu() * dP / (pv.getPosition().getNormSq() * pv.getVelocity().getNorm());
final PVCoordinates pvcM =
propagateInType(initialState, dP, OrbitType.CARTESIAN, PositionAngle.MEAN);
final PVCoordinates pviM =
propagateInType(initialState, dP, OrbitType.CIRCULAR, PositionAngle.MEAN);
final PVCoordinates pveM =
propagateInType(initialState, dP, OrbitType.EQUINOCTIAL, PositionAngle.MEAN);
final PVCoordinates pvkM =
propagateInType(initialState, dP, OrbitType.KEPLERIAN, PositionAngle.MEAN);
final PVCoordinates pvcE =
propagateInType(initialState, dP, OrbitType.CARTESIAN, PositionAngle.ECCENTRIC);
final PVCoordinates pviE =
propagateInType(initialState, dP, OrbitType.CIRCULAR, PositionAngle.ECCENTRIC);
final PVCoordinates pveE =
propagateInType(initialState, dP, OrbitType.EQUINOCTIAL, PositionAngle.ECCENTRIC);
final PVCoordinates pvkE =
propagateInType(initialState, dP, OrbitType.KEPLERIAN, PositionAngle.ECCENTRIC);
final PVCoordinates pvcT =
propagateInType(initialState, dP, OrbitType.CARTESIAN, PositionAngle.TRUE);
final PVCoordinates pviT =
propagateInType(initialState, dP, OrbitType.CIRCULAR, PositionAngle.TRUE);
final PVCoordinates pveT =
propagateInType(initialState, dP, OrbitType.EQUINOCTIAL, PositionAngle.TRUE);
final PVCoordinates pvkT =
propagateInType(initialState, dP, OrbitType.KEPLERIAN, PositionAngle.TRUE);
Assert.assertEquals(0, pvcM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 3.0);
Assert.assertEquals(0, pvcM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 2.0);
Assert.assertEquals(0, pviM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.6);
Assert.assertEquals(0, pviM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.4);
Assert.assertEquals(0, pvkM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.5);
Assert.assertEquals(0, pvkM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.3);
Assert.assertEquals(0, pveM.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.2);
Assert.assertEquals(0, pveM.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.2);
Assert.assertEquals(0, pvcE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 3.0);
Assert.assertEquals(0, pvcE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 2.0);
Assert.assertEquals(0, pviE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.03);
Assert.assertEquals(0, pviE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.04);
Assert.assertEquals(0, pvkE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.4);
Assert.assertEquals(0, pvkE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.3);
Assert.assertEquals(0, pveE.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.2);
Assert.assertEquals(0, pveE.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.07);
Assert.assertEquals(0, pvcT.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 3.0);
Assert.assertEquals(0, pvcT.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 2.0);
Assert.assertEquals(0, pviT.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.3);
Assert.assertEquals(0, pviT.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.2);
Assert.assertEquals(0, pvkT.getPosition().subtract(pveT.getPosition()).getNorm() / dP, 0.4);
Assert.assertEquals(0, pvkT.getVelocity().subtract(pveT.getVelocity()).getNorm() / dV, 0.2);
}