/**
* 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 testResetStateEvent() throws OrekitException {
final AbsoluteDate resetDate = initDate.shiftedBy(1000);
CheckingHandler<DateDetector> checking =
new CheckingHandler<DateDetector>(Action.RESET_STATE) {
public SpacecraftState resetState(DateDetector detector, SpacecraftState oldState) {
return new SpacecraftState(
oldState.getOrbit(), oldState.getAttitude(), oldState.getMass() - 200.0);
}
};
propagator.addEventDetector(new DateDetector(resetDate).withHandler(checking));
checking.assertEvent(false);
final SpacecraftState finalState = propagator.propagate(initDate.shiftedBy(3200));
checking.assertEvent(true);
Assert.assertEquals(initialState.getMass() - 200, finalState.getMass(), 1.0e-10);
}
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;
}
/** {@inheritDoc} */
public SpacecraftState apply(final SpacecraftState state1) {
if (state1.getDate().compareTo(referenceDate) <= 0 && !applyBefore) {
// the orbit change has not occurred yet, don't change anything
return state1;
}
return new SpacecraftState(
updateOrbit(state1.getOrbit()), state1.getAttitude(), state1.getMass());
}
/**
* Compute mean state from osculating state.
*
* <p>Compute in a DSST sense the mean state corresponding to the input osculating state.
*
* <p>The computing is done through a fixed-point iteration process.
*
* @param osculating initial osculating state
* @return mean state
* @throws OrekitException if the underlying computation of short periodic variation fails
*/
private Orbit computeMeanOrbit(final SpacecraftState osculating) throws OrekitException {
// rough initialization of the mean parameters
EquinoctialOrbit meanOrbit = new EquinoctialOrbit(osculating.getOrbit());
// threshold for each parameter
final double epsilon = 1.0e-13;
final double thresholdA = epsilon * (1 + FastMath.abs(meanOrbit.getA()));
final double thresholdE = epsilon * (1 + meanOrbit.getE());
final double thresholdAngles = epsilon * FastMath.PI;
int i = 0;
while (i++ < 200) {
final SpacecraftState meanState =
new SpacecraftState(meanOrbit, osculating.getAttitude(), osculating.getMass());
// recompute the osculating parameters from the current mean parameters
final EquinoctialOrbit rebuilt = (EquinoctialOrbit) computeOsculatingOrbit(meanState);
// adapted parameters residuals
final double deltaA = osculating.getA() - rebuilt.getA();
final double deltaEx = osculating.getEquinoctialEx() - rebuilt.getEquinoctialEx();
final double deltaEy = osculating.getEquinoctialEy() - rebuilt.getEquinoctialEy();
final double deltaHx = osculating.getHx() - rebuilt.getHx();
final double deltaHy = osculating.getHy() - rebuilt.getHy();
final double deltaLm = MathUtils.normalizeAngle(osculating.getLM() - rebuilt.getLM(), 0.0);
// check convergence
if ((FastMath.abs(deltaA) < thresholdA)
&& (FastMath.abs(deltaEx) < thresholdE)
&& (FastMath.abs(deltaEy) < thresholdE)
&& (FastMath.abs(deltaLm) < thresholdAngles)) {
return meanOrbit;
}
// update mean parameters
meanOrbit =
new EquinoctialOrbit(
meanOrbit.getA() + deltaA,
meanOrbit.getEquinoctialEx() + deltaEx,
meanOrbit.getEquinoctialEy() + deltaEy,
meanOrbit.getHx() + deltaHx,
meanOrbit.getHy() + deltaHy,
meanOrbit.getLM() + deltaLm,
PositionAngle.MEAN,
meanOrbit.getFrame(),
meanOrbit.getDate(),
meanOrbit.getMu());
}
throw new PropagationException(OrekitMessages.UNABLE_TO_COMPUTE_DSST_MEAN_PARAMETERS, i);
}
/** {@inheritDoc} */
@Override
public void mapStateToArray(final SpacecraftState state, final double[] y)
throws OrekitException {
final Orbit meanOrbit;
if (!initialIsOsculating) {
// the state is considered to be already a mean state
meanOrbit = state.getOrbit();
} else {
// the state is considered to be an osculating state
meanOrbit = computeMeanState(state, forceModels).getOrbit();
}
OrbitType.EQUINOCTIAL.mapOrbitToArray(meanOrbit, PositionAngle.MEAN, y);
y[6] = state.getMass();
}
/** {@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 {
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());
}