/** {@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());
}
/** {@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);
}
private SpacecraftState shiftState(
SpacecraftState state,
OrbitType orbitType,
PositionAngle angleType,
double delta,
int column) {
double[] array = stateToArray(state, orbitType, angleType, true);
array[column] += delta;
return arrayToState(
array,
orbitType,
angleType,
state.getFrame(),
state.getDate(),
state.getMu(),
state.getAttitude());
}
/** {@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);
}
/**
* Compute osculating state from mean state.
*
* <p>Compute and add the short periodic variation to the mean {@link SpacecraftState}.
*
* @param meanState initial mean state
* @return osculating state
* @throws OrekitException if the computation of the short-periodic variation fails
*/
private Orbit computeOsculatingOrbit(final SpacecraftState meanState) throws OrekitException {
resetShortPeriodicsCoefficients();
computeShortPeriodicsCoefficients(meanState);
final double[] mean = new double[6];
OrbitType.EQUINOCTIAL.mapOrbitToArray(meanState.getOrbit(), PositionAngle.MEAN, mean);
final double[] y = mean.clone();
for (final DSSTForceModel forceModel : this.forceModels) {
final double[] shortPeriodic =
forceModel.getShortPeriodicVariations(meanState.getDate(), mean);
for (int i = 0; i < shortPeriodic.length; i++) {
y[i] += shortPeriodic[i];
}
}
return OrbitType.EQUINOCTIAL.mapArrayToOrbit(
y, PositionAngle.MEAN, meanState.getDate(), meanState.getMu(), meanState.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];
}
