/**
* 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());
}
/** {@inheritDoc} */
@Override
public double[] computeDerivatives(final SpacecraftState state) throws OrekitException {
// compute common auxiliary elements
final AuxiliaryElements aux = new AuxiliaryElements(state.getOrbit(), I);
// initialize all perturbing forces
for (final DSSTForceModel force : mapper.getForceModels()) {
force.initializeStep(aux);
}
Arrays.fill(yDot, 0.0);
// compute the contributions of all perturbing forces
for (final DSSTForceModel forceModel : mapper.getForceModels()) {
final double[] daidt = forceModel.getMeanElementRate(state);
for (int i = 0; i < daidt.length; i++) {
yDot[i] += daidt[i];
}
}
// finalize derivatives by adding the Kepler contribution
final EquinoctialOrbit orbit =
(EquinoctialOrbit) OrbitType.EQUINOCTIAL.convertType(state.getOrbit());
orbit.addKeplerContribution(PositionAngle.MEAN, getMu(), yDot);
return yDot.clone();
}
@Test
public void testEventDetectionBug() throws OrekitException, IOException, ParseException {
TimeScale utc = TimeScalesFactory.getUTC();
AbsoluteDate initialDate = new AbsoluteDate(2005, 1, 1, 0, 0, 0.0, utc);
double duration = 100000.;
AbsoluteDate endDate = new AbsoluteDate(initialDate, duration);
// Initialization of the frame EME2000
Frame EME2000 = FramesFactory.getEME2000();
// Initial orbit
double a = 35786000. + 6378137.0;
double e = 0.70;
double rApogee = a * (1 + e);
double vApogee = FastMath.sqrt(mu * (1 - e) / (a * (1 + e)));
Orbit geo =
new CartesianOrbit(
new PVCoordinates(new Vector3D(rApogee, 0., 0.), new Vector3D(0., vApogee, 0.)),
EME2000,
initialDate,
mu);
duration = geo.getKeplerianPeriod();
endDate = new AbsoluteDate(initialDate, duration);
// Numerical Integration
final double minStep = 0.001;
final double maxStep = 1000;
final double initStep = 60;
final double[] absTolerance = {0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001};
final 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(minStep, maxStep, absTolerance, relTolerance);
integrator.setInitialStepSize(initStep);
// Numerical propagator based on the integrator
propagator = new NumericalPropagator(integrator);
double mass = 1000.;
SpacecraftState initialState = new SpacecraftState(geo, mass);
propagator.setInitialState(initialState);
propagator.setOrbitType(OrbitType.CARTESIAN);
// Set the events Detectors
ApsideDetector event1 = new ApsideDetector(geo);
propagator.addEventDetector(event1);
// Set the propagation mode
propagator.setSlaveMode();
// Propagate
SpacecraftState finalState = propagator.propagate(endDate);
// we should stop long before endDate
Assert.assertTrue(endDate.durationFrom(finalState.getDate()) > 40000.0);
}
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());
}
@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 testStopEvent() throws OrekitException {
final AbsoluteDate stopDate = initDate.shiftedBy(1000);
CheckingHandler<DateDetector> checking = new CheckingHandler<DateDetector>(Action.STOP);
propagator.addEventDetector(new DateDetector(stopDate).withHandler(checking));
Assert.assertEquals(1, propagator.getEventsDetectors().size());
checking.assertEvent(false);
final SpacecraftState finalState = propagator.propagate(initDate.shiftedBy(3200));
checking.assertEvent(true);
Assert.assertEquals(0, finalState.getDate().durationFrom(stopDate), 1.0e-10);
propagator.clearEventsDetectors();
Assert.assertEquals(0, propagator.getEventsDetectors().size());
}
@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
}
}
/** {@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();
}
/**
* Method called just before integration.
*
* <p>The default implementation does nothing, it may be specialized in subclasses.
*
* @param initialState initial state
* @param tEnd target date at which state should be propagated
* @exception OrekitException if hook cannot be run
*/
@Override
protected void beforeIntegration(final SpacecraftState initialState, final AbsoluteDate tEnd)
throws OrekitException {
// compute common auxiliary elements
final AuxiliaryElements aux = new AuxiliaryElements(initialState.getOrbit(), I);
// check if only mean elements must be used
final boolean meanOnly = isMeanOrbit();
// initialize all perturbing forces
for (final DSSTForceModel force : mapper.getForceModels()) {
force.initialize(aux, meanOnly);
}
// if required, insert the special short periodics step handler
if (!meanOnly) {
final InterpolationGrid grid =
new VariableStepInterpolationGrid(INTERPOLATION_POINTS_PER_STEP);
final ShortPeriodicsHandler spHandler = new ShortPeriodicsHandler(grid);
final Collection<StepHandler> stepHandlers = new ArrayList<StepHandler>();
stepHandlers.add(spHandler);
final AbstractIntegrator integrator = getIntegrator();
final Collection<StepHandler> existing = integrator.getStepHandlers();
stepHandlers.addAll(existing);
integrator.clearStepHandlers();
// add back the existing handlers after the short periodics one
for (final StepHandler sp : stepHandlers) {
integrator.addStepHandler(sp);
}
}
}
/**
* 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);
}
/** {@inheritDoc} */
@Override
public double[] getMeanElementRate(final SpacecraftState spacecraftState) throws OrekitException {
// Compute potential derivatives
final double[] dU = computeUDerivatives(spacecraftState.getDate());
final double dUda = dU[0];
final double dUdh = dU[1];
final double dUdk = dU[2];
final double dUdl = dU[3];
final double dUdAl = dU[4];
final double dUdBe = dU[5];
final double dUdGa = dU[6];
// Compute the cross derivative operator :
final double UAlphaGamma = alpha * dUdGa - gamma * dUdAl;
final double UAlphaBeta = alpha * dUdBe - beta * dUdAl;
final double UBetaGamma = beta * dUdGa - gamma * dUdBe;
final double Uhk = h * dUdk - k * dUdh;
final double pUagmIqUbgoAB = (p * UAlphaGamma - I * q * UBetaGamma) * ooAB;
final double UhkmUabmdUdl = Uhk - UAlphaBeta - dUdl;
final double da = ax2oA * dUdl;
final double dh = BoA * dUdk + k * pUagmIqUbgoAB - h * BoABpo * dUdl;
final double dk = -(BoA * dUdh + h * pUagmIqUbgoAB + k * BoABpo * dUdl);
final double dp = Co2AB * (p * UhkmUabmdUdl - UBetaGamma);
final double dq = Co2AB * (q * UhkmUabmdUdl - I * UAlphaGamma);
final double dM = -ax2oA * dUda + BoABpo * (h * dUdh + k * dUdk) + pUagmIqUbgoAB;
return new double[] {da, dk, dh, dq, dp, dM};
}
@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);
}
/**
* 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);
}
/**
* Launch the computation of short periodics coefficients.
*
* @param state input state
* @throws OrekitException if the computation of a short periodic coefficient fails
* @see {@link DSSTForceModel#computeShortPeriodicsCoefficients(AuxiliaryElements)}
*/
private void computeShortPeriodicsCoefficients(final SpacecraftState state)
throws OrekitException {
final AuxiliaryElements aux = new AuxiliaryElements(state.getOrbit(), I);
for (final DSSTForceModel forceModel : forceModels) {
forceModel.initializeStep(aux);
forceModel.computeShortPeriodicsCoefficients(state);
}
}
/** {@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());
}
/** check propagation succeeds when two events are within the tolerance of each other. */
@Test
public void testCloseEventDates() throws PropagationException {
// setup
DateDetector d1 =
new DateDetector(10, 1, initDate.shiftedBy(15))
.withHandler(new ContinueOnEvent<DateDetector>());
DateDetector d2 =
new DateDetector(10, 1, initDate.shiftedBy(15.5))
.withHandler(new ContinueOnEvent<DateDetector>());
propagator.addEventDetector(d1);
propagator.addEventDetector(d2);
// action
AbsoluteDate end = initDate.shiftedBy(30);
SpacecraftState actual = propagator.propagate(end);
// verify
Assert.assertEquals(actual.getDate().durationFrom(end), 0.0, 0.0);
}
/* (non-Javadoc)
* @see org.orekit.propagation.sampling.OrekitFixedStepHandler#handleStep(org.orekit.propagation.SpacecraftState, boolean)
*/
public void handleStep(SpacecraftState currentState, boolean isLast) throws PropagationException {
/** Create position vector. */
D3Vector position =
new D3Vector(
"",
"Position",
"Position",
"The orbital position of the satellite at the given time.",
currentState.getOrbit().getPVCoordinates().getPosition().getX(),
currentState.getOrbit().getPVCoordinates().getPosition().getY(),
currentState.getOrbit().getPVCoordinates().getPosition().getZ());
/** Create velocity vector. */
D3Vector velocity =
new D3Vector(
"",
"Velocity",
"Velocity",
"The orbital velocity of the satellite at the given time",
currentState.getOrbit().getPVCoordinates().getVelocity().getX(),
currentState.getOrbit().getPVCoordinates().getVelocity().getY(),
currentState.getOrbit().getPVCoordinates().getVelocity().getZ());
try {
/** Create orbital state. */
OrbitalState state =
new OrbitalState(
predictor.name,
orbitalStateName,
orbitalStateDescription,
currentState.getDate().toDate(TimeScalesFactory.getUTC()).getTime(),
datasetidentifier,
satellite,
position,
velocity);
/** Send the orbital state on the response stream. */
predictor.addResult(state);
} catch (OrekitException e) {
e.printStackTrace();
}
}
/**
* 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());
}
/** {@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 {
// 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());
}
private PVCoordinates propagateInType(
SpacecraftState state, double dP, OrbitType type, PositionAngle angle)
throws PropagationException {
final double dt = 3200;
final double minStep = 0.001;
final double maxStep = 1000;
double[][] tol = NumericalPropagator.tolerances(dP, state.getOrbit(), type);
AdaptiveStepsizeIntegrator integrator =
new DormandPrince853Integrator(minStep, maxStep, tol[0], tol[1]);
NumericalPropagator newPropagator = new NumericalPropagator(integrator);
newPropagator.setOrbitType(type);
newPropagator.setPositionAngleType(angle);
newPropagator.setInitialState(state);
for (ForceModel force : propagator.getForceModels()) {
newPropagator.addForceModel(force);
}
return newPropagator.propagate(state.getDate().shiftedBy(dt)).getPVCoordinates();
}
/** {@inheritDoc} */
public void init(final SpacecraftState s0, final AbsoluteDate t) {
// delegate to raw detector
rawDetector.init(s0, t);
// initialize events triggering logic
forward = t.compareTo(s0.getDate()) >= 0;
extremeT = forward ? AbsoluteDate.PAST_INFINITY : AbsoluteDate.FUTURE_INFINITY;
extremeG = Double.NaN;
Arrays.fill(transformers, Transformer.UNINITIALIZED);
Arrays.fill(updates, extremeT);
}
@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);
}
/**
* 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 gravityField gravity field to use
* @exception OrekitException if gravity field does not contain J2 coefficient
*/
public J2DifferentialEffect(
final SpacecraftState original,
final AdapterPropagator.DifferentialEffect directEffect,
final boolean applyBefore,
final UnnormalizedSphericalHarmonicsProvider gravityField)
throws OrekitException {
this(
original,
directEffect,
applyBefore,
gravityField.getAe(),
gravityField.getMu(),
-gravityField.onDate(original.getDate()).getUnnormalizedCnm(2, 0));
}
@Test
public void testResetAdditionalStateEvent() throws OrekitException {
propagator.addAdditionalEquations(
new AdditionalEquations() {
public String getName() {
return "linear";
}
public double[] computeDerivatives(SpacecraftState s, double[] pDot) {
pDot[0] = 1.0;
return null;
}
});
propagator.setInitialState(propagator.getInitialState().addAdditionalState("linear", 1.5));
CheckingHandler<AdditionalStateLinearDetector> checking =
new CheckingHandler<AdditionalStateLinearDetector>(Action.RESET_STATE) {
public SpacecraftState resetState(
AdditionalStateLinearDetector detector, SpacecraftState oldState)
throws OrekitException {
return oldState.addAdditionalState(
"linear", oldState.getAdditionalState("linear")[0] * 2);
}
};
propagator.addEventDetector(
new AdditionalStateLinearDetector(10.0, 1.0e-8).withHandler(checking));
final double dt = 3200;
checking.assertEvent(false);
final SpacecraftState finalState = propagator.propagate(initDate.shiftedBy(dt));
checking.assertEvent(true);
Assert.assertEquals(dt + 4.5, finalState.getAdditionalState("linear")[0], 1.0e-8);
Assert.assertEquals(dt, finalState.getDate().durationFrom(initDate), 1.0e-8);
}
/** {@inheritDoc} */
public void computeShortPeriodicsCoefficients(final SpacecraftState state)
throws OrekitException {
// Initialise the Hansen coefficients
for (int s = -maxDegree; s <= maxDegree; s++) {
// coefficients with j == 0 are always needed
this.hansenObjects[s + maxDegree][0].computeInitValues(e2, chi, chi2);
if (!mDailiesOnly) {
// initialize other objects only if required
for (int j = 1; j <= jMax; j++) {
this.hansenObjects[s + maxDegree][j].computeInitValues(e2, chi, chi2);
}
}
}
// Compute coefficients
tesseralSPCoefs.computeCoefficients(state.getDate());
}
/** {@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());
}
@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);
}
