/** {@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 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);
}
/** {@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 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 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 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());
}
/**
* 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));
}
/** {@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 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);
}
/* (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();
}
}
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);
}
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 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());
}
@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);
}
@Test
public void testAdditionalStateEvent() throws OrekitException {
propagator.addAdditionalEquations(
new AdditionalEquations() {
public String getName() {
return "linear";
}
public double[] computeDerivatives(SpacecraftState s, double[] pDot) {
pDot[0] = 1.0;
return new double[7];
}
});
try {
propagator.addAdditionalEquations(
new AdditionalEquations() {
public String getName() {
return "linear";
}
public double[] computeDerivatives(SpacecraftState s, double[] pDot) {
pDot[0] = 1.0;
return new double[7];
}
});
Assert.fail("an exception should have been thrown");
} catch (OrekitException oe) {
Assert.assertEquals(oe.getSpecifier(), OrekitMessages.ADDITIONAL_STATE_NAME_ALREADY_IN_USE);
}
try {
propagator.addAdditionalStateProvider(
new AdditionalStateProvider() {
public String getName() {
return "linear";
}
public double[] getAdditionalState(SpacecraftState state) {
return null;
}
});
Assert.fail("an exception should have been thrown");
} catch (OrekitException oe) {
Assert.assertEquals(oe.getSpecifier(), OrekitMessages.ADDITIONAL_STATE_NAME_ALREADY_IN_USE);
}
propagator.addAdditionalStateProvider(
new AdditionalStateProvider() {
public String getName() {
return "constant";
}
public double[] getAdditionalState(SpacecraftState state) {
return new double[] {1.0};
}
});
Assert.assertTrue(propagator.isAdditionalStateManaged("linear"));
Assert.assertTrue(propagator.isAdditionalStateManaged("constant"));
Assert.assertFalse(propagator.isAdditionalStateManaged("non-managed"));
Assert.assertEquals(2, propagator.getManagedAdditionalStates().length);
propagator.setInitialState(propagator.getInitialState().addAdditionalState("linear", 1.5));
CheckingHandler<AdditionalStateLinearDetector> checking =
new CheckingHandler<AdditionalStateLinearDetector>(Action.STOP);
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(3.0, finalState.getAdditionalState("linear")[0], 1.0e-8);
Assert.assertEquals(1.5, finalState.getDate().durationFrom(initDate), 1.0e-8);
}
@Test
public void testPropagationTypesElliptical() throws OrekitException, ParseException, IOException {
NormalizedSphericalHarmonicsProvider provider = GravityFieldFactory.getNormalizedProvider(5, 5);
ForceModel gravityField =
new HolmesFeatherstoneAttractionModel(
FramesFactory.getITRF(IERSConventions.IERS_2010, true), provider);
Orbit initialOrbit =
new KeplerianOrbit(
8000000.0,
0.01,
0.1,
0.7,
0,
1.2,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
AbsoluteDate.J2000_EPOCH,
provider.getMu());
double dt = 900;
double dP = 0.001;
for (OrbitType orbitType : OrbitType.values()) {
for (PositionAngle angleType : PositionAngle.values()) {
// compute state Jacobian using PartialDerivatives
NumericalPropagator propagator =
setUpPropagator(initialOrbit, dP, orbitType, angleType, gravityField);
PartialDerivativesEquations partials =
new PartialDerivativesEquations("partials", propagator);
final SpacecraftState initialState =
partials.setInitialJacobians(new SpacecraftState(initialOrbit), 6, 0);
propagator.setInitialState(initialState);
final JacobiansMapper mapper = partials.getMapper();
PickUpHandler pickUp = new PickUpHandler(mapper, null);
propagator.setMasterMode(pickUp);
propagator.propagate(initialState.getDate().shiftedBy(dt));
double[][] dYdY0 = pickUp.getdYdY0();
// compute reference state Jacobian using finite differences
double[][] dYdY0Ref = new double[6][6];
AbstractIntegratedPropagator propagator2 =
setUpPropagator(initialOrbit, dP, orbitType, angleType, gravityField);
double[] steps = NumericalPropagator.tolerances(1000000 * dP, initialOrbit, orbitType)[0];
for (int i = 0; i < 6; ++i) {
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -4 * steps[i], i));
SpacecraftState sM4h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -3 * steps[i], i));
SpacecraftState sM3h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -2 * steps[i], i));
SpacecraftState sM2h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, -1 * steps[i], i));
SpacecraftState sM1h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 1 * steps[i], i));
SpacecraftState sP1h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 2 * steps[i], i));
SpacecraftState sP2h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 3 * steps[i], i));
SpacecraftState sP3h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
propagator2.resetInitialState(
shiftState(initialState, orbitType, angleType, 4 * steps[i], i));
SpacecraftState sP4h = propagator2.propagate(initialState.getDate().shiftedBy(dt));
fillJacobianColumn(
dYdY0Ref, i, orbitType, angleType, steps[i], sM4h, sM3h, sM2h, sM1h, sP1h, sP2h, sP3h,
sP4h);
}
for (int i = 0; i < 6; ++i) {
for (int j = 0; j < 6; ++j) {
double error = FastMath.abs((dYdY0[i][j] - dYdY0Ref[i][j]) / dYdY0Ref[i][j]);
Assert.assertEquals(0, error, 6.0e-2);
}
}
}
}
}
@Test
public void testJacobianIssue18() throws OrekitException {
// Body mu
final double mu = 3.9860047e14;
final double isp = 318;
final double mass = 2500;
final double a = 24396159;
final double e = 0.72831215;
final double i = FastMath.toRadians(7);
final double omega = FastMath.toRadians(180);
final double OMEGA = FastMath.toRadians(261);
final double lv = 0;
final double duration = 3653.99;
final double f = 420;
final double delta = FastMath.toRadians(-7.4978);
final double alpha = FastMath.toRadians(351);
final AttitudeProvider law =
new InertialProvider(new Rotation(new Vector3D(alpha, delta), Vector3D.PLUS_I));
final AbsoluteDate initDate =
new AbsoluteDate(
new DateComponents(2004, 01, 01),
new TimeComponents(23, 30, 00.000),
TimeScalesFactory.getUTC());
final Orbit orbit =
new KeplerianOrbit(
a,
e,
i,
omega,
OMEGA,
lv,
PositionAngle.TRUE,
FramesFactory.getEME2000(),
initDate,
mu);
final SpacecraftState initialState =
new SpacecraftState(orbit, law.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass);
final AbsoluteDate fireDate =
new AbsoluteDate(
new DateComponents(2004, 01, 02),
new TimeComponents(04, 15, 34.080),
TimeScalesFactory.getUTC());
final ConstantThrustManeuver maneuver =
new ConstantThrustManeuver(fireDate, duration, f, isp, Vector3D.PLUS_I);
double[] absTolerance = {0.001, 1.0e-9, 1.0e-9, 1.0e-6, 1.0e-6, 1.0e-6, 0.001};
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(0.001, 1000, absTolerance, relTolerance);
integrator.setInitialStepSize(60);
final NumericalPropagator propagator = new NumericalPropagator(integrator);
propagator.setAttitudeProvider(law);
propagator.addForceModel(maneuver);
propagator.setOrbitType(OrbitType.CARTESIAN);
PartialDerivativesEquations PDE = new PartialDerivativesEquations("derivatives", propagator);
PDE.selectParamAndStep("thrust", Double.NaN);
Assert.assertEquals(3, PDE.getAvailableParameters().size());
Assert.assertEquals("central attraction coefficient", PDE.getAvailableParameters().get(0));
Assert.assertEquals("thrust", PDE.getAvailableParameters().get(1));
Assert.assertEquals("flow rate", PDE.getAvailableParameters().get(2));
propagator.setInitialState(PDE.setInitialJacobians(initialState, 7, 1));
final AbsoluteDate finalDate = fireDate.shiftedBy(3800);
final SpacecraftState finalorb = propagator.propagate(finalDate);
Assert.assertEquals(0, finalDate.durationFrom(finalorb.getDate()), 1.0e-11);
}
/**
* 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];
}
/** {@inheritDoc} */
public double g(final SpacecraftState s) throws OrekitException {
final double rawG = rawDetector.g(s);
final boolean isEnabled = enabler.eventIsEnabled(s, rawDetector, rawG);
if (Double.isNaN(extremeG)) {
extremeG = rawG;
}
// search which transformer should be applied to g
if (forward) {
final int last = transformers.length - 1;
if (extremeT.compareTo(s.getDate()) < 0) {
// we are at the forward end of the history
// check if enabled status has changed
final Transformer previous = transformers[last];
final Transformer next = selectTransformer(previous, extremeG, isEnabled);
if (next != previous) {
// there is a status change somewhere between extremeT and t.
// the new transformer is valid for t (this is how we have just computed
// it above), but it is in fact valid on both sides of the change, so
// it was already valid before t and even up to previous time. We store
// the switch at extremeT for safety, to ensure the previous transformer
// is not applied too close of the root
System.arraycopy(updates, 1, updates, 0, last);
System.arraycopy(transformers, 1, transformers, 0, last);
updates[last] = extremeT;
transformers[last] = next;
}
extremeT = s.getDate();
extremeG = rawG;
// apply the transform
return next.transformed(rawG);
} else {
// we are in the middle of the history
// select the transformer
for (int i = last; i > 0; --i) {
if (updates[i].compareTo(s.getDate()) <= 0) {
// apply the transform
return transformers[i].transformed(rawG);
}
}
return transformers[0].transformed(rawG);
}
} else {
if (s.getDate().compareTo(extremeT) < 0) {
// we are at the backward end of the history
// check if a new rough root has been crossed
final Transformer previous = transformers[0];
final Transformer next = selectTransformer(previous, extremeG, isEnabled);
if (next != previous) {
// there is a status change somewhere between extremeT and t.
// the new transformer is valid for t (this is how we have just computed
// it above), but it is in fact valid on both sides of the change, so
// it was already valid before t and even up to previous time. We store
// the switch at extremeT for safety, to ensure the previous transformer
// is not applied too close of the root
System.arraycopy(updates, 0, updates, 1, updates.length - 1);
System.arraycopy(transformers, 0, transformers, 1, transformers.length - 1);
updates[0] = extremeT;
transformers[0] = next;
}
extremeT = s.getDate();
extremeG = rawG;
// apply the transform
return next.transformed(rawG);
} else {
// we are in the middle of the history
// select the transformer
for (int i = 0; i < updates.length - 1; ++i) {
if (s.getDate().compareTo(updates[i]) <= 0) {
// apply the transform
return transformers[i].transformed(rawG);
}
}
return transformers[updates.length - 1].transformed(rawG);
}
}
}
/**
* Reset the initial state.
*
* @param state new initial state
* @throws PropagationException if initial state cannot be reset
*/
@Override
public void resetInitialState(final SpacecraftState state) throws PropagationException {
super.setStartDate(state.getDate());
super.resetInitialState(state);
}