@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);
}
@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);
}
/** create {@link #cache} and {@link #data} with neighborsSize = 3 */
@Before
public void setUp() {
data =
Arrays.asList(
date,
date.shiftedBy(1),
date.shiftedBy(2),
date.shiftedBy(3),
date.shiftedBy(4),
date.shiftedBy(5));
cache = new ImmutableTimeStampedCache<AbsoluteDate>(3, data);
}
@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 testEphemerisGenerationIssue14() throws OrekitException, IOException {
// Propagation of the initial at t + dt
final double dt = 3200;
propagator.getInitialState();
propagator.setOrbitType(OrbitType.CARTESIAN);
propagator.setEphemerisMode();
propagator.propagate(initDate.shiftedBy(dt));
final BoundedPropagator ephemeris1 = propagator.getGeneratedEphemeris();
Assert.assertEquals(initDate, ephemeris1.getMinDate());
Assert.assertEquals(initDate.shiftedBy(dt), ephemeris1.getMaxDate());
propagator.getPVCoordinates(initDate.shiftedBy(2 * dt), FramesFactory.getEME2000());
propagator.getPVCoordinates(initDate.shiftedBy(-2 * dt), FramesFactory.getEME2000());
// the new propagations should not have changed ephemeris1
Assert.assertEquals(initDate, ephemeris1.getMinDate());
Assert.assertEquals(initDate.shiftedBy(dt), ephemeris1.getMaxDate());
final BoundedPropagator ephemeris2 = propagator.getGeneratedEphemeris();
Assert.assertEquals(initDate.shiftedBy(-2 * dt), ephemeris2.getMinDate());
Assert.assertEquals(initDate.shiftedBy(2 * dt), ephemeris2.getMaxDate());
// generating ephemeris2 should not have changed ephemeris1
Assert.assertEquals(initDate, ephemeris1.getMinDate());
Assert.assertEquals(initDate.shiftedBy(dt), ephemeris1.getMaxDate());
}
private double[] derivativesErrors(
PVCoordinatesProvider provider, AbsoluteDate date, Frame frame, OneAxisEllipsoid model)
throws OrekitException {
List<TimeStampedPVCoordinates> pvList = new ArrayList<TimeStampedPVCoordinates>();
List<TimeStampedPVCoordinates> groundPVList = new ArrayList<TimeStampedPVCoordinates>();
for (double dt = -0.25; dt <= 0.25; dt += 0.125) {
TimeStampedPVCoordinates shiftedPV = provider.getPVCoordinates(date.shiftedBy(dt), frame);
Vector3D p = model.projectToGround(shiftedPV.getPosition(), shiftedPV.getDate(), frame);
pvList.add(shiftedPV);
groundPVList.add(
new TimeStampedPVCoordinates(shiftedPV.getDate(), p, Vector3D.ZERO, Vector3D.ZERO));
}
TimeStampedPVCoordinates computed =
model.projectToGround(
TimeStampedPVCoordinates.interpolate(date, CartesianDerivativesFilter.USE_P, pvList),
frame);
TimeStampedPVCoordinates reference =
TimeStampedPVCoordinates.interpolate(date, CartesianDerivativesFilter.USE_P, groundPVList);
TimeStampedPVCoordinates pv0 = provider.getPVCoordinates(date, frame);
Vector3D p0 = pv0.getPosition();
Vector3D v0 = pv0.getVelocity();
Vector3D a0 = pv0.getAcceleration();
return new double[] {
Vector3D.distance(computed.getPosition(), reference.getPosition()) / p0.getNorm(),
Vector3D.distance(computed.getVelocity(), reference.getVelocity()) / v0.getNorm(),
Vector3D.distance(computed.getAcceleration(), reference.getAcceleration()) / a0.getNorm(),
};
}
/**
* check that the cache is immutable by changing the data passed in the constructor and returned
* from methods.
*
* @throws TimeStampedCacheException
*/
@Test
public void testImmutable() throws TimeStampedCacheException {
// setup
List<AbsoluteDate> actuals;
List<AbsoluteDate> expecteds = new ArrayList<AbsoluteDate>(data);
AbsoluteDate different = date.shiftedBy(-50);
// actions + verify
// check constructor
data.set(0, different);
Assert.assertArrayEquals(cache.getAll().toArray(), expecteds.toArray());
// check getAll
actuals = cache.getAll();
try {
actuals.set(0, different);
} catch (UnsupportedOperationException e) {
// exception ok
}
Assert.assertArrayEquals(cache.getAll().toArray(), expecteds.toArray());
// check getNeighbors
actuals = cache.getNeighbors(date);
try {
actuals.set(0, different);
} catch (UnsupportedOperationException e) {
// exception ok
}
Assert.assertArrayEquals(cache.getAll().toArray(), expecteds.toArray());
}
@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);
}
@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
}
}
@Test
public void testFrance() throws OrekitException {
final BodyShape earth =
new OneAxisEllipsoid(
Constants.WGS84_EARTH_EQUATORIAL_RADIUS,
Constants.WGS84_EARTH_FLATTENING,
FramesFactory.getITRF(IERSConventions.IERS_2010, true));
GeographicZoneDetector d =
new GeographicZoneDetector(20.0, 1.e-3, earth, buildFrance(), FastMath.toRadians(0.5))
.withHandler(new ContinueOnEvent<GeographicZoneDetector>());
Assert.assertEquals(20.0, d.getMaxCheckInterval(), 1.0e-15);
Assert.assertEquals(1.0e-3, d.getThreshold(), 1.0e-15);
Assert.assertEquals(0.5, FastMath.toDegrees(d.getMargin()), 1.0e-15);
Assert.assertEquals(AbstractDetector.DEFAULT_MAX_ITER, d.getMaxIterationCount());
final TimeScale utc = TimeScalesFactory.getUTC();
final Vector3D position = new Vector3D(-6142438.668, 3492467.56, -25767.257);
final Vector3D velocity = new Vector3D(505.848, 942.781, 7435.922);
final AbsoluteDate date = new AbsoluteDate(2003, 9, 16, utc);
final Orbit orbit =
new EquinoctialOrbit(
new PVCoordinates(position, velocity),
FramesFactory.getEME2000(),
date,
Constants.EIGEN5C_EARTH_MU);
Propagator propagator =
new EcksteinHechlerPropagator(
orbit,
Constants.EIGEN5C_EARTH_EQUATORIAL_RADIUS,
Constants.EIGEN5C_EARTH_MU,
Constants.EIGEN5C_EARTH_C20,
Constants.EIGEN5C_EARTH_C30,
Constants.EIGEN5C_EARTH_C40,
Constants.EIGEN5C_EARTH_C50,
Constants.EIGEN5C_EARTH_C60);
EventsLogger logger = new EventsLogger();
propagator.addEventDetector(logger.monitorDetector(d));
propagator.propagate(date.shiftedBy(10 * Constants.JULIAN_DAY));
Assert.assertEquals(26, logger.getLoggedEvents().size());
}
@Test
public void testEphemerisDatesBackward() throws OrekitException {
// setup
TimeScale tai = TimeScalesFactory.getTAI();
AbsoluteDate initialDate = new AbsoluteDate("2015-07-05", tai);
AbsoluteDate startDate = new AbsoluteDate("2015-07-03", tai).shiftedBy(-0.1);
AbsoluteDate endDate = new AbsoluteDate("2015-07-04", tai);
Frame eci = FramesFactory.getGCRF();
KeplerianOrbit orbit =
new KeplerianOrbit(
600e3 + Constants.WGS84_EARTH_EQUATORIAL_RADIUS,
0,
0,
0,
0,
0,
PositionAngle.TRUE,
eci,
initialDate,
mu);
double[][] tol = NumericalPropagator.tolerances(1, orbit, OrbitType.CARTESIAN);
Propagator prop =
new NumericalPropagator(new DormandPrince853Integrator(0.1, 500, tol[0], tol[1]));
prop.resetInitialState(new SpacecraftState(new CartesianOrbit(orbit)));
// action
prop.setEphemerisMode();
prop.propagate(endDate, startDate);
BoundedPropagator ephemeris = prop.getGeneratedEphemeris();
// verify
TimeStampedPVCoordinates actualPV = ephemeris.getPVCoordinates(startDate, eci);
TimeStampedPVCoordinates expectedPV = orbit.getPVCoordinates(startDate, eci);
MatcherAssert.assertThat(
actualPV.getPosition(), OrekitMatchers.vectorCloseTo(expectedPV.getPosition(), 1.0));
MatcherAssert.assertThat(
actualPV.getVelocity(), OrekitMatchers.vectorCloseTo(expectedPV.getVelocity(), 1.0));
MatcherAssert.assertThat(
ephemeris.getMinDate().durationFrom(startDate), OrekitMatchers.closeTo(0, 0));
MatcherAssert.assertThat(
ephemeris.getMaxDate().durationFrom(endDate), OrekitMatchers.closeTo(0, 0));
// test date
AbsoluteDate date = endDate.shiftedBy(-0.11);
Assert.assertEquals(ephemeris.propagate(date).getDate().durationFrom(date), 0, 0);
}
@Test
public void testKepler() throws OrekitException {
// Propagation of the initial at t + dt
final double dt = 3200;
final SpacecraftState finalState =
propagator.propagate(initDate.shiftedBy(-60), initDate.shiftedBy(dt));
// Check results
final double n =
FastMath.sqrt(initialState.getMu() / initialState.getA()) / initialState.getA();
Assert.assertEquals(initialState.getA(), finalState.getA(), 1.0e-10);
Assert.assertEquals(initialState.getEquinoctialEx(), finalState.getEquinoctialEx(), 1.0e-10);
Assert.assertEquals(initialState.getEquinoctialEy(), finalState.getEquinoctialEy(), 1.0e-10);
Assert.assertEquals(initialState.getHx(), finalState.getHx(), 1.0e-10);
Assert.assertEquals(initialState.getHy(), finalState.getHy(), 1.0e-10);
Assert.assertEquals(initialState.getLM() + n * dt, finalState.getLM(), 2.0e-9);
}
/**
* Approximate an already fitted model to polynomial only terms.
*
* <p>This method is mainly used in order to combine the large amplitude long periods with the
* secular part as a new approximate polynomial model over some time range. This should be used
* rather than simply extracting the polynomial coefficients from {@link #getFittedParameters()}
* when some periodic terms amplitudes are large (for example Sun resonance effects on local solar
* time in sun synchronous orbits). In theses cases, the pure polynomial secular part in the
* coefficients may be far from the mean model.
*
* @param combinedDegree desired degree for the combined polynomial
* @param combinedReference desired reference date for the combined polynomial
* @param meanDegree degree of polynomial secular part to consider
* @param meanHarmonics number of harmonics terms to consider
* @param start start date of the approximation time range
* @param end end date of the approximation time range
* @param step sampling step
* @return coefficients of the approximate polynomial (in increasing degree order), using the user
* provided reference date
*/
public double[] approximateAsPolynomialOnly(
final int combinedDegree,
final AbsoluteDate combinedReference,
final int meanDegree,
final int meanHarmonics,
final AbsoluteDate start,
final AbsoluteDate end,
final double step) {
final List<WeightedObservedPoint> points = new ArrayList<WeightedObservedPoint>();
for (AbsoluteDate date = start; date.compareTo(end) < 0; date = date.shiftedBy(step)) {
points.add(
new WeightedObservedPoint(
1.0,
date.durationFrom(combinedReference),
meanValue(date, meanDegree, meanHarmonics)));
}
return PolynomialCurveFitter.create(combinedDegree).fit(points);
}
@Test
public void testContinueEvent() throws OrekitException {
final AbsoluteDate resetDate = initDate.shiftedBy(1000);
CheckingHandler<DateDetector> checking = new CheckingHandler<DateDetector>(Action.CONTINUE);
propagator.addEventDetector(new DateDetector(resetDate).withHandler(checking));
final double dt = 3200;
checking.assertEvent(false);
final SpacecraftState finalState = propagator.propagate(initDate.shiftedBy(dt));
checking.assertEvent(true);
final double n =
FastMath.sqrt(initialState.getMu() / initialState.getA()) / initialState.getA();
Assert.assertEquals(initialState.getA(), finalState.getA(), 1.0e-10);
Assert.assertEquals(initialState.getEquinoctialEx(), finalState.getEquinoctialEx(), 1.0e-10);
Assert.assertEquals(initialState.getEquinoctialEy(), finalState.getEquinoctialEy(), 1.0e-10);
Assert.assertEquals(initialState.getHx(), finalState.getHx(), 1.0e-10);
Assert.assertEquals(initialState.getHy(), finalState.getHy(), 1.0e-10);
Assert.assertEquals(initialState.getLM() + n * dt, finalState.getLM(), 6.0e-10);
}
/** 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);
}
/** {@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);
}
/**
* Compute the differential effect of J2 on an orbit.
*
* @param orbit1 original orbit at t₁, without differential J2
* @return orbit at t₁, always taking the effect into account
*/
private Orbit updateOrbit(final Orbit orbit1) {
// convert current orbital state to equinoctial elements
final EquinoctialOrbit original = (EquinoctialOrbit) OrbitType.EQUINOCTIAL.convertType(orbit1);
// compute differential effect
final AbsoluteDate date = original.getDate();
final double dt = date.durationFrom(referenceDate);
final double dPaRaan = (dPaDot + dRaanDot) * dt;
final double cPaRaan = FastMath.cos(dPaRaan);
final double sPaRaan = FastMath.sin(dPaRaan);
final double dRaan = dRaanDot * dt;
final double cRaan = FastMath.cos(dRaan);
final double sRaan = FastMath.sin(dRaan);
final double ex = original.getEquinoctialEx() * cPaRaan - original.getEquinoctialEy() * sPaRaan;
final double ey = original.getEquinoctialEx() * sPaRaan + original.getEquinoctialEy() * cPaRaan;
final double hx = original.getHx() * cRaan - original.getHy() * sRaan;
final double hy = original.getHx() * sRaan + original.getHy() * cRaan;
final double lambda = original.getLv() + dPaRaan;
// build updated orbit
final EquinoctialOrbit updated =
new EquinoctialOrbit(
original.getA(),
ex,
ey,
hx,
hy,
lambda,
PositionAngle.TRUE,
original.getFrame(),
date,
original.getMu());
// convert to required type
return orbit1.getType().convertType(updated);
}
public void handleStep(OrekitStepInterpolator interpolator, boolean isLast)
throws PropagationException {
try {
if (pickUpDate == null) {
// we want to pick up the Jacobians at the end of last step
if (isLast) {
interpolator.setInterpolatedDate(interpolator.getCurrentDate());
}
} else {
// we want to pick up some intermediate Jacobians
double dt0 = pickUpDate.durationFrom(interpolator.getPreviousDate());
double dt1 = pickUpDate.durationFrom(interpolator.getCurrentDate());
if (dt0 * dt1 > 0) {
// the current step does not cover the pickup date
return;
} else {
interpolator.setInterpolatedDate(pickUpDate);
}
}
Assert.assertEquals(1, interpolator.getInterpolatedState().getAdditionalStates().size());
Assert.assertTrue(
interpolator
.getInterpolatedState()
.getAdditionalStates()
.containsKey(mapper.getName()));
SpacecraftState state = interpolator.getInterpolatedState();
mapper.getStateJacobian(state, dYdY0);
mapper.getParametersJacobian(state, dYdP);
} catch (PropagationException pe) {
throw pe;
} catch (OrekitException oe) {
throw new PropagationException(oe);
}
}
/**
* Compute the transform at some date.
*
* @param date date at which the transform is desired
* @return computed transform at specified date
*/
public Transform getTransform(final AbsoluteDate date) {
// compute parameters evolution since reference epoch
final double dt = date.durationFrom(epoch);
final Vector3D dR = new Vector3D(1, rotationVector, dt, rotationRate);
// build translation part
final Transform translationTransform = new Transform(date, cartesian.shiftedBy(dt));
// build rotation part
final double angle = dR.getNorm();
final Transform rotationTransform =
new Transform(
date,
(angle < Precision.SAFE_MIN) ? Rotation.IDENTITY : new Rotation(dR, angle),
rotationRate);
// combine both parts
return new Transform(date, translationTransform, rotationTransform);
}
@Test(expected = OrekitException.class)
public void testException() throws OrekitException {
propagator.setMasterMode(
new OrekitStepHandler() {
private int countDown = 3;
private AbsoluteDate previousCall = null;
public void init(SpacecraftState s0, AbsoluteDate t) {}
public void handleStep(OrekitStepInterpolator interpolator, boolean isLast)
throws PropagationException {
if (previousCall != null) {
Assert.assertTrue(interpolator.getInterpolatedDate().compareTo(previousCall) < 0);
}
if (--countDown == 0) {
throw new PropagationException(LocalizedFormats.SIMPLE_MESSAGE, "dummy error");
}
}
});
propagator.propagate(initDate.shiftedBy(-3600));
}
@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);
}
/**
* Add a fitting point.
*
* @param date date of the point
* @param osculatingValue osculating value
*/
public void addPoint(final AbsoluteDate date, final double osculatingValue) {
observedPoints.add(
new WeightedObservedPoint(1.0, date.durationFrom(reference), osculatingValue));
}
/**
* Get mean second derivative, truncated to first components.
*
* @param date current date
* @param degree degree of polynomial secular part
* @param harmonics number of harmonics terms to consider
* @return mean second derivative at current date
*/
public double meanSecondDerivative(
final AbsoluteDate date, final int degree, final int harmonics) {
return truncatedSecondDerivative(degree, harmonics, date.durationFrom(reference), fitted);
}
/**
* Get fitted osculating second derivative.
*
* @param date current date
* @return osculating second derivative at current date
*/
public double osculatingSecondDerivative(final AbsoluteDate date) {
return truncatedSecondDerivative(
secularDegree, pulsations.length, date.durationFrom(reference), fitted);
}
@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);
}
/** {@inheritDoc} */
public TimeStampedPVCoordinates getPVCoordinates(
final AbsoluteDate otherDate, final Frame otherFrame) throws OrekitException {
return shiftedBy(otherDate.durationFrom(getDate())).getPVCoordinates(otherFrame);
}
/** {@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);
}
}
}
@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 testEphemerisAdditionalState() throws OrekitException, IOException {
// Propagation of the initial at t + dt
final double dt = -3200;
final double rate = 2.0;
propagator.addAdditionalStateProvider(
new AdditionalStateProvider() {
public String getName() {
return "squaredA";
}
public double[] getAdditionalState(SpacecraftState state) {
return new double[] {state.getA() * state.getA()};
}
});
propagator.addAdditionalEquations(
new AdditionalEquations() {
public String getName() {
return "extra";
}
public double[] computeDerivatives(SpacecraftState s, double[] pDot) {
pDot[0] = rate;
return null;
}
});
propagator.setInitialState(propagator.getInitialState().addAdditionalState("extra", 1.5));
propagator.setOrbitType(OrbitType.CARTESIAN);
propagator.setEphemerisMode();
propagator.propagate(initDate.shiftedBy(dt));
final BoundedPropagator ephemeris1 = propagator.getGeneratedEphemeris();
Assert.assertEquals(initDate.shiftedBy(dt), ephemeris1.getMinDate());
Assert.assertEquals(initDate, ephemeris1.getMaxDate());
try {
ephemeris1.propagate(ephemeris1.getMinDate().shiftedBy(-10.0));
Assert.fail("an exception should have been thrown");
} catch (PropagationException pe) {
Assert.assertEquals(OrekitMessages.OUT_OF_RANGE_EPHEMERIDES_DATE, pe.getSpecifier());
}
try {
ephemeris1.propagate(ephemeris1.getMaxDate().shiftedBy(+10.0));
Assert.fail("an exception should have been thrown");
} catch (PropagationException pe) {
Assert.assertEquals(OrekitMessages.OUT_OF_RANGE_EPHEMERIDES_DATE, pe.getSpecifier());
}
double shift = -60;
SpacecraftState s = ephemeris1.propagate(initDate.shiftedBy(shift));
Assert.assertEquals(2, s.getAdditionalStates().size());
Assert.assertTrue(s.hasAdditionalState("squaredA"));
Assert.assertTrue(s.hasAdditionalState("extra"));
Assert.assertEquals(s.getA() * s.getA(), s.getAdditionalState("squaredA")[0], 1.0e-10);
Assert.assertEquals(1.5 + shift * rate, s.getAdditionalState("extra")[0], 1.0e-10);
try {
ephemeris1.resetInitialState(s);
Assert.fail("an exception should have been thrown");
} catch (OrekitException oe) {
Assert.assertEquals(OrekitMessages.NON_RESETABLE_STATE, oe.getSpecifier());
}
}
