NumericalPropagator.java
/* Copyright 2002-2019 CS Systèmes d'Information
* Licensed to CS Systèmes d'Information (CS) under one or more
* contributor license agreements. See the NOTICE file distributed with
* this work for additional information regarding copyright ownership.
* CS licenses this file to You under the Apache License, Version 2.0
* (the "License"); you may not use this file except in compliance with
* the License. You may obtain a copy of the License at
*
* http://www.apache.org/licenses/LICENSE-2.0
*
* Unless required by applicable law or agreed to in writing, software
* distributed under the License is distributed on an "AS IS" BASIS,
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* See the License for the specific language governing permissions and
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package org.orekit.propagation.numerical;
import java.util.ArrayList;
import java.util.Arrays;
import java.util.Collections;
import java.util.List;
import org.hipparchus.geometry.euclidean.threed.Vector3D;
import org.hipparchus.ode.ODEIntegrator;
import org.hipparchus.util.FastMath;
import org.orekit.attitudes.Attitude;
import org.orekit.attitudes.AttitudeProvider;
import org.orekit.errors.OrekitException;
import org.orekit.errors.OrekitIllegalArgumentException;
import org.orekit.errors.OrekitInternalError;
import org.orekit.errors.OrekitMessages;
import org.orekit.forces.ForceModel;
import org.orekit.forces.gravity.NewtonianAttraction;
import org.orekit.forces.inertia.InertialForces;
import org.orekit.frames.Frame;
import org.orekit.orbits.Orbit;
import org.orekit.orbits.OrbitType;
import org.orekit.orbits.PositionAngle;
import org.orekit.propagation.PropagationType;
import org.orekit.propagation.SpacecraftState;
import org.orekit.propagation.events.EventDetector;
import org.orekit.propagation.integration.AbstractIntegratedPropagator;
import org.orekit.propagation.integration.StateMapper;
import org.orekit.time.AbsoluteDate;
import org.orekit.utils.AbsolutePVCoordinates;
import org.orekit.utils.PVCoordinates;
import org.orekit.utils.ParameterDriver;
import org.orekit.utils.ParameterObserver;
import org.orekit.utils.TimeStampedPVCoordinates;
/** This class propagates {@link org.orekit.orbits.Orbit orbits} using
* numerical integration.
* <p>Numerical propagation is much more accurate than analytical propagation
* like for example {@link org.orekit.propagation.analytical.KeplerianPropagator
* Keplerian} or {@link org.orekit.propagation.analytical.EcksteinHechlerPropagator
* Eckstein-Hechler}, but requires a few more steps to set up to be used properly.
* Whereas analytical propagators are configured only thanks to their various
* constructors and can be used immediately after construction, numerical propagators
* configuration involve setting several parameters between construction time
* and propagation time.</p>
* <p>The configuration parameters that can be set are:</p>
* <ul>
* <li>the initial spacecraft state ({@link #setInitialState(SpacecraftState)})</li>
* <li>the central attraction coefficient ({@link #setMu(double)})</li>
* <li>the various force models ({@link #addForceModel(ForceModel)},
* {@link #removeForceModels()})</li>
* <li>the {@link OrbitType type} of orbital parameters to be used for propagation
* ({@link #setOrbitType(OrbitType)}),
* <li>the {@link PositionAngle type} of position angle to be used in orbital parameters
* to be used for propagation where it is relevant ({@link
* #setPositionAngleType(PositionAngle)}),
* <li>whether {@link org.orekit.propagation.integration.AdditionalEquations additional equations}
* (for example {@link PartialDerivativesEquations Jacobians}) should be propagated along with orbital state
* ({@link #addAdditionalEquations(org.orekit.propagation.integration.AdditionalEquations)}),
* <li>the discrete events that should be triggered during propagation
* ({@link #addEventDetector(EventDetector)},
* {@link #clearEventsDetectors()})</li>
* <li>the binding logic with the rest of the application ({@link #setSlaveMode()},
* {@link #setMasterMode(double, org.orekit.propagation.sampling.OrekitFixedStepHandler)},
* {@link #setMasterMode(org.orekit.propagation.sampling.OrekitStepHandler)},
* {@link #setEphemerisMode()}, {@link #getGeneratedEphemeris()})</li>
* </ul>
* <p>From these configuration parameters, only the initial state is mandatory. The default
* propagation settings are in {@link OrbitType#EQUINOCTIAL equinoctial} parameters with
* {@link PositionAngle#TRUE true} longitude argument. If the central attraction coefficient
* is not explicitly specified, the one used to define the initial orbit will be used.
* However, specifying only the initial state and perhaps the central attraction coefficient
* would mean the propagator would use only Keplerian forces. In this case, the simpler {@link
* org.orekit.propagation.analytical.KeplerianPropagator KeplerianPropagator} class would
* perhaps be more effective.</p>
* <p>The underlying numerical integrator set up in the constructor may also have its own
* configuration parameters. Typical configuration parameters for adaptive stepsize integrators
* are the min, max and perhaps start step size as well as the absolute and/or relative errors
* thresholds.</p>
* <p>The state that is seen by the integrator is a simple seven elements double array.
* The six first elements are either:
* <ul>
* <li>the {@link org.orekit.orbits.EquinoctialOrbit equinoctial orbit parameters} (a, e<sub>x</sub>,
* e<sub>y</sub>, h<sub>x</sub>, h<sub>y</sub>, λ<sub>M</sub> or λ<sub>E</sub>
* or λ<sub>v</sub>) in meters and radians,</li>
* <li>the {@link org.orekit.orbits.KeplerianOrbit Keplerian orbit parameters} (a, e, i, ω, Ω,
* M or E or v) in meters and radians,</li>
* <li>the {@link org.orekit.orbits.CircularOrbit circular orbit parameters} (a, e<sub>x</sub>, e<sub>y</sub>, i,
* Ω, α<sub>M</sub> or α<sub>E</sub> or α<sub>v</sub>) in meters
* and radians,</li>
* <li>the {@link org.orekit.orbits.CartesianOrbit Cartesian orbit parameters} (x, y, z, v<sub>x</sub>,
* v<sub>y</sub>, v<sub>z</sub>) in meters and meters per seconds.
* </ul>
* <p> The last element is the mass in kilograms.
*
* <p>The following code snippet shows a typical setting for Low Earth Orbit propagation in
* equinoctial parameters and true longitude argument:</p>
* <pre>
* final double dP = 0.001;
* final double minStep = 0.001;
* final double maxStep = 500;
* final double initStep = 60;
* final double[][] tolerance = NumericalPropagator.tolerances(dP, orbit, OrbitType.EQUINOCTIAL);
* AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tolerance[0], tolerance[1]);
* integrator.setInitialStepSize(initStep);
* propagator = new NumericalPropagator(integrator);
* </pre>
* <p>By default, at the end of the propagation, the propagator resets the initial state to the final state,
* thus allowing a new propagation to be started from there without recomputing the part already performed.
* This behaviour can be chenged by calling {@link #setResetAtEnd(boolean)}.
* </p>
* <p>Beware the same instance cannot be used simultaneously by different threads, the class is <em>not</em>
* thread-safe.</p>
*
* @see SpacecraftState
* @see ForceModel
* @see org.orekit.propagation.sampling.OrekitStepHandler
* @see org.orekit.propagation.sampling.OrekitFixedStepHandler
* @see org.orekit.propagation.integration.IntegratedEphemeris
* @see TimeDerivativesEquations
*
* @author Mathieu Roméro
* @author Luc Maisonobe
* @author Guylaine Prat
* @author Fabien Maussion
* @author Véronique Pommier-Maurussane
*/
public class NumericalPropagator extends AbstractIntegratedPropagator {
/** Force models used during the extrapolation of the orbit. */
private final List<ForceModel> forceModels;
/** boolean to ignore or not the creation of a NewtonianAttraction. */
private boolean ignoreCentralAttraction = false;
/** Create a new instance of NumericalPropagator, based on orbit definition mu.
* After creation, the instance is empty, i.e. the attitude provider is set to an
* unspecified default law and there are no perturbing forces at all.
* This means that if {@link #addForceModel addForceModel} is not
* called after creation, the integrated orbit will follow a Keplerian
* evolution only. The defaults are {@link OrbitType#EQUINOCTIAL}
* for {@link #setOrbitType(OrbitType) propagation
* orbit type} and {@link PositionAngle#TRUE} for {@link
* #setPositionAngleType(PositionAngle) position angle type}.
* @param integrator numerical integrator to use for propagation.
*/
public NumericalPropagator(final ODEIntegrator integrator) {
super(integrator, PropagationType.MEAN);
forceModels = new ArrayList<ForceModel>();
initMapper();
setAttitudeProvider(DEFAULT_LAW);
setSlaveMode();
setOrbitType(OrbitType.EQUINOCTIAL);
setPositionAngleType(PositionAngle.TRUE);
}
public void setIgnoreCentralAttraction(final boolean ignoreCentralAttraction) {
this.ignoreCentralAttraction = ignoreCentralAttraction;
}
/** Set the central attraction coefficient μ.
* <p>
* Setting the central attraction coefficient is
* equivalent to {@link #addForceModel(ForceModel) add}
* a {@link NewtonianAttraction} force model.
* </p>
* @param mu central attraction coefficient (m³/s²)
* @see #addForceModel(ForceModel)
* @see #getAllForceModels()
*/
public void setMu(final double mu) {
if (ignoreCentralAttraction) {
superSetMu(mu);
} else {
addForceModel(new NewtonianAttraction(mu));
}
}
/** Set the central attraction coefficient μ only in upper class.
* @param mu central attraction coefficient (m³/s²)
*/
private void superSetMu(final double mu) {
super.setMu(mu);
}
/** Check if Newtonian attraction force model is available.
* <p>
* Newtonian attraction is always the last force model in the list.
* </p>
* @return true if Newtonian attraction force model is available
*/
private boolean hasNewtonianAttraction() {
final int last = forceModels.size() - 1;
return last >= 0 && forceModels.get(last) instanceof NewtonianAttraction;
}
/** Add a force model.
* <p>If this method is not called at all, the integrated orbit will follow
* a Keplerian evolution only.</p>
* @param model {@link ForceModel} to add (it can be either a perturbing force
* model or an instance of {@link NewtonianAttraction})
* @see #removeForceModels()
* @see #setMu(double)
*/
public void addForceModel(final ForceModel model) {
if (model instanceof NewtonianAttraction) {
// we want to add the central attraction force model
try {
// ensure we are notified of any mu change
model.getParametersDrivers()[0].addObserver(new ParameterObserver() {
/** {@inheritDoc} */
@Override
public void valueChanged(final double previousValue, final ParameterDriver driver) {
superSetMu(driver.getValue());
}
});
} catch (OrekitException oe) {
// this should never happen
throw new OrekitInternalError(oe);
}
if (hasNewtonianAttraction()) {
// there is already a central attraction model, replace it
forceModels.set(forceModels.size() - 1, model);
} else {
// there are no central attraction model yet, add it at the end of the list
forceModels.add(model);
}
} else {
// we want to add a perturbing force model
if (hasNewtonianAttraction()) {
// insert the new force model before Newtonian attraction,
// which should always be the last one in the list
forceModels.add(forceModels.size() - 1, model);
} else {
// we only have perturbing force models up to now, just append at the end of the list
forceModels.add(model);
}
}
}
/** Remove all force models (except central attraction).
* <p>Once all perturbing forces have been removed (and as long as no new force
* model is added), the integrated orbit will follow a Keplerian evolution
* only.</p>
* @see #addForceModel(ForceModel)
*/
public void removeForceModels() {
final int last = forceModels.size() - 1;
if (hasNewtonianAttraction()) {
// preserve the Newtonian attraction model at the end
final ForceModel newton = forceModels.get(last);
forceModels.clear();
forceModels.add(newton);
} else {
forceModels.clear();
}
}
/** Get all the force models, perturbing forces and Newtonian attraction included.
* @return list of perturbing force models, with Newtonian attraction being the
* last one
* @see #addForceModel(ForceModel)
* @see #setMu(double)
*/
public List<ForceModel> getAllForceModels() {
return Collections.unmodifiableList(forceModels);
}
/** Set propagation orbit type.
* @param orbitType orbit type to use for propagation, null for
* propagating using {@link org.orekit.utils.AbsolutePVCoordinates} rather than {@link Orbit}
*/
public void setOrbitType(final OrbitType orbitType) {
super.setOrbitType(orbitType);
}
/** Get propagation parameter type.
* @return orbit type used for propagation, null for
* propagating using {@link org.orekit.utils.AbsolutePVCoordinates} rather than {@link Orbit}
*/
public OrbitType getOrbitType() {
return super.getOrbitType();
}
/** Set position angle type.
* <p>
* The position parameter type is meaningful only if {@link
* #getOrbitType() propagation orbit type}
* support it. As an example, it is not meaningful for propagation
* in {@link OrbitType#CARTESIAN Cartesian} parameters.
* </p>
* @param positionAngleType angle type to use for propagation
*/
public void setPositionAngleType(final PositionAngle positionAngleType) {
super.setPositionAngleType(positionAngleType);
}
/** Get propagation parameter type.
* @return angle type to use for propagation
*/
public PositionAngle getPositionAngleType() {
return super.getPositionAngleType();
}
/** Set the initial state.
* @param initialState initial state
*/
public void setInitialState(final SpacecraftState initialState) {
resetInitialState(initialState);
}
/** {@inheritDoc} */
public void resetInitialState(final SpacecraftState state) {
super.resetInitialState(state);
if (!hasNewtonianAttraction()) {
// use the state to define central attraction
setMu(state.getMu());
}
setStartDate(state.getDate());
}
/** {@inheritDoc} */
public TimeStampedPVCoordinates getPVCoordinates(final AbsoluteDate date, final Frame frame) {
return propagate(date).getPVCoordinates(frame);
}
/** {@inheritDoc} */
protected StateMapper createMapper(final AbsoluteDate referenceDate, final double mu,
final OrbitType orbitType, final PositionAngle positionAngleType,
final AttitudeProvider attitudeProvider, final Frame frame) {
return new OsculatingMapper(referenceDate, mu, orbitType, positionAngleType, attitudeProvider, frame);
}
/** Internal mapper using directly osculating parameters. */
private static class OsculatingMapper extends StateMapper {
/** Simple constructor.
* <p>
* The position parameter type is meaningful only if {@link
* #getOrbitType() propagation orbit type}
* support it. As an example, it is not meaningful for propagation
* in {@link OrbitType#CARTESIAN Cartesian} parameters.
* </p>
* @param referenceDate reference date
* @param mu central attraction coefficient (m³/s²)
* @param orbitType orbit type to use for mapping (can be null for {@link AbsolutePVCoordinates})
* @param positionAngleType angle type to use for propagation
* @param attitudeProvider attitude provider
* @param frame inertial frame
*/
OsculatingMapper(final AbsoluteDate referenceDate, final double mu,
final OrbitType orbitType, final PositionAngle positionAngleType,
final AttitudeProvider attitudeProvider, final Frame frame) {
super(referenceDate, mu, orbitType, positionAngleType, attitudeProvider, frame);
}
/** {@inheritDoc} */
public SpacecraftState mapArrayToState(final AbsoluteDate date, final double[] y, final double[] yDot,
final PropagationType type) {
// the parameter type is ignored for the Numerical Propagator
final double mass = y[6];
if (mass <= 0.0) {
throw new OrekitException(OrekitMessages.SPACECRAFT_MASS_BECOMES_NEGATIVE, mass);
}
if (getOrbitType() == null) {
// propagation uses absolute position-velocity-acceleration
final Vector3D p = new Vector3D(y[0], y[1], y[2]);
final Vector3D v = new Vector3D(y[3], y[4], y[5]);
final Vector3D a;
final AbsolutePVCoordinates absPva;
if (yDot == null) {
absPva = new AbsolutePVCoordinates(getFrame(), new TimeStampedPVCoordinates(date, p, v));
} else {
a = new Vector3D(yDot[3], yDot[4], yDot[5]);
absPva = new AbsolutePVCoordinates(getFrame(), new TimeStampedPVCoordinates(date, p, v, a));
}
final Attitude attitude = getAttitudeProvider().getAttitude(absPva, date, getFrame());
return new SpacecraftState(absPva, attitude, mass);
} else {
// propagation uses regular orbits
final Orbit orbit = getOrbitType().mapArrayToOrbit(y, yDot, getPositionAngleType(), date, getMu(), getFrame());
final Attitude attitude = getAttitudeProvider().getAttitude(orbit, date, getFrame());
return new SpacecraftState(orbit, attitude, mass);
}
}
/** {@inheritDoc} */
public void mapStateToArray(final SpacecraftState state, final double[] y, final double[] yDot) {
if (getOrbitType() == null) {
// propagation uses absolute position-velocity-acceleration
final Vector3D p = state.getAbsPVA().getPosition();
final Vector3D v = state.getAbsPVA().getVelocity();
y[0] = p.getX();
y[1] = p.getY();
y[2] = p.getZ();
y[3] = v.getX();
y[4] = v.getY();
y[5] = v.getZ();
y[6] = state.getMass();
}
else {
getOrbitType().mapOrbitToArray(state.getOrbit(), getPositionAngleType(), y, yDot);
y[6] = state.getMass();
}
}
}
/** {@inheritDoc} */
protected MainStateEquations getMainStateEquations(final ODEIntegrator integrator) {
return new Main(integrator);
}
/** Internal class for osculating parameters integration. */
private class Main implements MainStateEquations, TimeDerivativesEquations {
/** Derivatives array. */
private final double[] yDot;
/** Current state. */
private SpacecraftState currentState;
/** Jacobian of the orbital parameters with respect to the Cartesian parameters. */
private double[][] jacobian;
/** Simple constructor.
* @param integrator numerical integrator to use for propagation.
*/
Main(final ODEIntegrator integrator) {
this.yDot = new double[7];
this.jacobian = new double[6][6];
for (final ForceModel forceModel : forceModels) {
forceModel.getEventsDetectors().forEach(detector -> setUpEventDetector(integrator, detector));
}
if (getOrbitType() == null) {
// propagation uses absolute position-velocity-acceleration
// we can set Jacobian once and for all
for (int i = 0; i < jacobian.length; ++i) {
Arrays.fill(jacobian[i], 0.0);
jacobian[i][i] = 1.0;
}
}
}
/** {@inheritDoc} */
@Override
public void init(final SpacecraftState initialState, final AbsoluteDate target) {
for (final ForceModel forceModel : forceModels) {
forceModel.init(initialState, target);
}
}
/** {@inheritDoc} */
@Override
public double[] computeDerivatives(final SpacecraftState state) {
currentState = state;
Arrays.fill(yDot, 0.0);
if (getOrbitType() != null) {
// propagation uses regular orbits
currentState.getOrbit().getJacobianWrtCartesian(getPositionAngleType(), jacobian);
}
// compute the contributions of all perturbing forces,
// using the Kepler contribution at the end since
// NewtonianAttraction is always the last instance in the list
for (final ForceModel forceModel : forceModels) {
forceModel.addContribution(state, this);
}
if (getOrbitType() == null) {
// position derivative is velocity, and was not added above in the force models
// (it is added when orbit type is non-null because NewtonianAttraction considers it)
final Vector3D velocity = currentState.getPVCoordinates().getVelocity();
yDot[0] += velocity.getX();
yDot[1] += velocity.getY();
yDot[2] += velocity.getZ();
}
return yDot.clone();
}
/** {@inheritDoc} */
@Override
public void addKeplerContribution(final double mu) {
if (getOrbitType() == null) {
// if mu is neither 0 nor NaN, we want to include Newtonian acceleration
if (mu > 0) {
// velocity derivative is Newtonian acceleration
final Vector3D position = currentState.getPVCoordinates().getPosition();
final double r2 = position.getNormSq();
final double coeff = -mu / (r2 * FastMath.sqrt(r2));
yDot[3] += coeff * position.getX();
yDot[4] += coeff * position.getY();
yDot[5] += coeff * position.getZ();
}
} else {
// propagation uses regular orbits
currentState.getOrbit().addKeplerContribution(getPositionAngleType(), mu, yDot);
}
}
/** {@inheritDoc} */
public void addNonKeplerianAcceleration(final Vector3D gamma) {
for (int i = 0; i < 6; ++i) {
final double[] jRow = jacobian[i];
yDot[i] += jRow[3] * gamma.getX() + jRow[4] * gamma.getY() + jRow[5] * gamma.getZ();
}
}
/** {@inheritDoc} */
@Override
public void addMassDerivative(final double q) {
if (q > 0) {
throw new OrekitIllegalArgumentException(OrekitMessages.POSITIVE_FLOW_RATE, q);
}
yDot[6] += q;
}
}
/** Estimate tolerance vectors for integrators when propagating in absolute position-velocity-acceleration.
* @param dP user specified position error
* @param absPva reference absolute position-velocity-acceleration
* @return a two rows array, row 0 being the absolute tolerance error and row 1
* being the relative tolerance error
* @see NumericalPropagator#tolerances(double, Orbit, OrbitType)
*/
public static double[][] tolerances(final double dP, final AbsolutePVCoordinates absPva) {
final double relative = dP / absPva.getPosition().getNorm();
final double dV = relative * absPva.getVelocity().getNorm();
final double[] absTol = new double[7];
final double[] relTol = new double[7];
absTol[0] = dP;
absTol[1] = dP;
absTol[2] = dP;
absTol[3] = dV;
absTol[4] = dV;
absTol[5] = dV;
// we set the mass tolerance arbitrarily to 1.0e-6 kg, as mass evolves linearly
// with trust, this often has no influence at all on propagation
absTol[6] = 1.0e-6;
Arrays.fill(relTol, relative);
return new double[][] {
absTol, relTol
};
}
/** Estimate tolerance vectors for integrators when propagating in orbits.
* <p>
* The errors are estimated from partial derivatives properties of orbits,
* starting from a scalar position error specified by the user.
* Considering the energy conservation equation V = sqrt(mu (2/r - 1/a)),
* we get at constant energy (i.e. on a Keplerian trajectory):
* <pre>
* V² r |dV| = mu |dr|
* </pre>
* <p> So we deduce a scalar velocity error consistent with the position error.
* From here, we apply orbits Jacobians matrices to get consistent errors
* on orbital parameters.
*
* <p>
* The tolerances are only <em>orders of magnitude</em>, and integrator tolerances
* are only local estimates, not global ones. So some care must be taken when using
* these tolerances. Setting 1mm as a position error does NOT mean the tolerances
* will guarantee a 1mm error position after several orbits integration.
* </p>
* @param dP user specified position error
* @param orbit reference orbit
* @param type propagation type for the meaning of the tolerance vectors elements
* (it may be different from {@code orbit.getType()})
* @return a two rows array, row 0 being the absolute tolerance error and row 1
* being the relative tolerance error
*/
public static double[][] tolerances(final double dP, final Orbit orbit, final OrbitType type) {
// estimate the scalar velocity error
final PVCoordinates pv = orbit.getPVCoordinates();
final double r2 = pv.getPosition().getNormSq();
final double v = pv.getVelocity().getNorm();
final double dV = orbit.getMu() * dP / (v * r2);
final double[] absTol = new double[7];
final double[] relTol = new double[7];
// we set the mass tolerance arbitrarily to 1.0e-6 kg, as mass evolves linearly
// with trust, this often has no influence at all on propagation
absTol[6] = 1.0e-6;
if (type == OrbitType.CARTESIAN) {
absTol[0] = dP;
absTol[1] = dP;
absTol[2] = dP;
absTol[3] = dV;
absTol[4] = dV;
absTol[5] = dV;
} else {
// convert the orbit to the desired type
final double[][] jacobian = new double[6][6];
final Orbit converted = type.convertType(orbit);
converted.getJacobianWrtCartesian(PositionAngle.TRUE, jacobian);
for (int i = 0; i < 6; ++i) {
final double[] row = jacobian[i];
absTol[i] = FastMath.abs(row[0]) * dP +
FastMath.abs(row[1]) * dP +
FastMath.abs(row[2]) * dP +
FastMath.abs(row[3]) * dV +
FastMath.abs(row[4]) * dV +
FastMath.abs(row[5]) * dV;
if (Double.isNaN(absTol[i])) {
throw new OrekitException(OrekitMessages.SINGULAR_JACOBIAN_FOR_ORBIT_TYPE, type);
}
}
}
Arrays.fill(relTol, dP / FastMath.sqrt(r2));
return new double[][] {
absTol, relTol
};
}
/** {@inheritDoc} */
@Override
protected void beforeIntegration(final SpacecraftState initialState, final AbsoluteDate tEnd) {
if (!getFrame().isPseudoInertial()) {
// inspect all force models to find InertialForces
for (ForceModel force : forceModels) {
if (force instanceof InertialForces) {
return;
}
}
// throw exception if no inertial forces found
throw new OrekitException(OrekitMessages.INERTIAL_FORCE_MODEL_MISSING, getFrame().getName());
}
}
}