/* Copyright 2002-2015 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,
   * WITHOUT WARRANTIES OR CONDITIONS OF ANY KIND, either express or implied.
   * See the License for the specific language governing permissions and
   * limitations under the License.
   */
  package org.orekit.propagation.numerical;
  
  import java.io.NotSerializableException;
  import java.io.Serializable;
  import java.util.ArrayList;
  import java.util.Arrays;
  import java.util.List;
  
  import org.apache.commons.math3.geometry.euclidean.threed.Vector3D;
  import org.apache.commons.math3.ode.AbstractIntegrator;
  import org.apache.commons.math3.util.FastMath;
  import org.orekit.attitudes.Attitude;
  import org.orekit.attitudes.AttitudeProvider;
  import org.orekit.errors.OrekitException;
  import org.orekit.errors.OrekitMessages;
  import org.orekit.errors.PropagationException;
  import org.orekit.forces.ForceModel;
  import org.orekit.forces.gravity.NewtonianAttraction;
  import org.orekit.frames.Frame;
  import org.orekit.frames.Transform;
  import org.orekit.orbits.Orbit;
  import org.orekit.orbits.OrbitType;
  import org.orekit.orbits.PositionAngle;
  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.PVCoordinates;
  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>
  * The last element is the mass in kilograms.
  * </p>
  * <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>The same propagator can be reused for several orbit extrapolations, by resetting
  * the initial state without modifying the other configuration parameters. However, 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 Rom&eacute;ro
  * @author Luc Maisonobe
  * @author Guylaine Prat
  * @author Fabien Maussion
  * @author V&eacute;ronique Pommier-Maurussane
  */
 public class NumericalPropagator extends AbstractIntegratedPropagator {
 
     /** Central body attraction. */
     private NewtonianAttraction newtonianAttraction;
 
     /** Force models used during the extrapolation of the Orbit, without jacobians. */
     private final List<ForceModel> forceModels;
 
     /** 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 AbstractIntegrator integrator) {
         super(integrator, true);
         forceModels = new ArrayList<ForceModel>();
         initMapper();
         setAttitudeProvider(DEFAULT_LAW);
         setMu(Double.NaN);
         setSlaveMode();
         setOrbitType(OrbitType.EQUINOCTIAL);
         setPositionAngleType(PositionAngle.TRUE);
     }
 
      /** Set the central attraction coefficient μ.
      * @param mu central attraction coefficient (m³/s²)
      * @see #addForceModel(ForceModel)
      */
     public void setMu(final double mu) {
         super.setMu(mu);
         newtonianAttraction = new NewtonianAttraction(mu);
     }
 
     /** Add a force model to the global perturbation model.
      * <p>If this method is not called at all, the integrated orbit will follow
      * a keplerian evolution only.</p>
      * @param model perturbing {@link ForceModel} to add
      * @see #removeForceModels()
      * @see #setMu(double)
      */
     public void addForceModel(final ForceModel model) {
         forceModels.add(model);
     }
 
     /** Remove all perturbing force models from the global perturbation model.
      * <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() {
         forceModels.clear();
     }
 
     /** Get perturbing force models list.
      * @return list of perturbing force models
      * @see #addForceModel(ForceModel)
      * @see #getNewtonianAttractionForceModel()
      */
     public List<ForceModel> getForceModels() {
         return forceModels;
     }
 
     /** Get the Newtonian attraction from the central body force model.
      * @return Newtonian attraction force model
      * @see #setMu(double)
      * @see #getForceModels()
      */
     public NewtonianAttraction getNewtonianAttractionForceModel() {
         return newtonianAttraction;
     }
 
     /** Set propagation orbit type.
      * @param orbitType orbit type to use for propagation
      */
     public void setOrbitType(final OrbitType orbitType) {
         super.setOrbitType(orbitType);
     }
 
     /** Get propagation parameter type.
      * @return orbit type used for propagation
      */
     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
      * @exception PropagationException if initial state cannot be set
      */
     public void setInitialState(final SpacecraftState initialState) throws PropagationException {
         resetInitialState(initialState);
     }
 
     /** {@inheritDoc} */
     public void resetInitialState(final SpacecraftState state) throws PropagationException {
         super.resetInitialState(state);
         if (newtonianAttraction == null) {
             setMu(state.getMu());
         }
         setStartDate(null);
     }
 
     /** {@inheritDoc} */
     public TimeStampedPVCoordinates getPVCoordinates(final AbsoluteDate date, final Frame frame)
         throws OrekitException {
         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 implements Serializable {
 
         /** Serializable UID. */
         private static final long serialVersionUID = 20130621L;
 
         /** 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
          * @param positionAngleType angle type to use for propagation
          * @param attitudeProvider attitude provider
          * @param frame inertial frame
          */
         public 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 double t, final double[] y, final boolean meanOnly)
             throws OrekitException {
             // the parameter meanOnly is ignored for the Numerical Propagator
 
             final double mass = y[6];
             if (mass <= 0.0) {
                 throw new PropagationException(OrekitMessages.SPACECRAFT_MASS_BECOMES_NEGATIVE, mass);
             }
 
             final AbsoluteDate date = mapDoubleToDate(t);
             final Orbit orbit       = getOrbitType().mapArrayToOrbit(y, 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) {
             getOrbitType().mapOrbitToArray(state.getOrbit(), getPositionAngleType(), y);
             y[6] = state.getMass();
         }
 
         /** Replace the instance with a data transfer object for serialization.
          * @return data transfer object that will be serialized
          * @exception NotSerializableException if the state mapper cannot be serialized (typically for DSST propagator)
          */
         private Object writeReplace() throws NotSerializableException {
             return new DataTransferObject(getReferenceDate(), getMu(), getOrbitType(),
                                           getPositionAngleType(), getAttitudeProvider(), getFrame());
         }
 
         /** Internal class used only for serialization. */
         private static class DataTransferObject implements Serializable {
 
             /** Serializable UID. */
             private static final long serialVersionUID = 20130621L;
 
             /** Reference date. */
             private final AbsoluteDate referenceDate;
 
             /** Central attraction coefficient (m³/s²). */
             private final double mu;
 
             /** Orbit type to use for mapping. */
             private final OrbitType orbitType;
 
             /** Angle type to use for propagation. */
             private final PositionAngle positionAngleType;
 
             /** Attitude provider. */
             private final AttitudeProvider attitudeProvider;
 
             /** Inertial frame. */
             private final Frame frame;
 
             /** Simple constructor.
              * @param referenceDate reference date
              * @param mu central attraction coefficient (m³/s²)
              * @param orbitType orbit type to use for mapping
              * @param positionAngleType angle type to use for propagation
              * @param attitudeProvider attitude provider
              * @param frame inertial frame
              */
             public DataTransferObject(final AbsoluteDate referenceDate, final double mu,
                                       final OrbitType orbitType, final PositionAngle positionAngleType,
                                       final AttitudeProvider attitudeProvider, final Frame frame) {
                 this.referenceDate     = referenceDate;
                 this.mu                = mu;
                 this.orbitType         = orbitType;
                 this.positionAngleType = positionAngleType;
                 this.attitudeProvider  = attitudeProvider;
                 this.frame             = frame;
             }
 
             /** Replace the deserialized data transfer object with a {@link OsculatingMapper}.
              * @return replacement {@link OsculatingMapper}
              */
             private Object readResolve() {
                 return new OsculatingMapper(referenceDate, mu, orbitType, positionAngleType, attitudeProvider, frame);
             }
         }
 
     }
 
     /** {@inheritDoc} */
     protected MainStateEquations getMainStateEquations(final AbstractIntegrator integrator) {
         return new Main(integrator);
     }
 
     /** Internal class for osculating parameters integration. */
     private class Main implements MainStateEquations, TimeDerivativesEquations {
 
         /** Derivatives array. */
         private final double[] yDot;
 
         /** Current orbit. */
         private Orbit orbit;
 
         /** Jacobian of the orbital parameters with respect to the cartesian parameters. */
         private double[][] jacobian;
 
         /** Simple constructor.
          * @param integrator numerical integrator to use for propagation.
          */
         public Main(final AbstractIntegrator integrator) {
 
             this.yDot     = new double[7];
             this.jacobian = new double[6][6];
 
             for (final ForceModel forceModel : forceModels) {
                 final EventDetector[] modelDetectors = forceModel.getEventsDetectors();
                 if (modelDetectors != null) {
                     for (final EventDetector detector : modelDetectors) {
                         setUpEventDetector(integrator, detector);
                     }
                 }
             }
 
         }
 
         /** {@inheritDoc} */
         public double[] computeDerivatives(final SpacecraftState state) throws OrekitException {
 
             orbit = state.getOrbit();
             Arrays.fill(yDot, 0.0);
             orbit.getJacobianWrtCartesian(getPositionAngleType(), jacobian);
 
             // compute the contributions of all perturbing forces
             for (final ForceModel forceModel : forceModels) {
                 forceModel.addContribution(state, this);
             }
 
             // finalize derivatives by adding the Kepler contribution
             newtonianAttraction.addContribution(state, this);
 
             return yDot.clone();
 
         }
 
         /** {@inheritDoc} */
         public void addKeplerContribution(final double mu) {
             orbit.addKeplerContribution(getPositionAngleType(), mu, yDot);
         }
 
         /** {@inheritDoc} */
         public void addXYZAcceleration(final double x, final double y, final double z) {
             for (int i = 0; i < 6; ++i) {
                 final double[] jRow = jacobian[i];
                 yDot[i] += jRow[3] * x + jRow[4] * y + jRow[5] * z;
             }
         }
 
         /** {@inheritDoc} */
         public void addAcceleration(final Vector3D gamma, final Frame frame)
             throws OrekitException {
             final Transform t = frame.getTransformTo(orbit.getFrame(), orbit.getDate());
             final Vector3D gammInRefFrame = t.transformVector(gamma);
             addXYZAcceleration(gammInRefFrame.getX(), gammInRefFrame.getY(), gammInRefFrame.getZ());
         }
 
         /** {@inheritDoc} */
         public void addMassDerivative(final double q) {
             if (q > 0) {
                 throw OrekitException.createIllegalArgumentException(OrekitMessages.POSITIVE_FLOW_RATE, q);
             }
             yDot[6] += q;
         }
 
     }
 
     /** Estimate tolerance vectors for integrators.
      * <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>
      * 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>
      * <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
      * @exception PropagationException if Jacobian is singular
      */
     public static double[][] tolerances(final double dP, final Orbit orbit, final OrbitType type)
         throws PropagationException {
 
         // 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 PropagationException(OrekitMessages.SINGULAR_JACOBIAN_FOR_ORBIT_TYPE, type);
                 }
             }
 
         }
 
         Arrays.fill(relTol, dP / FastMath.sqrt(r2));
 
         return new double[][] {
             absTol, relTol
         };
 
     }
 
 }