DSSTPropagator.java
/* Copyright 2002-2020 CS Group
* Licensed to CS Group (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
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package org.orekit.propagation.semianalytical.dsst;
import java.util.ArrayList;
import java.util.Arrays;
import java.util.Collection;
import java.util.Collections;
import java.util.HashMap;
import java.util.HashSet;
import java.util.List;
import java.util.Map;
import java.util.Set;
import org.hipparchus.ode.ODEIntegrator;
import org.hipparchus.ode.ODEStateAndDerivative;
import org.hipparchus.ode.sampling.ODEStateInterpolator;
import org.hipparchus.ode.sampling.ODEStepHandler;
import org.hipparchus.util.FastMath;
import org.hipparchus.util.MathUtils;
import org.orekit.annotation.DefaultDataContext;
import org.orekit.attitudes.Attitude;
import org.orekit.attitudes.AttitudeProvider;
import org.orekit.data.DataContext;
import org.orekit.errors.OrekitException;
import org.orekit.errors.OrekitInternalError;
import org.orekit.errors.OrekitMessages;
import org.orekit.frames.Frame;
import org.orekit.orbits.EquinoctialOrbit;
import org.orekit.orbits.Orbit;
import org.orekit.orbits.OrbitType;
import org.orekit.orbits.PositionAngle;
import org.orekit.propagation.PropagationType;
import org.orekit.propagation.Propagator;
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.propagation.numerical.NumericalPropagator;
import org.orekit.propagation.semianalytical.dsst.forces.DSSTForceModel;
import org.orekit.propagation.semianalytical.dsst.forces.DSSTNewtonianAttraction;
import org.orekit.propagation.semianalytical.dsst.forces.ShortPeriodTerms;
import org.orekit.propagation.semianalytical.dsst.utilities.AuxiliaryElements;
import org.orekit.propagation.semianalytical.dsst.utilities.FixedNumberInterpolationGrid;
import org.orekit.propagation.semianalytical.dsst.utilities.InterpolationGrid;
import org.orekit.propagation.semianalytical.dsst.utilities.MaxGapInterpolationGrid;
import org.orekit.time.AbsoluteDate;
import org.orekit.utils.ParameterDriver;
import org.orekit.utils.ParameterObserver;
/**
* This class propagates {@link org.orekit.orbits.Orbit orbits} using the DSST theory.
* <p>
* Whereas analytical propagators are configured only thanks to their various
* constructors and can be used immediately after construction, such a semianalytical
* propagator configuration involves setting several parameters between construction
* time and propagation time, just as numerical propagators.
* </p>
* <p>
* The configuration parameters that can be set are:
* </p>
* <ul>
* <li>the initial spacecraft state ({@link #setInitialState(SpacecraftState)})</li>
* <li>the various force models ({@link #addForceModel(DSSTForceModel)},
* {@link #removeForceModels()})</li>
* <li>the discrete events that should be triggered during propagation (
* {@link #addEventDetector(org.orekit.propagation.events.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.
* The central attraction coefficient used to define the initial orbit will be used.
* However, specifying only the initial state would mean the propagator would use
* only Keplerian forces. In this case, the simpler
* {@link org.orekit.propagation.analytical.KeplerianPropagator KeplerianPropagator}
* class would 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 six elements double array.
* These six elements are:
* <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>)
* in meters and radians,</li>
* </ul>
*
* <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 DSSTForceModel
* @author Romain Di Costanzo
* @author Pascal Parraud
*/
public class DSSTPropagator extends AbstractIntegratedPropagator {
/** Retrograde factor I.
* <p>
* DSST model needs equinoctial orbit as internal representation.
* Classical equinoctial elements have discontinuities when inclination
* is close to zero. In this representation, I = +1. <br>
* To avoid this discontinuity, another representation exists and equinoctial
* elements can be expressed in a different way, called "retrograde" orbit.
* This implies I = -1. <br>
* As Orekit doesn't implement the retrograde orbit, I is always set to +1.
* But for the sake of consistency with the theory, the retrograde factor
* has been kept in the formulas.
* </p>
*/
private static final int I = 1;
/** Default value for epsilon. */
private static final double EPSILON_DEFAULT = 1.0e-13;
/** Default value for maxIterations. */
private static final int MAX_ITERATIONS_DEFAULT = 200;
/** Number of grid points per integration step to be used in interpolation of short periodics coefficients.*/
private static final int INTERPOLATION_POINTS_PER_STEP = 3;
/** Flag specifying whether the initial orbital state is given with osculating elements. */
private boolean initialIsOsculating;
/** Force models used to compute short periodic terms. */
private final transient List<DSSTForceModel> forceModels;
/** State mapper holding the force models. */
private MeanPlusShortPeriodicMapper mapper;
/** Generator for the interpolation grid. */
private InterpolationGrid interpolationgrid;
/** Create a new instance of DSSTPropagator.
* <p>
* After creation, 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.
* </p>
*
* <p>This constructor uses the {@link DataContext#getDefault() default data context}.
*
* @param integrator numerical integrator to use for propagation.
* @param propagationType type of orbit to output (mean or osculating).
* @see #DSSTPropagator(ODEIntegrator, PropagationType, AttitudeProvider)
*/
@DefaultDataContext
public DSSTPropagator(final ODEIntegrator integrator, final PropagationType propagationType) {
this(integrator, propagationType,
Propagator.getDefaultLaw(DataContext.getDefault().getFrames()));
}
/** Create a new instance of DSSTPropagator.
* <p>
* After creation, 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.
* </p>
* @param integrator numerical integrator to use for propagation.
* @param propagationType type of orbit to output (mean or osculating).
* @param attitudeProvider the attitude law.
* @since 10.1
*/
public DSSTPropagator(final ODEIntegrator integrator,
final PropagationType propagationType,
final AttitudeProvider attitudeProvider) {
super(integrator, propagationType);
forceModels = new ArrayList<DSSTForceModel>();
initMapper();
// DSST uses only equinoctial orbits and mean longitude argument
setOrbitType(OrbitType.EQUINOCTIAL);
setPositionAngleType(PositionAngle.MEAN);
setAttitudeProvider(attitudeProvider);
setInterpolationGridToFixedNumberOfPoints(INTERPOLATION_POINTS_PER_STEP);
}
/** Create a new instance of DSSTPropagator.
* <p>
* After creation, 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. Only the mean orbits
* will be generated.
* </p>
*
* <p>This constructor uses the {@link DataContext#getDefault() default data context}.
*
* @param integrator numerical integrator to use for propagation.
* @see #DSSTPropagator(ODEIntegrator, PropagationType, AttitudeProvider)
*/
@DefaultDataContext
public DSSTPropagator(final ODEIntegrator integrator) {
this(integrator, PropagationType.MEAN);
}
/** Set the central attraction coefficient μ.
* <p>
* Setting the central attraction coefficient is
* equivalent to {@link #addForceModel(DSSTForceModel) add}
* a {@link DSSTNewtonianAttraction} force model.
* </p>
* @param mu central attraction coefficient (m³/s²)
* @see #addForceModel(DSSTForceModel)
* @see #getAllForceModels()
*/
public void setMu(final double mu) {
addForceModel(new DSSTNewtonianAttraction(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 DSSTNewtonianAttraction;
}
/** Set the initial state with osculating orbital elements.
* @param initialState initial state (defined with osculating elements)
*/
public void setInitialState(final SpacecraftState initialState) {
setInitialState(initialState, PropagationType.OSCULATING);
}
/** Set the initial state.
* @param initialState initial state
* @param stateType defined if the orbital state is defined with osculating or mean elements
*/
public void setInitialState(final SpacecraftState initialState,
final PropagationType stateType) {
switch (stateType) {
case MEAN:
initialIsOsculating = false;
break;
case OSCULATING:
initialIsOsculating = true;
break;
default:
throw new OrekitInternalError(null);
}
resetInitialState(initialState);
}
/** Reset the initial state.
*
* @param state new initial state
*/
@Override
public void resetInitialState(final SpacecraftState state) {
super.resetInitialState(state);
if (!hasNewtonianAttraction()) {
// use the state to define central attraction
setMu(state.getMu());
}
super.setStartDate(state.getDate());
}
/** Set the selected short periodic coefficients that must be stored as additional states.
* @param selectedCoefficients short periodic coefficients that must be stored as additional states
* (null means no coefficients are selected, empty set means all coefficients are selected)
*/
public void setSelectedCoefficients(final Set<String> selectedCoefficients) {
mapper.setSelectedCoefficients(selectedCoefficients == null ?
null : new HashSet<String>(selectedCoefficients));
}
/** Get the selected short periodic coefficients that must be stored as additional states.
* @return short periodic coefficients that must be stored as additional states
* (null means no coefficients are selected, empty set means all coefficients are selected)
*/
public Set<String> getSelectedCoefficients() {
final Set<String> set = mapper.getSelectedCoefficients();
return set == null ? null : Collections.unmodifiableSet(set);
}
/** Check if the initial state is provided in osculating elements.
* @return true if initial state is provided in osculating elements
*/
public boolean initialIsOsculating() {
return initialIsOsculating;
}
/** Set the interpolation grid generator.
* <p>
* The generator will create an interpolation grid with a fixed
* number of points for each mean element integration step.
* </p>
* <p>
* If neither {@link #setInterpolationGridToFixedNumberOfPoints(int)}
* nor {@link #setInterpolationGridToMaxTimeGap(double)} has been called,
* by default the propagator is set as to 3 interpolations points per step.
* </p>
* @param interpolationPoints number of interpolation points at
* each integration step
* @see #setInterpolationGridToMaxTimeGap(double)
* @since 7.1
*/
public void setInterpolationGridToFixedNumberOfPoints(final int interpolationPoints) {
interpolationgrid = new FixedNumberInterpolationGrid(interpolationPoints);
}
/** Set the interpolation grid generator.
* <p>
* The generator will create an interpolation grid with a maximum
* time gap between interpolation points.
* </p>
* <p>
* If neither {@link #setInterpolationGridToFixedNumberOfPoints(int)}
* nor {@link #setInterpolationGridToMaxTimeGap(double)} has been called,
* by default the propagator is set as to 3 interpolations points per step.
* </p>
* @param maxGap maximum time gap between interpolation points (seconds)
* @see #setInterpolationGridToFixedNumberOfPoints(int)
* @since 7.1
*/
public void setInterpolationGridToMaxTimeGap(final double maxGap) {
interpolationgrid = new MaxGapInterpolationGrid(maxGap);
}
/** 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 force perturbing {@link DSSTForceModel force} to add
* @see #removeForceModels()
* @see #setMu(double)
*/
public void addForceModel(final DSSTForceModel force) {
if (force instanceof DSSTNewtonianAttraction) {
// we want to add the central attraction force model
// ensure we are notified of any mu change
force.getParametersDrivers()[0].addObserver(new ParameterObserver() {
/** {@inheritDoc} */
@Override
public void valueChanged(final double previousValue, final ParameterDriver driver) {
superSetMu(driver.getValue());
}
});
if (hasNewtonianAttraction()) {
// there is already a central attraction model, replace it
forceModels.set(forceModels.size() - 1, force);
} else {
// there are no central attraction model yet, add it at the end of the list
forceModels.add(force);
}
} 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, force);
} else {
// we only have perturbing force models up to now, just append at the end of the list
forceModels.add(force);
}
}
force.registerAttitudeProvider(getAttitudeProvider());
}
/** Remove all perturbing force models from the global perturbation model
* (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(DSSTForceModel)
*/
public void removeForceModels() {
final int last = forceModels.size() - 1;
if (hasNewtonianAttraction()) {
// preserve the Newtonian attraction model at the end
final DSSTForceModel 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(DSSTForceModel)
* @see #setMu(double)
*/
public List<DSSTForceModel> getAllForceModels() {
return Collections.unmodifiableList(forceModels);
}
/** Get propagation parameter type.
* @return orbit type used for propagation
*/
public OrbitType getOrbitType() {
return super.getOrbitType();
}
/** Get propagation parameter type.
* @return angle type to use for propagation
*/
public PositionAngle getPositionAngleType() {
return super.getPositionAngleType();
}
/** Conversion from mean to osculating orbit.
* <p>
* Compute osculating state <b>in a DSST sense</b>, corresponding to the
* mean SpacecraftState in input, and according to the Force models taken
* into account.
* </p><p>
* Since the osculating state is obtained by adding short-periodic variation
* of each force model, the resulting output will depend on the
* force models parameterized in input.
* </p>
* @param mean Mean state to convert
* @param forces Forces to take into account
* @param attitudeProvider attitude provider (may be null if there are no Gaussian force models
* like atmospheric drag, radiation pressure or specific user-defined models)
* @return osculating state in a DSST sense
*/
public static SpacecraftState computeOsculatingState(final SpacecraftState mean,
final AttitudeProvider attitudeProvider,
final Collection<DSSTForceModel> forces) {
//Create the auxiliary object
final AuxiliaryElements aux = new AuxiliaryElements(mean.getOrbit(), I);
// Set the force models
final List<ShortPeriodTerms> shortPeriodTerms = new ArrayList<ShortPeriodTerms>();
for (final DSSTForceModel force : forces) {
force.registerAttitudeProvider(attitudeProvider);
shortPeriodTerms.addAll(force.initialize(aux, PropagationType.OSCULATING, force.getParameters()));
force.updateShortPeriodTerms(force.getParameters(), mean);
}
final EquinoctialOrbit osculatingOrbit = computeOsculatingOrbit(mean, shortPeriodTerms);
return new SpacecraftState(osculatingOrbit, mean.getAttitude(), mean.getMass(),
mean.getAdditionalStates());
}
/** Conversion from osculating to mean orbit.
* <p>
* Compute mean state <b>in a DSST sense</b>, corresponding to the
* osculating SpacecraftState in input, and according to the Force models
* taken into account.
* </p><p>
* Since the osculating state is obtained with the computation of
* short-periodic variation of each force model, the resulting output will
* depend on the force models parameterized in input.
* </p><p>
* The computation is done through a fixed-point iteration process.
* </p>
* @param osculating Osculating state to convert
* @param attitudeProvider attitude provider (may be null if there are no Gaussian force models
* like atmospheric drag, radiation pressure or specific user-defined models)
* @param forceModels Forces to take into account
* @return mean state in a DSST sense
*/
public static SpacecraftState computeMeanState(final SpacecraftState osculating,
final AttitudeProvider attitudeProvider,
final Collection<DSSTForceModel> forceModels) {
return computeMeanState(osculating, attitudeProvider, forceModels, EPSILON_DEFAULT, MAX_ITERATIONS_DEFAULT);
}
/** Conversion from osculating to mean orbit.
* <p>
* Compute mean state <b>in a DSST sense</b>, corresponding to the
* osculating SpacecraftState in input, and according to the Force models
* taken into account.
* </p><p>
* Since the osculating state is obtained with the computation of
* short-periodic variation of each force model, the resulting output will
* depend on the force models parameterized in input.
* </p><p>
* The computation is done through a fixed-point iteration process.
* </p>
* @param osculating Osculating state to convert
* @param attitudeProvider attitude provider (may be null if there are no Gaussian force models
* like atmospheric drag, radiation pressure or specific user-defined models)
* @param forceModels Forces to take into account
* @param epsilon convergence threshold for mean parameters conversion
* @param maxIterations maximum iterations for mean parameters conversion
* @return mean state in a DSST sense
* @since 10.1
*/
public static SpacecraftState computeMeanState(final SpacecraftState osculating,
final AttitudeProvider attitudeProvider,
final Collection<DSSTForceModel> forceModels,
final double epsilon,
final int maxIterations) {
final Orbit meanOrbit = computeMeanOrbit(osculating, attitudeProvider, forceModels, epsilon, maxIterations);
return new SpacecraftState(meanOrbit, osculating.getAttitude(), osculating.getMass(), osculating.getAdditionalStates());
}
/** Override the default value of the parameter.
* <p>
* By default, if the initial orbit is defined as osculating,
* it will be averaged over 2 satellite revolutions.
* This can be changed by using this method.
* </p>
* @param satelliteRevolution number of satellite revolutions to use for converting osculating to mean
* elements
*/
public void setSatelliteRevolution(final int satelliteRevolution) {
mapper.setSatelliteRevolution(satelliteRevolution);
}
/** Get the number of satellite revolutions to use for converting osculating to mean elements.
* @return number of satellite revolutions to use for converting osculating to mean elements
*/
public int getSatelliteRevolution() {
return mapper.getSatelliteRevolution();
}
/** {@inheritDoc} */
@Override
public void setAttitudeProvider(final AttitudeProvider attitudeProvider) {
super.setAttitudeProvider(attitudeProvider);
//Register the attitude provider for each force model
for (final DSSTForceModel force : forceModels) {
force.registerAttitudeProvider(attitudeProvider);
}
}
/** Method called just before integration.
* <p>
* The default implementation does nothing, it may be specialized in subclasses.
* </p>
* @param initialState initial state
* @param tEnd target date at which state should be propagated
*/
@Override
protected void beforeIntegration(final SpacecraftState initialState,
final AbsoluteDate tEnd) {
// check if only mean elements must be used
final PropagationType type = isMeanOrbit();
// compute common auxiliary elements
final AuxiliaryElements aux = new AuxiliaryElements(initialState.getOrbit(), I);
// initialize all perturbing forces
final List<ShortPeriodTerms> shortPeriodTerms = new ArrayList<ShortPeriodTerms>();
for (final DSSTForceModel force : forceModels) {
shortPeriodTerms.addAll(force.initialize(aux, type, force.getParameters()));
}
mapper.setShortPeriodTerms(shortPeriodTerms);
// if required, insert the special short periodics step handler
if (type == PropagationType.OSCULATING) {
final ShortPeriodicsHandler spHandler = new ShortPeriodicsHandler(forceModels);
final Collection<ODEStepHandler> stepHandlers = new ArrayList<ODEStepHandler>();
stepHandlers.add(spHandler);
final ODEIntegrator integrator = getIntegrator();
final Collection<ODEStepHandler> existing = integrator.getStepHandlers();
stepHandlers.addAll(existing);
integrator.clearStepHandlers();
// add back the existing handlers after the short periodics one
for (final ODEStepHandler sp : stepHandlers) {
integrator.addStepHandler(sp);
}
}
}
/** {@inheritDoc} */
@Override
protected void afterIntegration() {
// remove the special short periodics step handler if added before
if (isMeanOrbit() == PropagationType.OSCULATING) {
final List<ODEStepHandler> preserved = new ArrayList<ODEStepHandler>();
final ODEIntegrator integrator = getIntegrator();
for (final ODEStepHandler sp : integrator.getStepHandlers()) {
if (!(sp instanceof ShortPeriodicsHandler)) {
preserved.add(sp);
}
}
// clear the list
integrator.clearStepHandlers();
// add back the step handlers that were important for the user
for (final ODEStepHandler sp : preserved) {
integrator.addStepHandler(sp);
}
}
}
/** Compute mean state from osculating state.
* <p>
* Compute in a DSST sense the mean state corresponding to the input osculating state.
* </p><p>
* The computing is done through a fixed-point iteration process.
* </p>
* @param osculating initial osculating state
* @param attitudeProvider attitude provider (may be null if there are no Gaussian force models
* like atmospheric drag, radiation pressure or specific user-defined models)
* @param forceModels force models
* @param epsilon convergence threshold for mean parameters conversion
* @param maxIterations maximum iterations for mean parameters conversion
* @return mean state
*/
private static Orbit computeMeanOrbit(final SpacecraftState osculating,
final AttitudeProvider attitudeProvider,
final Collection<DSSTForceModel> forceModels, final double epsilon, final int maxIterations) {
// rough initialization of the mean parameters
EquinoctialOrbit meanOrbit = (EquinoctialOrbit) OrbitType.EQUINOCTIAL.convertType(osculating.getOrbit());
// threshold for each parameter
final double thresholdA = epsilon * (1 + FastMath.abs(meanOrbit.getA()));
final double thresholdE = epsilon * (1 + meanOrbit.getE());
final double thresholdI = epsilon * (1 + meanOrbit.getI());
final double thresholdL = epsilon * FastMath.PI;
// ensure all Gaussian force models can rely on attitude
for (final DSSTForceModel force : forceModels) {
force.registerAttitudeProvider(attitudeProvider);
}
int i = 0;
while (i++ < maxIterations) {
final SpacecraftState meanState = new SpacecraftState(meanOrbit, osculating.getAttitude(), osculating.getMass());
//Create the auxiliary object
final AuxiliaryElements aux = new AuxiliaryElements(meanOrbit, I);
// Set the force models
final List<ShortPeriodTerms> shortPeriodTerms = new ArrayList<ShortPeriodTerms>();
for (final DSSTForceModel force : forceModels) {
shortPeriodTerms.addAll(force.initialize(aux, PropagationType.OSCULATING, force.getParameters()));
force.updateShortPeriodTerms(force.getParameters(), meanState);
}
// recompute the osculating parameters from the current mean parameters
final EquinoctialOrbit rebuilt = computeOsculatingOrbit(meanState, shortPeriodTerms);
// adapted parameters residuals
final double deltaA = osculating.getA() - rebuilt.getA();
final double deltaEx = osculating.getEquinoctialEx() - rebuilt.getEquinoctialEx();
final double deltaEy = osculating.getEquinoctialEy() - rebuilt.getEquinoctialEy();
final double deltaHx = osculating.getHx() - rebuilt.getHx();
final double deltaHy = osculating.getHy() - rebuilt.getHy();
final double deltaLv = MathUtils.normalizeAngle(osculating.getLv() - rebuilt.getLv(), 0.0);
// check convergence
if (FastMath.abs(deltaA) < thresholdA &&
FastMath.abs(deltaEx) < thresholdE &&
FastMath.abs(deltaEy) < thresholdE &&
FastMath.abs(deltaHx) < thresholdI &&
FastMath.abs(deltaHy) < thresholdI &&
FastMath.abs(deltaLv) < thresholdL) {
return meanOrbit;
}
// update mean parameters
meanOrbit = new EquinoctialOrbit(meanOrbit.getA() + deltaA,
meanOrbit.getEquinoctialEx() + deltaEx,
meanOrbit.getEquinoctialEy() + deltaEy,
meanOrbit.getHx() + deltaHx,
meanOrbit.getHy() + deltaHy,
meanOrbit.getLv() + deltaLv,
PositionAngle.TRUE, meanOrbit.getFrame(),
meanOrbit.getDate(), meanOrbit.getMu());
}
throw new OrekitException(OrekitMessages.UNABLE_TO_COMPUTE_DSST_MEAN_PARAMETERS, i);
}
/** Compute osculating state from mean state.
* <p>
* Compute and add the short periodic variation to the mean {@link SpacecraftState}.
* </p>
* @param meanState initial mean state
* @param shortPeriodTerms short period terms
* @return osculating state
*/
private static EquinoctialOrbit computeOsculatingOrbit(final SpacecraftState meanState,
final List<ShortPeriodTerms> shortPeriodTerms) {
final double[] mean = new double[6];
final double[] meanDot = new double[6];
OrbitType.EQUINOCTIAL.mapOrbitToArray(meanState.getOrbit(), PositionAngle.MEAN, mean, meanDot);
final double[] y = mean.clone();
for (final ShortPeriodTerms spt : shortPeriodTerms) {
final double[] shortPeriodic = spt.value(meanState.getOrbit());
for (int i = 0; i < shortPeriodic.length; i++) {
y[i] += shortPeriodic[i];
}
}
return (EquinoctialOrbit) OrbitType.EQUINOCTIAL.mapArrayToOrbit(y, meanDot,
PositionAngle.MEAN, meanState.getDate(),
meanState.getMu(), meanState.getFrame());
}
/** {@inheritDoc} */
@Override
protected SpacecraftState getInitialIntegrationState() {
if (initialIsOsculating) {
// the initial state is an osculating state,
// it must be converted to mean state
return computeMeanState(getInitialState(), getAttitudeProvider(), forceModels);
} else {
// the initial state is already a mean state
return getInitialState();
}
}
/** {@inheritDoc}
* <p>
* Note that for DSST, orbit type is hardcoded to {@link OrbitType#EQUINOCTIAL}
* and position angle type is hardcoded to {@link PositionAngle#MEAN}, so
* the corresponding parameters are ignored.
* </p>
*/
@Override
protected StateMapper createMapper(final AbsoluteDate referenceDate, final double mu,
final OrbitType ignoredOrbitType, final PositionAngle ignoredPositionAngleType,
final AttitudeProvider attitudeProvider, final Frame frame) {
// create a mapper with the common settings provided as arguments
final MeanPlusShortPeriodicMapper newMapper =
new MeanPlusShortPeriodicMapper(referenceDate, mu, attitudeProvider, frame);
// copy the specific settings from the existing mapper
if (mapper != null) {
newMapper.setSatelliteRevolution(mapper.getSatelliteRevolution());
newMapper.setSelectedCoefficients(mapper.getSelectedCoefficients());
newMapper.setShortPeriodTerms(mapper.getShortPeriodTerms());
}
mapper = newMapper;
return mapper;
}
/** Internal mapper using mean parameters plus short periodic terms. */
private static class MeanPlusShortPeriodicMapper extends StateMapper {
/** Short periodic coefficients that must be stored as additional states. */
private Set<String> selectedCoefficients;
/** Number of satellite revolutions in the averaging interval. */
private int satelliteRevolution;
/** Short period terms. */
private List<ShortPeriodTerms> shortPeriodTerms;
/** Simple constructor.
* @param referenceDate reference date
* @param mu central attraction coefficient (m³/s²)
* @param attitudeProvider attitude provider
* @param frame inertial frame
*/
MeanPlusShortPeriodicMapper(final AbsoluteDate referenceDate, final double mu,
final AttitudeProvider attitudeProvider, final Frame frame) {
super(referenceDate, mu, OrbitType.EQUINOCTIAL, PositionAngle.MEAN, attitudeProvider, frame);
this.selectedCoefficients = null;
// Default averaging period for conversion from osculating to mean elements
this.satelliteRevolution = 2;
this.shortPeriodTerms = Collections.emptyList();
}
/** {@inheritDoc} */
@Override
public SpacecraftState mapArrayToState(final AbsoluteDate date,
final double[] y, final double[] yDot,
final PropagationType type) {
// add short periodic variations to mean elements to get osculating elements
// (the loop may not be performed if there are no force models and in the
// case we want to remain in mean parameters only)
final double[] elements = y.clone();
final Map<String, double[]> coefficients;
switch (type) {
case MEAN:
coefficients = null;
break;
case OSCULATING:
final Orbit meanOrbit = OrbitType.EQUINOCTIAL.mapArrayToOrbit(elements, yDot, PositionAngle.MEAN, date, getMu(), getFrame());
coefficients = selectedCoefficients == null ? null : new HashMap<String, double[]>();
for (final ShortPeriodTerms spt : shortPeriodTerms) {
final double[] shortPeriodic = spt.value(meanOrbit);
for (int i = 0; i < shortPeriodic.length; i++) {
elements[i] += shortPeriodic[i];
}
if (selectedCoefficients != null) {
coefficients.putAll(spt.getCoefficients(date, selectedCoefficients));
}
}
break;
default:
throw new OrekitInternalError(null);
}
final double mass = elements[6];
if (mass <= 0.0) {
throw new OrekitException(OrekitMessages.SPACECRAFT_MASS_BECOMES_NEGATIVE, mass);
}
final Orbit orbit = OrbitType.EQUINOCTIAL.mapArrayToOrbit(elements, yDot, PositionAngle.MEAN, date, getMu(), getFrame());
final Attitude attitude = getAttitudeProvider().getAttitude(orbit, date, getFrame());
if (coefficients == null) {
return new SpacecraftState(orbit, attitude, mass);
} else {
return new SpacecraftState(orbit, attitude, mass, coefficients);
}
}
/** {@inheritDoc} */
@Override
public void mapStateToArray(final SpacecraftState state, final double[] y, final double[] yDot) {
OrbitType.EQUINOCTIAL.mapOrbitToArray(state.getOrbit(), PositionAngle.MEAN, y, yDot);
y[6] = state.getMass();
}
/** Set the number of satellite revolutions to use for converting osculating to mean elements.
* <p>
* By default, if the initial orbit is defined as osculating,
* it will be averaged over 2 satellite revolutions.
* This can be changed by using this method.
* </p>
* @param satelliteRevolution number of satellite revolutions to use for converting osculating to mean
* elements
*/
public void setSatelliteRevolution(final int satelliteRevolution) {
this.satelliteRevolution = satelliteRevolution;
}
/** Get the number of satellite revolutions to use for converting osculating to mean elements.
* @return number of satellite revolutions to use for converting osculating to mean elements
*/
public int getSatelliteRevolution() {
return satelliteRevolution;
}
/** Set the selected short periodic coefficients that must be stored as additional states.
* @param selectedCoefficients short periodic coefficients that must be stored as additional states
* (null means no coefficients are selected, empty set means all coefficients are selected)
*/
public void setSelectedCoefficients(final Set<String> selectedCoefficients) {
this.selectedCoefficients = selectedCoefficients;
}
/** Get the selected short periodic coefficients that must be stored as additional states.
* @return short periodic coefficients that must be stored as additional states
* (null means no coefficients are selected, empty set means all coefficients are selected)
*/
public Set<String> getSelectedCoefficients() {
return selectedCoefficients;
}
/** Set the short period terms.
* @param shortPeriodTerms short period terms
* @since 7.1
*/
public void setShortPeriodTerms(final List<ShortPeriodTerms> shortPeriodTerms) {
this.shortPeriodTerms = shortPeriodTerms;
}
/** Get the short period terms.
* @return shortPeriodTerms short period terms
* @since 7.1
*/
public List<ShortPeriodTerms> getShortPeriodTerms() {
return shortPeriodTerms;
}
}
/** {@inheritDoc} */
@Override
protected MainStateEquations getMainStateEquations(final ODEIntegrator integrator) {
return new Main(integrator);
}
/** Internal class for mean parameters integration. */
private class Main implements MainStateEquations {
/** Derivatives array. */
private final double[] yDot;
/** Simple constructor.
* @param integrator numerical integrator to use for propagation.
*/
Main(final ODEIntegrator integrator) {
yDot = new double[7];
for (final DSSTForceModel forceModel : forceModels) {
final EventDetector[] modelDetectors = forceModel.getEventsDetectors();
if (modelDetectors != null) {
for (final EventDetector detector : modelDetectors) {
setUpEventDetector(integrator, detector);
}
}
}
}
/** {@inheritDoc} */
@Override
public double[] computeDerivatives(final SpacecraftState state) {
Arrays.fill(yDot, 0.0);
// compute common auxiliary elements
final AuxiliaryElements auxiliaryElements = new AuxiliaryElements(state.getOrbit(), I);
// compute the contributions of all perturbing forces
for (final DSSTForceModel forceModel : forceModels) {
final double[] daidt = elementRates(forceModel, state, auxiliaryElements, forceModel.getParameters());
for (int i = 0; i < daidt.length; i++) {
yDot[i] += daidt[i];
}
}
return yDot.clone();
}
/** This method allows to compute the mean equinoctial elements rates da<sub>i</sub> / dt
* for a specific force model.
* @param forceModel force to take into account
* @param state current state
* @param auxiliaryElements auxiliary elements related to the current orbit
* @param parameters force model parameters
* @return the mean equinoctial elements rates da<sub>i</sub> / dt
*/
private double[] elementRates(final DSSTForceModel forceModel,
final SpacecraftState state,
final AuxiliaryElements auxiliaryElements,
final double[] parameters) {
return forceModel.getMeanElementRate(state, auxiliaryElements, parameters);
}
}
/** Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
* <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 (m)
* @param orbit reference orbit
* @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) {
return NumericalPropagator.tolerances(dP, orbit, OrbitType.EQUINOCTIAL);
}
/** Step handler used to compute the parameters for the short periodic contributions.
* @author Lucian Barbulescu
*/
private class ShortPeriodicsHandler implements ODEStepHandler {
/** Force models used to compute short periodic terms. */
private final List<DSSTForceModel> forceModels;
/** Constructor.
* @param forceModels force models
*/
ShortPeriodicsHandler(final List<DSSTForceModel> forceModels) {
this.forceModels = forceModels;
}
/** {@inheritDoc} */
@Override
public void init(final ODEStateAndDerivative initialState, final double finalTime) {
// Build the mean state interpolated at initial point
final SpacecraftState meanStates = mapper.mapArrayToState(0.0,
initialState.getPrimaryState(),
initialState.getPrimaryDerivative(),
PropagationType.MEAN);
// Compute short periodic coefficients for this point
for (DSSTForceModel forceModel : forceModels) {
forceModel.updateShortPeriodTerms(forceModel.getParameters(), meanStates);
}
}
/** {@inheritDoc} */
@Override
public void handleStep(final ODEStateInterpolator interpolator, final boolean isLast) {
// Get the grid points to compute
final double[] interpolationPoints =
interpolationgrid.getGridPoints(interpolator.getPreviousState().getTime(),
interpolator.getCurrentState().getTime());
final SpacecraftState[] meanStates = new SpacecraftState[interpolationPoints.length];
for (int i = 0; i < interpolationPoints.length; ++i) {
// Build the mean state interpolated at grid point
final double time = interpolationPoints[i];
final ODEStateAndDerivative sd = interpolator.getInterpolatedState(time);
meanStates[i] = mapper.mapArrayToState(time,
sd.getPrimaryState(),
sd.getPrimaryDerivative(),
PropagationType.MEAN);
}
// Computate short periodic coefficients for this step
for (DSSTForceModel forceModel : forceModels) {
forceModel.updateShortPeriodTerms(forceModel.getParameters(), meanStates);
}
}
}
}