java.lang.Object
- org.orekit.propagation.AbstractPropagator
- - org.orekit.propagation.integration.AbstractIntegratedPropagator
  - - org.orekit.propagation.semianalytical.dsst.DSSTPropagator

All Implemented Interfaces:

Propagator, PVCoordinatesProvider
```
public class DSSTPropagator
extends AbstractIntegratedPropagator
```
This class propagates orbits using the DSST theory.
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.

The configuration parameters that can be set are:
- the initial spacecraft state (setInitialState(SpacecraftState))
- the various force models (addForceModel(DSSTForceModel), removeForceModels())
- the discrete events that should be triggered during propagation ( AbstractIntegratedPropagator.addEventDetector(org.orekit.propagation.events.EventDetector), AbstractIntegratedPropagator.clearEventsDetectors())
- the binding logic with the rest of the application (AbstractPropagator.getMultiplexer())
From these configuration parameters, only the initial state is mandatory. The default propagation settings are in equinoctial parameters with 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 KeplerianPropagator class would be more effective.

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.

The state that is seen by the integrator is a simple six elements double array. These six elements are:
- the equinoctial orbit parameters (a, e_x, e_y, h_x, h_y, λ_m) in meters and radians,
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 AbstractIntegratedPropagator.setResetAtEnd(boolean).

Beware the same instance cannot be used simultaneously by different threads, the class is not thread-safe.
Author:

Romain Di Costanzo, Pascal Parraud

See Also:

SpacecraftState, DSSTForceModel

Nested Class Summary
- Nested classes/interfaces inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator
  AbstractIntegratedPropagator.MainStateEquations

Field Summary
- Fields inherited from interface org.orekit.propagation.Propagator
  DEFAULT_MASS

Constructor Summary

Constructors
Constructor	Description
`DSSTPropagator(ODEIntegrator integrator)`	Create a new instance of DSSTPropagator.
`DSSTPropagator(ODEIntegrator integrator, PropagationType propagationType)`	Create a new instance of DSSTPropagator.
`DSSTPropagator(ODEIntegrator integrator, PropagationType propagationType, AttitudeProvider attitudeProvider)`	Create a new instance of DSSTPropagator.

Method Summary

All Methods Static Methods Instance Methods Concrete Methods
Modifier and Type	Method	Description
`void`	`addForceModel(DSSTForceModel force)`	Add a force model to the global perturbation model.
`protected void`	`afterIntegration()`	Method called just after integration.
`protected void`	`beforeIntegration(SpacecraftState initialState, AbsoluteDate tEnd)`	Method called just before integration.
`static SpacecraftState`	`computeMeanState(SpacecraftState osculating, AttitudeProvider attitudeProvider, Collection<DSSTForceModel> forceModels)`	Conversion from osculating to mean orbit.
`static SpacecraftState`	`computeMeanState(SpacecraftState osculating, AttitudeProvider attitudeProvider, Collection<DSSTForceModel> forceModels, double epsilon, int maxIterations)`	Conversion from osculating to mean orbit.
`static SpacecraftState`	`computeOsculatingState(SpacecraftState mean, AttitudeProvider attitudeProvider, Collection<DSSTForceModel> forces)`	Conversion from mean to osculating orbit.
`protected DSSTHarvester`	`createHarvester(String stmName, RealMatrix initialStm, DoubleArrayDictionary initialJacobianColumns)`	Create the harvester suitable for propagator.
`protected StateMapper`	`createMapper(AbsoluteDate referenceDate, double mu, OrbitType ignoredOrbitType, PositionAngleType ignoredPositionAngleType, AttitudeProvider attitudeProvider, Frame frame)`	Create a mapper between raw double components and spacecraft state.
`List<DSSTForceModel>`	`getAllForceModels()`	Get all the force models, perturbing forces and Newtonian attraction included.
`protected DSSTHarvester`	`getHarvester()`	Get the harvester.
`protected SpacecraftState`	`getInitialIntegrationState()`	Get the initial state for integration.
`protected List<String>`	`getJacobiansColumnsNames()`	Get the names of the parameters in the matrix returned by `MatricesHarvester.getParametersJacobian(org.orekit.propagation.SpacecraftState)`.
`protected AbstractIntegratedPropagator.MainStateEquations`	`getMainStateEquations(ODEIntegrator integrator)`	Get the differential equations to integrate (for main state only).
`OrbitType`	`getOrbitType()`	Get propagation parameter type.
`PositionAngleType`	`getPositionAngleType()`	Get propagation parameter type.
`int`	`getSatelliteRevolution()`	Get the number of satellite revolutions to use for converting osculating to mean elements.
`Set<String>`	`getSelectedCoefficients()`	Get the selected short periodic coefficients that must be stored as additional states.
`List<ShortPeriodTerms>`	`getShortPeriodTerms()`	Get the short periodic terms.
`double[]`	`getShortPeriodTermsValue(SpacecraftState meanState)`	Get the short period terms value.
`boolean`	`initialIsOsculating()`	Check if the initial state is provided in osculating elements.
`void`	`removeForceModels()`	Remove all perturbing force models from the global perturbation model (except central attraction).
`void`	`resetInitialState(SpacecraftState state)`	Reset the initial state.
`void`	`setAttitudeProvider(AttitudeProvider attitudeProvider)`	Set attitude provider.
`void`	`setInitialState(SpacecraftState initialState)`	Set the initial state with osculating orbital elements.
`void`	`setInitialState(SpacecraftState initialState, PropagationType stateType)`	Set the initial state.
`void`	`setInterpolationGridToFixedNumberOfPoints(int interpolationPoints)`	Set the interpolation grid generator.
`void`	`setInterpolationGridToMaxTimeGap(double maxGap)`	Set the interpolation grid generator.
`void`	`setMu(double mu)`	Set the central attraction coefficient μ.
`void`	`setSatelliteRevolution(int satelliteRevolution)`	Override the default value of the parameter.
`void`	`setSelectedCoefficients(Set<String> selectedCoefficients)`	Set the selected short periodic coefficients that must be stored as additional states.
`void`	`setShortPeriodTerms(List<ShortPeriodTerms> shortPeriodTerms)`	Override the default value short periodic terms.
`DSSTHarvester`	`setupMatricesComputation(String stmName, RealMatrix initialStm, DoubleArrayDictionary initialJacobianColumns)`	Set up computation of State Transition Matrix and Jacobians matrix with respect to parameters.
`protected void`	`setUpStmAndJacobianGenerators()`	Set up State Transition Matrix and Jacobian matrix handling.
`static double[][]`	`tolerances(double dP, double dV, Orbit orbit)`	Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
`static double[][]`	`tolerances(double dP, Orbit orbit)`	Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.

Methods inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator
addAdditionalDerivativesProvider, addEventDetector, clearEventsDetectors, getAdditionalDerivativesProviders, getBasicDimension, getCalls, getEphemerisGenerator, getEventsDetectors, getIntegrator, getIntegratorName, getManagedAdditionalStates, getMu, getPropagationType, getResetAtEnd, initMapper, isAdditionalStateManaged, propagate, propagate, setOrbitType, setPositionAngleType, setResetAtEnd, setUpEventDetector, setUpUserEventDetectors

Methods inherited from class org.orekit.propagation.AbstractPropagator
addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFrame, getInitialState, getMultiplexer, getPVCoordinates, getStartDate, initializeAdditionalStates, initializePropagation, setStartDate, stateChanged, updateAdditionalStates, updateUnmanagedStates

Methods inherited from class java.lang.Object
clone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, wait

Methods inherited from interface org.orekit.propagation.Propagator
clearStepHandlers, setStepHandler, setStepHandler

Methods inherited from interface org.orekit.utils.PVCoordinatesProvider
getPosition

- Constructor Detail
  - DSSTPropagator
```
@DefaultDataContext
public DSSTPropagator(ODEIntegrator integrator,
                      PropagationType propagationType)
```
    Create a new instance of DSSTPropagator.
    After creation, there are no perturbing forces at all. This means that if addForceModel is not called after creation, the integrated orbit will follow a Keplerian evolution only.
    
    This constructor uses the default data context.
    
    Parameters:
    
    integrator - numerical integrator to use for propagation.
    
    propagationType - type of orbit to output (mean or osculating).
    
    See Also:
    
    DSSTPropagator(ODEIntegrator, PropagationType, AttitudeProvider)
  - DSSTPropagator
```
public DSSTPropagator(ODEIntegrator integrator,
                      PropagationType propagationType,
                      AttitudeProvider attitudeProvider)
```
    Create a new instance of DSSTPropagator.
    After creation, there are no perturbing forces at all. This means that if addForceModel is not called after creation, the integrated orbit will follow a Keplerian evolution only.
    
    Parameters:
    
    integrator - numerical integrator to use for propagation.
    
    propagationType - type of orbit to output (mean or osculating).
    
    attitudeProvider - the attitude law.
    
    Since:
    
    10.1
  - DSSTPropagator
```
@DefaultDataContext
public DSSTPropagator(ODEIntegrator integrator)
```
    Create a new instance of DSSTPropagator.
    After creation, there are no perturbing forces at all. This means that if addForceModel is not called after creation, the integrated orbit will follow a Keplerian evolution only. Only the mean orbits will be generated.
    
    This constructor uses the default data context.
    
    Parameters:
    
    integrator - numerical integrator to use for propagation.
    
    See Also:
    
    DSSTPropagator(ODEIntegrator, PropagationType, AttitudeProvider)
- Method Detail
  - setMu
```
public void setMu(double mu)
```
    Set the central attraction coefficient μ.
    Setting the central attraction coefficient is equivalent to add a DSSTNewtonianAttraction force model.
    
    Overrides:
    
    setMu in class AbstractIntegratedPropagator
    
    Parameters:
    
    mu - central attraction coefficient (m³/s²)
    
    See Also:
    
    addForceModel(DSSTForceModel), getAllForceModels()
  - setInitialState
```
public void setInitialState(SpacecraftState initialState)
```
    Set the initial state with osculating orbital elements.
    
    Parameters:
    
    initialState - initial state (defined with osculating elements)
  - setInitialState
```
public void setInitialState(SpacecraftState initialState,
                            PropagationType stateType)
```
    Set the initial state.
    
    Parameters:
    
    initialState - initial state
    
    stateType - defined if the orbital state is defined with osculating or mean elements
  - resetInitialState
```
public void resetInitialState(SpacecraftState state)
```
    Reset the initial state.
    
    Specified by:
    
    resetInitialState in interface Propagator
    
    Overrides:
    
    resetInitialState in class AbstractPropagator
    
    Parameters:
    
    state - new initial state
  - setSelectedCoefficients
```
public void setSelectedCoefficients(Set<String> selectedCoefficients)
```
    Set the selected short periodic coefficients that must be stored as additional states.
    
    Parameters:
    
    selectedCoefficients - short periodic coefficients that must be stored as additional states (null means no coefficients are selected, empty set means all coefficients are selected)
  - getSelectedCoefficients
```
public Set<String> getSelectedCoefficients()
```
    Get the selected short periodic coefficients that must be stored as additional states.
    
    Returns:
    
    short periodic coefficients that must be stored as additional states (null means no coefficients are selected, empty set means all coefficients are selected)
  - getJacobiansColumnsNames
```
protected List<String> getJacobiansColumnsNames()
```
    Get the names of the parameters in the matrix returned by MatricesHarvester.getParametersJacobian(org.orekit.propagation.SpacecraftState).
    
    Returns:
    
    names of the parameters (i.e. columns) of the Jacobian matrix
  - setupMatricesComputation
```
public DSSTHarvester setupMatricesComputation(String stmName,
                                              RealMatrix initialStm,
                                              DoubleArrayDictionary initialJacobianColumns)
```
    Set up computation of State Transition Matrix and Jacobians matrix with respect to parameters.
    If this method is called, both State Transition Matrix and Jacobians with respect to the force models parameters that will be selected when propagation starts will be automatically computed, and the harvester will allow to retrieve them.
    
    The arguments for initial matrices must be compatible with the orbit type and position angle that will be used by the propagator.
    
    The default implementation throws an exception as the method is not supported by all propagators.
    
    Specified by:
    
    setupMatricesComputation in interface Propagator
    
    Overrides:
    
    setupMatricesComputation in class AbstractPropagator
    
    Parameters:
    
    stmName - State Transition Matrix state name
    
    initialStm - initial State Transition Matrix ∂Y/∂Y₀, if null (which is the most frequent case), assumed to be 6x6 identity
    
    initialJacobianColumns - initial columns of the Jacobians matrix with respect to parameters, if null or if some selected parameters are missing from the dictionary, the corresponding initial column is assumed to be 0
    
    Returns:
    
    harvester to retrieve computed matrices during and after propagation
  - createHarvester
```
protected DSSTHarvester createHarvester(String stmName,
                                        RealMatrix initialStm,
                                        DoubleArrayDictionary initialJacobianColumns)
```
    Create the harvester suitable for propagator.
    
    Overrides:
    
    createHarvester in class AbstractPropagator
    
    Parameters:
    
    stmName - State Transition Matrix state name
    
    initialStm - initial State Transition Matrix ∂Y/∂Y₀, if null (which is the most frequent case), assumed to be 6x6 identity
    
    initialJacobianColumns - initial columns of the Jacobians matrix with respect to parameters, if null or if some selected parameters are missing from the dictionary, the corresponding initial column is assumed to be 0
    
    Returns:
    
    harvester to retrieve computed matrices during and after propagation
  - getHarvester
```
protected DSSTHarvester getHarvester()
```
    Get the harvester.
    
    Overrides:
    
    getHarvester in class AbstractPropagator
    
    Returns:
    
    harvester, or null if it was not created
  - setUpStmAndJacobianGenerators
```
protected void setUpStmAndJacobianGenerators()
```
    Set up State Transition Matrix and Jacobian matrix handling.
    
    Overrides:
    
    setUpStmAndJacobianGenerators in class AbstractIntegratedPropagator
  - initialIsOsculating
```
public boolean initialIsOsculating()
```
    Check if the initial state is provided in osculating elements.
    
    Returns:
    
    true if initial state is provided in osculating elements
  - setInterpolationGridToFixedNumberOfPoints
```
public void setInterpolationGridToFixedNumberOfPoints(int interpolationPoints)
```
    Set the interpolation grid generator.
    The generator will create an interpolation grid with a fixed number of points for each mean element integration step.
    
    If neither setInterpolationGridToFixedNumberOfPoints(int) nor setInterpolationGridToMaxTimeGap(double) has been called, by default the propagator is set as to 3 interpolations points per step.
    
    Parameters:
    
    interpolationPoints - number of interpolation points at each integration step
    
    Since:
    
    7.1
    
    See Also:
    
    setInterpolationGridToMaxTimeGap(double)
  - setInterpolationGridToMaxTimeGap
```
public void setInterpolationGridToMaxTimeGap(double maxGap)
```
    Set the interpolation grid generator.
    The generator will create an interpolation grid with a maximum time gap between interpolation points.
    
    If neither setInterpolationGridToFixedNumberOfPoints(int) nor setInterpolationGridToMaxTimeGap(double) has been called, by default the propagator is set as to 3 interpolations points per step.
    
    Parameters:
    
    maxGap - maximum time gap between interpolation points (seconds)
    
    Since:
    
    7.1
    
    See Also:
    
    setInterpolationGridToFixedNumberOfPoints(int)
  - addForceModel
```
public void addForceModel(DSSTForceModel force)
```
    Add a force model to the global perturbation model.
    If this method is not called at all, the integrated orbit will follow a Keplerian evolution only.
    
    Parameters:
    
    force - perturbing force to add
    
    See Also:
    
    removeForceModels(), setMu(double)
  - removeForceModels
```
public void removeForceModels()
```
    Remove all perturbing force models from the global perturbation model (except central attraction).
    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.
    
    See Also:
    
    addForceModel(DSSTForceModel)
  - getAllForceModels
```
public List<DSSTForceModel> getAllForceModels()
```
    Get all the force models, perturbing forces and Newtonian attraction included.
    
    Returns:
    
    list of perturbing force models, with Newtonian attraction being the last one
    
    See Also:
    
    addForceModel(DSSTForceModel), setMu(double)
  - getOrbitType
```
public OrbitType getOrbitType()
```
    Get propagation parameter type.
    
    Overrides:
    
    getOrbitType in class AbstractIntegratedPropagator
    
    Returns:
    
    orbit type used for propagation
  - getPositionAngleType
```
public PositionAngleType getPositionAngleType()
```
    Get propagation parameter type.
    
    Overrides:
    
    getPositionAngleType in class AbstractIntegratedPropagator
    
    Returns:
    
    angle type to use for propagation
  - computeOsculatingState
```
public static SpacecraftState computeOsculatingState(SpacecraftState mean,
                                                     AttitudeProvider attitudeProvider,
                                                     Collection<DSSTForceModel> forces)
```
    Conversion from mean to osculating orbit.
    Compute osculating state in a DSST sense, corresponding to the mean SpacecraftState in input, and according to the Force models taken into account.
    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.
    
    Parameters:
    
    mean - Mean state to convert
    
    forces - Forces to take into account
    
    attitudeProvider - attitude provider (may be null if there are no Gaussian force models like atmospheric drag, radiation pressure or specific user-defined models)
    
    Returns:
    
    osculating state in a DSST sense
  - computeMeanState
```
public static SpacecraftState computeMeanState(SpacecraftState osculating,
                                               AttitudeProvider attitudeProvider,
                                               Collection<DSSTForceModel> forceModels)
```
    Conversion from osculating to mean orbit.
    Compute mean state in a DSST sense, corresponding to the osculating SpacecraftState in input, and according to the Force models taken into account.
    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.
    The computation is done through a fixed-point iteration process.
    
    Parameters:
    
    osculating - Osculating state to convert
    
    attitudeProvider - attitude provider (may be null if there are no Gaussian force models like atmospheric drag, radiation pressure or specific user-defined models)
    
    forceModels - Forces to take into account
    
    Returns:
    
    mean state in a DSST sense
  - computeMeanState
```
public static SpacecraftState computeMeanState(SpacecraftState osculating,
                                               AttitudeProvider attitudeProvider,
                                               Collection<DSSTForceModel> forceModels,
                                               double epsilon,
                                               int maxIterations)
```
    Conversion from osculating to mean orbit.
    Compute mean state in a DSST sense, corresponding to the osculating SpacecraftState in input, and according to the Force models taken into account.
    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.
    The computation is done through a fixed-point iteration process.
    
    Parameters:
    
    osculating - Osculating state to convert
    
    attitudeProvider - attitude provider (may be null if there are no Gaussian force models like atmospheric drag, radiation pressure or specific user-defined models)
    
    forceModels - Forces to take into account
    
    epsilon - convergence threshold for mean parameters conversion
    
    maxIterations - maximum iterations for mean parameters conversion
    
    Returns:
    
    mean state in a DSST sense
    
    Since:
    
    10.1
  - setSatelliteRevolution
```
public void setSatelliteRevolution(int satelliteRevolution)
```
    Override the default value of the parameter.
    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.
    
    Parameters:
    
    satelliteRevolution - number of satellite revolutions to use for converting osculating to mean elements
  - getSatelliteRevolution
```
public int getSatelliteRevolution()
```
    Get the number of satellite revolutions to use for converting osculating to mean elements.
    
    Returns:
    
    number of satellite revolutions to use for converting osculating to mean elements
  - setShortPeriodTerms
```
public void setShortPeriodTerms(List<ShortPeriodTerms> shortPeriodTerms)
```
    Override the default value short periodic terms.
    By default, short periodic terms are initialized before the numerical integration of the mean orbital elements.
    
    Parameters:
    
    shortPeriodTerms - short periodic terms
  - getShortPeriodTerms
```
public List<ShortPeriodTerms> getShortPeriodTerms()
```
    Get the short periodic terms.
    
    Returns:
    
    the short periodic terms
  - setAttitudeProvider
```
public void setAttitudeProvider(AttitudeProvider attitudeProvider)
```
    Set attitude provider.
    
    Specified by:
    
    setAttitudeProvider in interface Propagator
    
    Overrides:
    
    setAttitudeProvider in class AbstractIntegratedPropagator
    
    Parameters:
    
    attitudeProvider - attitude provider
  - beforeIntegration
```
protected void beforeIntegration(SpacecraftState initialState,
                                 AbsoluteDate tEnd)
```
    Method called just before integration.
    The default implementation does nothing, it may be specialized in subclasses.
    
    Overrides:
    
    beforeIntegration in class AbstractIntegratedPropagator
    
    Parameters:
    
    initialState - initial state
    
    tEnd - target date at which state should be propagated
  - afterIntegration
```
protected void afterIntegration()
```
    Method called just after integration.
    The default implementation does nothing, it may be specialized in subclasses.
    
    Overrides:
    
    afterIntegration in class AbstractIntegratedPropagator
  - getInitialIntegrationState
```
protected SpacecraftState getInitialIntegrationState()
```
    Get the initial state for integration.
    
    Overrides:
    
    getInitialIntegrationState in class AbstractIntegratedPropagator
    
    Returns:
    
    initial state for integration
  - createMapper
```
protected StateMapper createMapper(AbsoluteDate referenceDate,
                                   double mu,
                                   OrbitType ignoredOrbitType,
                                   PositionAngleType ignoredPositionAngleType,
                                   AttitudeProvider attitudeProvider,
                                   Frame frame)
```
    Create a mapper between raw double components and spacecraft state. /** Simple constructor.
    The position parameter type is meaningful only if propagation orbit type support it. As an example, it is not meaningful for propagation in Cartesian parameters.
    
    Note that for DSST, orbit type is hardcoded to OrbitType.EQUINOCTIAL and position angle type is hardcoded to PositionAngleType.MEAN, so the corresponding parameters are ignored.
    
    Specified by:
    
    createMapper in class AbstractIntegratedPropagator
    
    Parameters:
    
    referenceDate - reference date
    
    mu - central attraction coefficient (m³/s²)
    
    ignoredOrbitType - orbit type to use for mapping
    
    ignoredPositionAngleType - angle type to use for propagation
    
    attitudeProvider - attitude provider
    
    frame - inertial frame
    
    Returns:
    
    new mapper
  - getShortPeriodTermsValue
```
public double[] getShortPeriodTermsValue(SpacecraftState meanState)
```
    Get the short period terms value.
    
    Parameters:
    
    meanState - the mean state
    
    Returns:
    
    shortPeriodTerms short period terms
    
    Since:
    
    7.1
  - getMainStateEquations
```
protected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(ODEIntegrator integrator)
```
    Get the differential equations to integrate (for main state only).
    
    Specified by:
    
    getMainStateEquations in class AbstractIntegratedPropagator
    
    Parameters:
    
    integrator - numerical integrator to use for propagation.
    
    Returns:
    
    differential equations for main state
  - tolerances
```
public static double[][] tolerances(double dP,
                                    Orbit orbit)
```
    Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
    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):
    V r² |dV| = mu |dr|
    
    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.
    The tolerances are only orders of magnitude, 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.
    Parameters:
    
    dP - user specified position error (m)
    
    orbit - reference orbit
    
    Returns:
    
    a two rows array, row 0 being the absolute tolerance error and row 1 being the relative tolerance error
  - tolerances
```
public static double[][] tolerances(double dP,
                                    double dV,
                                    Orbit orbit)
```
    Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
    The errors are estimated from partial derivatives properties of orbits, starting from scalar position and velocity errors specified by the user.
    The tolerances are only orders of magnitude, 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.
    
    Parameters:
    
    dP - user specified position error (m)
    
    dV - user specified velocity error (m/s)
    
    orbit - reference orbit
    
    Returns:
    
    a two rows array, row 0 being the absolute tolerance error and row 1 being the relative tolerance error
    
    Since:
    
    10.3

Class DSSTPropagator

Nested Class Summary

Nested classes/interfaces inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator

Field Summary

Fields inherited from interface org.orekit.propagation.Propagator

Constructor Summary

Method Summary

Methods inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator

Methods inherited from class org.orekit.propagation.AbstractPropagator

Methods inherited from class java.lang.Object

Methods inherited from interface org.orekit.propagation.Propagator

Methods inherited from interface org.orekit.utils.PVCoordinatesProvider

Constructor Detail

DSSTPropagator

DSSTPropagator

DSSTPropagator

Method Detail

setMu

setInitialState

setInitialState

resetInitialState

setSelectedCoefficients

getSelectedCoefficients

getJacobiansColumnsNames

setupMatricesComputation

createHarvester

getHarvester

setUpStmAndJacobianGenerators

initialIsOsculating

setInterpolationGridToFixedNumberOfPoints

setInterpolationGridToMaxTimeGap

addForceModel

removeForceModels

getAllForceModels

getOrbitType

getPositionAngleType

computeOsculatingState

computeMeanState

computeMeanState

setSatelliteRevolution

getSatelliteRevolution

setShortPeriodTerms

getShortPeriodTerms

setAttitudeProvider

beforeIntegration

afterIntegration

getInitialIntegrationState

createMapper

getShortPeriodTermsValue

getMainStateEquations

tolerances

tolerances