Class DSSTPropagator
- java.lang.Object
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- org.orekit.propagation.AbstractPropagator
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- org.orekit.propagation.integration.AbstractIntegratedPropagator
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- org.orekit.propagation.semianalytical.dsst.DSSTPropagator
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- All Implemented Interfaces:
Propagator
,PVCoordinatesProvider
public class DSSTPropagator extends AbstractIntegratedPropagator
This class propagatesorbits
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 withtrue
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 simplerKeplerianPropagator
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, ex, ey, hx, hy, λm) 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
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Nested Class Summary
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Nested classes/interfaces inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator
AbstractIntegratedPropagator.MainStateEquations
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Field Summary
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Fields inherited from interface org.orekit.propagation.Propagator
DEFAULT_MASS
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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.
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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 byMatricesHarvester.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, initializeAttitudeProviderForDerivatives, initMapper, isAdditionalStateManaged, propagate, propagate, setOrbitType, setPositionAngleType, setResetAtEnd, setUpEventDetector, setUpUserEventDetectors
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Methods inherited from class org.orekit.propagation.AbstractPropagator
addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFrame, getInitialState, getMultiplexer, getStartDate, initializeAdditionalStates, initializePropagation, setStartDate, stateChanged, updateAdditionalStates, updateUnmanagedStates
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Methods inherited from class java.lang.Object
clone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, wait
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Methods inherited from interface org.orekit.propagation.Propagator
clearStepHandlers, getPosition, getPVCoordinates, setStepHandler, setStepHandler
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Constructor Detail
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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)
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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
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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)
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Method Detail
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setMu
public void setMu(double mu)
Set the central attraction coefficient μ.Setting the central attraction coefficient is equivalent to
add
aDSSTNewtonianAttraction
force model.- Overrides:
setMu
in classAbstractIntegratedPropagator
- Parameters:
mu
- central attraction coefficient (m³/s²)- See Also:
addForceModel(DSSTForceModel)
,getAllForceModels()
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setInitialState
public void setInitialState(SpacecraftState initialState)
Set the initial state with osculating orbital elements.- Parameters:
initialState
- initial state (defined with osculating elements)
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setInitialState
public void setInitialState(SpacecraftState initialState, PropagationType stateType)
Set the initial state.- Parameters:
initialState
- initial statestateType
- defined if the orbital state is defined with osculating or mean elements
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resetInitialState
public void resetInitialState(SpacecraftState state)
Reset the initial state.- Specified by:
resetInitialState
in interfacePropagator
- Overrides:
resetInitialState
in classAbstractPropagator
- Parameters:
state
- new initial state
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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)
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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)
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getJacobiansColumnsNames
protected List<String> getJacobiansColumnsNames()
Get the names of the parameters in the matrix returned byMatricesHarvester.getParametersJacobian(org.orekit.propagation.SpacecraftState)
.- Returns:
- names of the parameters (i.e. columns) of the Jacobian matrix
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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
andposition 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 interfacePropagator
- Overrides:
setupMatricesComputation
in classAbstractPropagator
- Parameters:
stmName
- State Transition Matrix state nameinitialStm
- initial State Transition Matrix ∂Y/∂Y₀, if null (which is the most frequent case), assumed to be 6x6 identityinitialJacobianColumns
- 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
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createHarvester
protected DSSTHarvester createHarvester(String stmName, RealMatrix initialStm, DoubleArrayDictionary initialJacobianColumns)
Create the harvester suitable for propagator.- Overrides:
createHarvester
in classAbstractPropagator
- Parameters:
stmName
- State Transition Matrix state nameinitialStm
- initial State Transition Matrix ∂Y/∂Y₀, if null (which is the most frequent case), assumed to be 6x6 identityinitialJacobianColumns
- 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
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getHarvester
protected DSSTHarvester getHarvester()
Get the harvester.- Overrides:
getHarvester
in classAbstractPropagator
- Returns:
- harvester, or null if it was not created
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setUpStmAndJacobianGenerators
protected void setUpStmAndJacobianGenerators()
Set up State Transition Matrix and Jacobian matrix handling.- Overrides:
setUpStmAndJacobianGenerators
in classAbstractIntegratedPropagator
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initialIsOsculating
public boolean initialIsOsculating()
Check if the initial state is provided in osculating elements.- Returns:
- true if initial state is provided in osculating elements
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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)
norsetInterpolationGridToMaxTimeGap(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)
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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)
norsetInterpolationGridToMaxTimeGap(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)
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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
- perturbingforce
to add- See Also:
removeForceModels()
,setMu(double)
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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)
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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)
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getOrbitType
public OrbitType getOrbitType()
Get propagation parameter type.- Overrides:
getOrbitType
in classAbstractIntegratedPropagator
- Returns:
- orbit type used for propagation
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getPositionAngleType
public PositionAngleType getPositionAngleType()
Get propagation parameter type.- Overrides:
getPositionAngleType
in classAbstractIntegratedPropagator
- Returns:
- angle type to use for propagation
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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 convertforces
- Forces to take into accountattitudeProvider
- 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
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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 convertattitudeProvider
- 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
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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 convertattitudeProvider
- 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 accountepsilon
- convergence threshold for mean parameters conversionmaxIterations
- maximum iterations for mean parameters conversion- Returns:
- mean state in a DSST sense
- Since:
- 10.1
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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
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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
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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
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getShortPeriodTerms
public List<ShortPeriodTerms> getShortPeriodTerms()
Get the short periodic terms.- Returns:
- the short periodic terms
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setAttitudeProvider
public void setAttitudeProvider(AttitudeProvider attitudeProvider)
Set attitude provider.- Specified by:
setAttitudeProvider
in interfacePropagator
- Overrides:
setAttitudeProvider
in classAbstractIntegratedPropagator
- Parameters:
attitudeProvider
- attitude provider
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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 classAbstractIntegratedPropagator
- Parameters:
initialState
- initial statetEnd
- target date at which state should be propagated
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afterIntegration
protected void afterIntegration()
Method called just after integration.The default implementation does nothing, it may be specialized in subclasses.
- Overrides:
afterIntegration
in classAbstractIntegratedPropagator
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getInitialIntegrationState
protected SpacecraftState getInitialIntegrationState()
Get the initial state for integration.- Overrides:
getInitialIntegrationState
in classAbstractIntegratedPropagator
- Returns:
- initial state for integration
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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 inCartesian
parameters.Note that for DSST, orbit type is hardcoded to
OrbitType.EQUINOCTIAL
and position angle type is hardcoded toPositionAngleType.MEAN
, so the corresponding parameters are ignored.- Specified by:
createMapper
in classAbstractIntegratedPropagator
- Parameters:
referenceDate
- reference datemu
- central attraction coefficient (m³/s²)ignoredOrbitType
- orbit type to use for mappingignoredPositionAngleType
- angle type to use for propagationattitudeProvider
- attitude providerframe
- inertial frame- Returns:
- new mapper
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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
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getMainStateEquations
protected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(ODEIntegrator integrator)
Get the differential equations to integrate (for main state only).- Specified by:
getMainStateEquations
in classAbstractIntegratedPropagator
- Parameters:
integrator
- numerical integrator to use for propagation.- Returns:
- differential equations for main state
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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
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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
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