Class NumericalPropagator
- 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.numerical.NumericalPropagator
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- All Implemented Interfaces:
Propagator
,PVCoordinatesProvider
public class NumericalPropagator extends AbstractIntegratedPropagator
This class propagatesorbits
using numerical integration.Numerical propagation is much more accurate than analytical propagation like for example
Keplerian
orEckstein-Hechler
, but requires a few more steps to set up to be used properly. Whereas analytical propagators are configured only thanks to their various constructors and can be used immediately after construction, numerical propagators configuration involve setting several parameters between construction time and propagation time.The configuration parameters that can be set are:
- the initial spacecraft state (
setInitialState(SpacecraftState)
) - the central attraction coefficient (
setMu(double)
) - the various force models (
addForceModel(ForceModel)
,removeForceModels()
) - the
type
of orbital parameters to be used for propagation (setOrbitType(OrbitType)
), - the
type
of position angle to be used in orbital parameters to be used for propagation where it is relevant (setPositionAngleType(PositionAngleType)
), - whether
state transition matrices and Jacobians matrices
should be propagated along with orbital state (AbstractPropagator.setupMatricesComputation(String, RealMatrix, DoubleArrayDictionary)
), - whether
additional derivatives
should be propagated along with orbital state (AbstractIntegratedPropagator.addAdditionalDerivativesProvider(AdditionalDerivativesProvider)
), - the discrete events that should be triggered during propagation
(
AbstractIntegratedPropagator.addEventDetector(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. If the central attraction coefficient is not explicitly specified, the one used to define the initial orbit will be used. However, specifying only the initial state and perhaps the central attraction coefficient would mean the propagator would use only Keplerian forces. In this case, the simplerKeplerianPropagator
class would perhaps 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 seven elements double array. The six first elements are either:
- the
equinoctial orbit parameters
(a, ex, ey, hx, hy, λM or λE or λv) in meters and radians, - the
Keplerian orbit parameters
(a, e, i, ω, Ω, M or E or v) in meters and radians, - the
circular orbit parameters
(a, ex, ey, i, Ω, αM or αE or αv) in meters and radians, - the
Cartesian orbit parameters
(x, y, z, vx, vy, vz) in meters and meters per seconds.
The last element is the mass in kilograms and changes only during thrusters firings
The following code snippet shows a typical setting for Low Earth Orbit propagation in equinoctial parameters and true longitude argument:
final double dP = 0.001; final double minStep = 0.001; final double maxStep = 500; final double initStep = 60; final double[][] tolerance = NumericalPropagator.tolerances(dP, orbit, OrbitType.EQUINOCTIAL); AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tolerance[0], tolerance[1]); integrator.setInitialStepSize(initStep); propagator = new NumericalPropagator(integrator);
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 changed by calling
AbstractIntegratedPropagator.setResetAtEnd(boolean)
.Beware the same instance cannot be used simultaneously by different threads, the class is not thread-safe.
- Author:
- Mathieu Roméro, Luc Maisonobe, Guylaine Prat, Fabien Maussion, Véronique Pommier-Maurussane
- See Also:
SpacecraftState
,ForceModel
,OrekitStepHandler
,OrekitFixedStepHandler
,IntegratedEphemeris
,TimeDerivativesEquations
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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 NumericalPropagator(ODEIntegrator integrator)
Create a new instance of NumericalPropagator, based on orbit definition mu.NumericalPropagator(ODEIntegrator integrator, AttitudeProvider attitudeProvider)
Create a new instance of NumericalPropagator, based on orbit definition mu.
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Method Summary
All Methods Static Methods Instance Methods Concrete Methods Modifier and Type Method Description void
addForceModel(ForceModel model)
Add a force model.protected void
beforeIntegration(SpacecraftState initialState, AbsoluteDate tEnd)
Method called just before integration.protected AbstractMatricesHarvester
createHarvester(String stmName, RealMatrix initialStm, DoubleArrayDictionary initialJacobianColumns)
Create the harvester suitable for propagator.protected StateMapper
createMapper(AbsoluteDate referenceDate, double mu, OrbitType orbitType, PositionAngleType positionAngleType, AttitudeProvider attitudeProvider, Frame frame)
Create a mapper between raw double components and spacecraft state.List<ForceModel>
getAllForceModels()
Get all the force models, perturbing forces and Newtonian attraction included.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.protected AttitudeProvider
initializeAttitudeProviderForDerivatives()
Method called when initializing the attitude provider used when evaluating derivatives.void
removeForceModels()
Remove all force models (except central attraction).void
resetInitialState(SpacecraftState state)
Reset the propagator initial state.void
setIgnoreCentralAttraction(boolean ignoreCentralAttraction)
Set the flag to ignore or not the creation of aNewtonianAttraction
.void
setInitialState(SpacecraftState initialState)
Set the initial state.void
setMu(double mu)
Set the central attraction coefficient μ.void
setOrbitType(OrbitType orbitType)
Set propagation orbit type.void
setPositionAngleType(PositionAngleType positionAngleType)
Set position angle type.protected void
setUpStmAndJacobianGenerators()
Set up State Transition Matrix and Jacobian matrix handling.static double[][]
tolerances(double dP, double dV, Orbit orbit, OrbitType type)
Estimate tolerance vectors for integrators when propagating in orbits.static double[][]
tolerances(double dP, Orbit orbit, OrbitType type)
Estimate tolerance vectors for integrators when propagating in orbits.static double[][]
tolerances(double dP, AbsolutePVCoordinates absPva)
Estimate tolerance vectors for integrators when propagating in absolute position-velocity-acceleration.-
Methods inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator
addAdditionalDerivativesProvider, addEventDetector, afterIntegration, clearEventsDetectors, getAdditionalDerivativesProviders, getBasicDimension, getCalls, getEphemerisGenerator, getEventsDetectors, getInitialIntegrationState, getIntegrator, getIntegratorName, getManagedAdditionalStates, getMu, getPropagationType, getResetAtEnd, initMapper, isAdditionalStateManaged, propagate, propagate, setAttitudeProvider, setResetAtEnd, setUpEventDetector, setUpUserEventDetectors
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Methods inherited from class org.orekit.propagation.AbstractPropagator
addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFrame, getHarvester, getInitialState, getMultiplexer, getStartDate, initializeAdditionalStates, initializePropagation, setStartDate, setupMatricesComputation, 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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NumericalPropagator
@DefaultDataContext public NumericalPropagator(ODEIntegrator integrator)
Create a new instance of NumericalPropagator, based on orbit definition mu. After creation, the instance is empty, i.e. the attitude provider is set to an unspecified default law and there are no perturbing forces at all. This means that ifaddForceModel
is not called after creation, the integrated orbit will follow a Keplerian evolution only. The defaults areOrbitType.EQUINOCTIAL
forpropagation orbit type
andPositionAngleType.TRUE
forposition angle type
.This constructor uses the
default data context
.- Parameters:
integrator
- numerical integrator to use for propagation.- See Also:
NumericalPropagator(ODEIntegrator, AttitudeProvider)
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NumericalPropagator
public NumericalPropagator(ODEIntegrator integrator, AttitudeProvider attitudeProvider)
Create a new instance of NumericalPropagator, based on orbit definition mu. After creation, the instance is empty, i.e. the attitude provider is set to an unspecified default law and there are no perturbing forces at all. This means that ifaddForceModel
is not called after creation, the integrated orbit will follow a Keplerian evolution only. The defaults areOrbitType.EQUINOCTIAL
forpropagation orbit type
andPositionAngleType.TRUE
forposition angle type
.- Parameters:
integrator
- numerical integrator to use for propagation.attitudeProvider
- the attitude law.- Since:
- 10.1
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Method Detail
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setIgnoreCentralAttraction
public void setIgnoreCentralAttraction(boolean ignoreCentralAttraction)
Set the flag to ignore or not the creation of aNewtonianAttraction
.- Parameters:
ignoreCentralAttraction
- if true,NewtonianAttraction
is not added automatically if missing
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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
aNewtonianAttraction
force model. *- Overrides:
setMu
in classAbstractIntegratedPropagator
- Parameters:
mu
- central attraction coefficient (m³/s²)- See Also:
addForceModel(ForceModel)
,getAllForceModels()
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addForceModel
public void addForceModel(ForceModel model)
Add a force model.If this method is not called at all, the integrated orbit will follow a Keplerian evolution only.
- Parameters:
model
-ForceModel
to add (it can be either a perturbing force model or an instance ofNewtonianAttraction
)- See Also:
removeForceModels()
,setMu(double)
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removeForceModels
public void removeForceModels()
Remove all force models (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(ForceModel)
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getAllForceModels
public List<ForceModel> 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(ForceModel)
,setMu(double)
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setOrbitType
public void setOrbitType(OrbitType orbitType)
Set propagation orbit type.- Overrides:
setOrbitType
in classAbstractIntegratedPropagator
- Parameters:
orbitType
- orbit type to use for propagation, null for propagating usingAbsolutePVCoordinates
rather thanOrbit
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getOrbitType
public OrbitType getOrbitType()
Get propagation parameter type.- Overrides:
getOrbitType
in classAbstractIntegratedPropagator
- Returns:
- orbit type used for propagation, null for
propagating using
AbsolutePVCoordinates
rather thanOrbit
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setPositionAngleType
public void setPositionAngleType(PositionAngleType positionAngleType)
Set position angle type.The position parameter type is meaningful only if
propagation orbit type
support it. As an example, it is not meaningful for propagation inCartesian
parameters.- Overrides:
setPositionAngleType
in classAbstractIntegratedPropagator
- Parameters:
positionAngleType
- angle type to use 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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setInitialState
public void setInitialState(SpacecraftState initialState)
Set the initial state.- Parameters:
initialState
- initial state
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resetInitialState
public void resetInitialState(SpacecraftState state)
Reset the propagator initial state.- Specified by:
resetInitialState
in interfacePropagator
- Overrides:
resetInitialState
in classAbstractPropagator
- Parameters:
state
- new initial state to consider
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createHarvester
protected AbstractMatricesHarvester 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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setUpStmAndJacobianGenerators
protected void setUpStmAndJacobianGenerators()
Set up State Transition Matrix and Jacobian matrix handling.- Overrides:
setUpStmAndJacobianGenerators
in classAbstractIntegratedPropagator
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initializeAttitudeProviderForDerivatives
protected AttitudeProvider initializeAttitudeProviderForDerivatives()
Method called when initializing the attitude provider used when evaluating derivatives.- Overrides:
initializeAttitudeProviderForDerivatives
in classAbstractIntegratedPropagator
- Returns:
- attitude provider for derivatives
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createMapper
protected StateMapper createMapper(AbsoluteDate referenceDate, double mu, OrbitType orbitType, PositionAngleType positionAngleType, 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.- Specified by:
createMapper
in classAbstractIntegratedPropagator
- Parameters:
referenceDate
- reference datemu
- central attraction coefficient (m³/s²)orbitType
- orbit type to use for mappingpositionAngleType
- angle type to use for propagationattitudeProvider
- attitude providerframe
- inertial frame- Returns:
- new mapper
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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, AbsolutePVCoordinates absPva)
Estimate tolerance vectors for integrators when propagating in absolute position-velocity-acceleration.- Parameters:
dP
- user specified position errorabsPva
- reference absolute position-velocity-acceleration- Returns:
- a two rows array, row 0 being the absolute tolerance error and row 1 being the relative tolerance error
- See Also:
tolerances(double, Orbit, OrbitType)
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tolerances
public static double[][] tolerances(double dP, Orbit orbit, OrbitType type)
Estimate tolerance vectors for integrators when propagating in orbits.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 errororbit
- reference orbittype
- propagation type for the meaning of the tolerance vectors elements (it may be different fromorbit.getType()
)- 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, OrbitType type)
Estimate tolerance vectors for integrators when propagating in orbits.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 errordV
- user specified velocity errororbit
- reference orbittype
- propagation type for the meaning of the tolerance vectors elements (it may be different fromorbit.getType()
)- 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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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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