public class DSSTPropagator extends AbstractIntegratedPropagator
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:
setInitialState(SpacecraftState))addForceModel(DSSTForceModel),
removeForceModels())AbstractIntegratedPropagator.addEventDetector(org.orekit.propagation.events.EventDetector),
AbstractIntegratedPropagator.clearEventsDetectors())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:
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.
SpacecraftState,
DSSTForceModelAbstractIntegratedPropagator.MainStateEquationsDEFAULT_MASS| Constructor and Description |
|---|
DSSTPropagator(ODEIntegrator integrator)
Create a new instance of DSSTPropagator.
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DSSTPropagator(ODEIntegrator integrator,
PropagationType propagationType)
Create a new instance of DSSTPropagator.
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DSSTPropagator(ODEIntegrator integrator,
PropagationType propagationType,
AttitudeProvider attitudeProvider)
Create a new instance of DSSTPropagator.
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| Modifier and Type | Method and Description |
|---|---|
void |
addForceModel(DSSTForceModel force)
Add a force model to the global perturbation model.
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protected void |
afterIntegration()
Method called just after integration.
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protected void |
beforeIntegration(SpacecraftState initialState,
AbsoluteDate tEnd)
Method called just before integration.
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static SpacecraftState |
computeMeanState(SpacecraftState osculating,
AttitudeProvider attitudeProvider,
Collection<DSSTForceModel> forceModels)
Conversion from osculating to mean orbit.
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static SpacecraftState |
computeMeanState(SpacecraftState osculating,
AttitudeProvider attitudeProvider,
Collection<DSSTForceModel> forceModels,
double epsilon,
int maxIterations)
Conversion from osculating to mean orbit.
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static SpacecraftState |
computeOsculatingState(SpacecraftState mean,
AttitudeProvider attitudeProvider,
Collection<DSSTForceModel> forces)
Conversion from mean to osculating orbit.
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protected AbstractMatricesHarvester |
createHarvester(String stmName,
RealMatrix initialStm,
DoubleArrayDictionary initialJacobianColumns)
Create the harvester suitable for propagator.
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protected StateMapper |
createMapper(AbsoluteDate referenceDate,
double mu,
OrbitType ignoredOrbitType,
PositionAngle ignoredPositionAngleType,
AttitudeProvider attitudeProvider,
Frame frame)
Create a mapper between raw double components and spacecraft state.
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List<DSSTForceModel> |
getAllForceModels()
Get all the force models, perturbing forces and Newtonian attraction included.
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protected SpacecraftState |
getInitialIntegrationState()
Get the initial state for integration.
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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).
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OrbitType |
getOrbitType()
Get propagation parameter type.
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PositionAngle |
getPositionAngleType()
Get propagation parameter type.
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int |
getSatelliteRevolution()
Get the number of satellite revolutions to use for converting osculating to mean elements.
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Set<String> |
getSelectedCoefficients()
Get the selected short periodic coefficients that must be stored as additional states.
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List<ShortPeriodTerms> |
getShortPeriodTerms()
Get the short periodic terms.
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double[] |
getShortPeriodTermsValue(SpacecraftState meanState)
Get the short period terms value.
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boolean |
initialIsOsculating()
Check if the initial state is provided in osculating elements.
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void |
removeForceModels()
Remove all perturbing force models from the global perturbation model
(except central attraction).
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void |
resetInitialState(SpacecraftState state)
Reset the initial state.
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void |
setAttitudeProvider(AttitudeProvider attitudeProvider)
Set attitude provider.
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void |
setInitialState(SpacecraftState initialState)
Set the initial state with osculating orbital elements.
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void |
setInitialState(SpacecraftState initialState,
PropagationType stateType)
Set the initial state.
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void |
setInterpolationGridToFixedNumberOfPoints(int interpolationPoints)
Set the interpolation grid generator.
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void |
setInterpolationGridToMaxTimeGap(double maxGap)
Set the interpolation grid generator.
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void |
setMu(double mu)
Set the central attraction coefficient μ.
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void |
setSatelliteRevolution(int satelliteRevolution)
Override the default value of the parameter.
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void |
setSelectedCoefficients(Set<String> selectedCoefficients)
Set the selected short periodic coefficients that must be stored as additional states.
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void |
setShortPeriodTerms(List<ShortPeriodTerms> shortPeriodTerms)
Override the default value short periodic terms.
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protected void |
setUpStmAndJacobianGenerators()
Set up State Transition Matrix and Jacobian matrix handling.
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static double[][] |
tolerances(double dP,
double dV,
Orbit orbit)
Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
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static double[][] |
tolerances(double dP,
Orbit orbit)
Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
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addAdditionalDerivativesProvider, addAdditionalEquations, addEventDetector, clearEventsDetectors, getAdditionalDerivativesProviders, getBasicDimension, getCalls, getEphemerisGenerator, getEventsDetectors, getIntegrator, getManagedAdditionalStates, getMu, getPropagationType, initMapper, isAdditionalStateManaged, isMeanOrbit, propagate, propagate, setOrbitType, setPositionAngleType, setResetAtEnd, setUpEventDetector, setUpUserEventDetectorsaddAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFrame, getHarvester, getInitialState, getMultiplexer, getPVCoordinates, getStartDate, initializePropagation, setStartDate, setupMatricesComputation, stateChanged, updateAdditionalStates, updateUnmanagedStatesclone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, waitclearStepHandlers, getDefaultLaw, setStepHandler, setStepHandler@DefaultDataContext public DSSTPropagator(ODEIntegrator integrator, PropagationType propagationType)
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.
integrator - numerical integrator to use for propagation.propagationType - type of orbit to output (mean or osculating).DSSTPropagator(ODEIntegrator, PropagationType, AttitudeProvider)public DSSTPropagator(ODEIntegrator integrator, PropagationType propagationType, AttitudeProvider attitudeProvider)
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.
integrator - numerical integrator to use for propagation.propagationType - type of orbit to output (mean or osculating).attitudeProvider - the attitude law.@DefaultDataContext public DSSTPropagator(ODEIntegrator integrator)
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.
integrator - numerical integrator to use for propagation.DSSTPropagator(ODEIntegrator, PropagationType, AttitudeProvider)public void setMu(double mu)
Setting the central attraction coefficient is
equivalent to add
a DSSTNewtonianAttraction force model.
setMu in class AbstractIntegratedPropagatormu - central attraction coefficient (m³/s²)addForceModel(DSSTForceModel),
getAllForceModels()public void setInitialState(SpacecraftState initialState)
initialState - initial state (defined with osculating elements)public void setInitialState(SpacecraftState initialState, PropagationType stateType)
initialState - initial statestateType - defined if the orbital state is defined with osculating or mean elementspublic void resetInitialState(SpacecraftState state)
resetInitialState in interface PropagatorresetInitialState in class AbstractPropagatorstate - new initial statepublic void setSelectedCoefficients(Set<String> selectedCoefficients)
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 Set<String> getSelectedCoefficients()
protected List<String> getJacobiansColumnsNames()
MatricesHarvester.getParametersJacobian(org.orekit.propagation.SpacecraftState).protected AbstractMatricesHarvester createHarvester(String stmName, RealMatrix initialStm, DoubleArrayDictionary initialJacobianColumns)
createHarvester in class AbstractPropagatorstmName - 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 0protected void setUpStmAndJacobianGenerators()
setUpStmAndJacobianGenerators in class AbstractIntegratedPropagatorpublic boolean initialIsOsculating()
public void setInterpolationGridToFixedNumberOfPoints(int interpolationPoints)
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.
interpolationPoints - number of interpolation points at
each integration stepsetInterpolationGridToMaxTimeGap(double)public void setInterpolationGridToMaxTimeGap(double maxGap)
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.
maxGap - maximum time gap between interpolation points (seconds)setInterpolationGridToFixedNumberOfPoints(int)public void addForceModel(DSSTForceModel force)
If this method is not called at all, the integrated orbit will follow a Keplerian evolution only.
force - perturbing force to addremoveForceModels(),
setMu(double)public void removeForceModels()
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.
addForceModel(DSSTForceModel)public List<DSSTForceModel> getAllForceModels()
addForceModel(DSSTForceModel),
setMu(double)public OrbitType getOrbitType()
getOrbitType in class AbstractIntegratedPropagatorpublic PositionAngle getPositionAngleType()
getPositionAngleType in class AbstractIntegratedPropagatorpublic static SpacecraftState computeOsculatingState(SpacecraftState mean, AttitudeProvider attitudeProvider, Collection<DSSTForceModel> forces)
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.
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)public static SpacecraftState computeMeanState(SpacecraftState osculating, AttitudeProvider attitudeProvider, Collection<DSSTForceModel> forceModels)
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.
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 accountpublic static SpacecraftState computeMeanState(SpacecraftState osculating, AttitudeProvider attitudeProvider, Collection<DSSTForceModel> forceModels, double epsilon, int maxIterations)
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.
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 conversionpublic void setSatelliteRevolution(int satelliteRevolution)
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.
satelliteRevolution - number of satellite revolutions to use for converting osculating to mean
elementspublic int getSatelliteRevolution()
public void setShortPeriodTerms(List<ShortPeriodTerms> shortPeriodTerms)
By default, short periodic terms are initialized before the numerical integration of the mean orbital elements.
shortPeriodTerms - short periodic termspublic List<ShortPeriodTerms> getShortPeriodTerms()
public void setAttitudeProvider(AttitudeProvider attitudeProvider)
setAttitudeProvider in interface PropagatorsetAttitudeProvider in class AbstractIntegratedPropagatorattitudeProvider - attitude providerprotected void beforeIntegration(SpacecraftState initialState, AbsoluteDate tEnd)
The default implementation does nothing, it may be specialized in subclasses.
beforeIntegration in class AbstractIntegratedPropagatorinitialState - initial statetEnd - target date at which state should be propagatedprotected void afterIntegration()
The default implementation does nothing, it may be specialized in subclasses.
afterIntegration in class AbstractIntegratedPropagatorprotected SpacecraftState getInitialIntegrationState()
getInitialIntegrationState in class AbstractIntegratedPropagatorprotected StateMapper createMapper(AbsoluteDate referenceDate, double mu, OrbitType ignoredOrbitType, PositionAngle ignoredPositionAngleType, AttitudeProvider attitudeProvider, Frame frame)
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 PositionAngle.MEAN, so
the corresponding parameters are ignored.
createMapper in class AbstractIntegratedPropagatorreferenceDate - reference datemu - central attraction coefficient (m³/s²)ignoredOrbitType - orbit type to use for mappingignoredPositionAngleType - angle type to use for propagationattitudeProvider - attitude providerframe - inertial framepublic double[] getShortPeriodTermsValue(SpacecraftState meanState)
meanState - the mean stateprotected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(ODEIntegrator integrator)
getMainStateEquations in class AbstractIntegratedPropagatorintegrator - numerical integrator to use for propagation.public static double[][] tolerances(double dP,
Orbit orbit)
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.
dP - user specified position error (m)orbit - reference orbitpublic static double[][] tolerances(double dP,
double dV,
Orbit orbit)
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.
dP - user specified position error (m)dV - user specified velocity error (m/s)orbit - reference orbitCopyright © 2002-2022 CS GROUP. All rights reserved.