org.orekit.propagation.semianalytical.dsst

## Class DSSTPropagator

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

AbstractIntegratedPropagator.MainStateEquations

• ### Fields inherited from interface org.orekit.propagation.Propagator

DEFAULT_LAW, DEFAULT_MASS, EPHEMERIS_GENERATION_MODE, MASTER_MODE, SLAVE_MODE
• ### Constructor Summary

Constructors
Constructor and 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
Modifier and Type Method and 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 StateMapper createMapper(AbsoluteDate referenceDate, double mu, OrbitType ignoredOrbitType, PositionAngle 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 SpacecraftState getInitialIntegrationState()
Get the initial state for integration.
protected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(ODEIntegrator integrator)
Get the differential equations to integrate (for main state only).
OrbitType getOrbitType()
Get propagation parameter type.
PositionAngle 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.
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.
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

addAdditionalEquations, addEventDetector, clearEventsDetectors, getBasicDimension, getCalls, getEventsDetectors, getGeneratedEphemeris, getIntegrator, getManagedAdditionalStates, getMu, initMapper, isAdditionalStateManaged, isMeanOrbit, propagate, propagate, propagate, setEphemerisMode, setEphemerisMode, setMasterMode, setOrbitType, setPositionAngleType, setResetAtEnd, setSlaveMode, setUpEventDetector, setUpUserEventDetectors
• ### Methods inherited from class org.orekit.propagation.AbstractPropagator

addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFixedStepSize, getFrame, getInitialState, getMode, getPVCoordinates, getStartDate, getStepHandler, initializePropagation, setMasterMode, setStartDate, stateChanged, updateAdditionalStates
• ### 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

getDefaultLaw
• ### 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).
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
• ### Method Detail

• #### 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)
• #### initialIsOsculating

public boolean initialIsOsculating()
Check if the initial state is provided in osculating elements.
Returns:
true if initial state is provided in osculating elements

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
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.

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
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 PositionAngle 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
• #### 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,
PositionAngle 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 PositionAngle.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
• #### 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):

  V² r |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