Package org.orekit.propagation.semianalytical.dsst

Class DSSTPropagator

    • Constructor Detail

      • DSSTPropagator

        public DSSTPropagator​(org.hipparchus.ode.ODEIntegrator integrator,
                      boolean meanOnly)
        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.
        meanOnly - output only the mean orbits.

      • DSSTPropagator

        public DSSTPropagator​(org.hipparchus.ode.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.

        Parameters:
        integrator - numerical integrator to use for propagation.

    • 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,
                            boolean isOsculating)
        Set the initial state.
        Parameters:
        initialState - initial state
        isOsculating - true if the orbital state is defined with osculating elements

      • 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

      • 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()

      • removeForceModels

        public void removeForceModels()
        Remove all perturbing force models from the global perturbation model.

        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)

      • 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

      • 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

      • 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

      • 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

      • 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