Package org.orekit.propagation.analytical

Class BrouwerLyddanePropagator

java.lang.Object
org.orekit.propagation.AbstractPropagator
org.orekit.propagation.analytical.AbstractAnalyticalPropagator
org.orekit.propagation.analytical.BrouwerLyddanePropagator
All Implemented Interfaces:
Propagator, ParameterDriversProvider, PVCoordinatesProvider
public class BrouwerLyddanePropagator extends AbstractAnalyticalPropagator implements ParameterDriversProvider
This class propagates a SpacecraftState using the analytical Brouwer-Lyddane model (from J2 to J5 zonal harmonics).

At the opposite of the EcksteinHechlerPropagator, the Brouwer-Lyddane model is suited for elliptical orbits, there is no problem having a rather small eccentricity or inclination (Lyddane helped to solve this issue with the Brouwer model). Singularity for the critical inclination i = 63.4° is avoided using the method developed in Warren Phipps' 1992 thesis.

By default, Brouwer-Lyddane model considers only the perturbations due to zonal harmonics. However, for low Earth orbits, the magnitude of the perturbative acceleration due to atmospheric drag can be significant. Warren Phipps' 1992 thesis considered the atmospheric drag by time derivatives of the mean mean anomaly using the catch-all coefficient M2Driver. Beware that M2Driver must have only 1 span on its TimeSpanMap value. Usually, M2 is adjusted during an orbit determination process and it represents the combination of all unmodeled secular along-track effects (i.e. not just the atmospheric drag). The behavior of M2 is close to the TLE.getBStar() parameter for the TLE. If the value of M2 is equal to 0.0, the along-track secular effects are not considered in the dynamical model. Typical values for M2 are not known. It depends on the orbit type. However, the value of M2 must be very small (e.g. between 1.0e-14 and 1.0e-15). The unit of M2 is rad/s². The along-track effects, represented by the secular rates of the mean semi-major axis and eccentricity, are computed following Eq. 2.38, 2.41, and 2.45 of Warren Phipps' thesis.

Since:
11.1
Author:
Melina Vanel, Bryan Cazabonne, Pascal Parraud
See Also:
  • "Brouwer, Dirk. Solution of the problem of artificial satellite theory without drag. YALE UNIV NEW HAVEN CT NEW HAVEN United States, 1959."
  • "Lyddane, R. H. Small eccentricities or inclinations in the Brouwer theory of the artificial satellite. The Astronomical Journal 68 (1963): 555."
  • "Phipps Jr, Warren E. Parallelization of the Navy Space Surveillance Center (NAVSPASUR) Satellite Model. NAVAL POSTGRADUATE SCHOOL MONTEREY CA, 1992."
  • "Solomon, Daniel, THE NAVSPASUR Satellite Motion Model, Naval Research Laboratory, August 8, 1991."

  • Field Details

    • M2_NAME

      public static final String M2_NAME
      Parameter name for M2 coefficient.
      See Also:

    • M2

      public static final double M2
      Default value for M2 coefficient.
      See Also:

    • EPSILON_DEFAULT

      public static final double EPSILON_DEFAULT
      Default convergence threshold for mean parameters conversion.
      See Also:

    • MAX_ITERATIONS_DEFAULT

      public static final int MAX_ITERATIONS_DEFAULT
      Default value for maxIterations.
      See Also:

  • Constructor Details

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, UnnormalizedSphericalHarmonicsProvider provider, double m2Value)
      Build a propagator from orbit and potential provider.

      Mass and attitude provider are set to unspecified non-null arbitrary values.

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      provider - for un-normalized zonal coefficients
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      See Also:

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double m2Value)
      Build a propagator from orbit and potential.

      Mass and attitude provider are set to unspecified non-null arbitrary values.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      See Also:

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, double mass, UnnormalizedSphericalHarmonicsProvider provider, double m2Value)
      Build a propagator from orbit, mass and potential provider.

      Attitude law is set to an unspecified non-null arbitrary value.

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      mass - spacecraft mass
      provider - for un-normalized zonal coefficients
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      See Also:

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double m2Value)
      Build a propagator from orbit, mass and potential.

      Attitude law is set to an unspecified non-null arbitrary value.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      mass - spacecraft mass
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      See Also:

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, UnnormalizedSphericalHarmonicsProvider provider, double m2Value)
      Build a propagator from orbit, attitude provider and potential provider.

      Mass is set to an unspecified non-null arbitrary value.

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      provider - for un-normalized zonal coefficients
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double m2Value)
      Build a propagator from orbit, attitude provider and potential.

      Mass is set to an unspecified non-null arbitrary value.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, UnnormalizedSphericalHarmonicsProvider provider, double m2Value)
      Build a propagator from orbit, attitude provider, mass and potential provider.

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      mass - spacecraft mass
      provider - for un-normalized zonal coefficients
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      See Also:

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double m2Value)
      Build a propagator from orbit, attitude provider, mass and potential.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      mass - spacecraft mass
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      See Also:

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, UnnormalizedSphericalHarmonicsProvider provider, PropagationType initialType, double m2Value)
      Build a propagator from orbit and potential provider.

      Mass and attitude provider are set to unspecified non-null arbitrary values.

      Using this constructor, it is possible to define the initial orbit as a mean Brouwer-Lyddane orbit or an osculating one.

      Parameters:
      initialOrbit - initial orbit
      provider - for un-normalized zonal coefficients
      initialType - initial orbit type (mean Brouwer-Lyddane orbit or osculating orbit)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, UnnormalizedSphericalHarmonicsProvider provider, PropagationType initialType, double m2Value)
      Build a propagator from orbit, attitude provider, mass and potential provider.

      Using this constructor, it is possible to define the initial orbit as a mean Brouwer-Lyddane orbit or an osculating one.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      mass - spacecraft mass
      provider - for un-normalized zonal coefficients
      initialType - initial orbit type (mean Brouwer-Lyddane orbit or osculating orbit)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, PropagationType initialType, double m2Value)
      Build a propagator from orbit, attitude provider, mass and potential.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, it is possible to define the initial orbit as a mean Brouwer-Lyddane orbit or an osculating one.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      mass - spacecraft mass
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      initialType - initial orbit type (mean Brouwer-Lyddane orbit or osculating orbit)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, PropagationType initialType, double m2Value, double epsilon, int maxIterations)
      Build a propagator from orbit, attitude provider, mass and potential.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, it is possible to define the initial orbit as a mean Brouwer-Lyddane orbit or an osculating one.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      mass - spacecraft mass
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      initialType - initial orbit type (mean Brouwer-Lyddane orbit or osculating orbit)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Since:
      11.2

    • BrouwerLyddanePropagator

      public BrouwerLyddanePropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, PropagationType initialType, double m2Value, OsculatingToMeanConverter converter)
      Build a propagator from orbit, attitude provider, mass and potential.

      The Cn,0 coefficients are the denormalized zonal coefficients, they are related to both the normalized coefficients Cn,0 and the Jn one as follows:

      Cn,0 = [(2-δ0,m)(2n+1)(n-m)!/(n+m)!]½ Cn,0

      Cn,0 = -Jn

      Using this constructor, it is possible to define the initial orbit as a mean Brouwer-Lyddane orbit or an osculating one.

      Parameters:
      initialOrbit - initial orbit
      attitudeProv - attitude provider
      mass - spacecraft mass
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      initialType - initial orbit type (mean Brouwer-Lyddane orbit or osculating orbit)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not computed
      converter - osculating to mean orbit converter
      Since:
      13.0

  • Method Details

    • computeMeanOrbit

      public static KeplerianOrbit computeMeanOrbit(Orbit osculating, UnnormalizedSphericalHarmonicsProvider provider, UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics harmonics, double m2Value)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Brouwer-Lyddane sense, corresponding to the osculating SpacecraftState in input.

      Since the osculating orbit is obtained with the computation of short-periodic variation, the resulting output will depend on both the gravity field parameterized in input and the atmospheric drag represented by the m2Value parameter.

      The computation is done through a fixed-point iteration process.

      Parameters:
      osculating - osculating orbit to convert
      provider - for un-normalized zonal coefficients
      harmonics - provider.onDate(osculating.getDate())
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not considered
      Returns:
      mean orbit in a Brouwer-Lyddane sense
      Since:
      11.2

    • computeMeanOrbit

      public static KeplerianOrbit computeMeanOrbit(Orbit osculating, UnnormalizedSphericalHarmonicsProvider provider, UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics harmonics, double m2Value, double epsilon, int maxIterations)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Brouwer-Lyddane sense, corresponding to the osculating SpacecraftState in input.

      Since the osculating orbit is obtained with the computation of short-periodic variation, the resulting output will depend on both the gravity field parameterized in input and the atmospheric drag represented by the m2Value parameter.

      The computation is done through a fixed-point iteration process.

      Parameters:
      osculating - osculating orbit to convert
      provider - for un-normalized zonal coefficients
      harmonics - provider.onDate(osculating.getDate())
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not considered
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Returns:
      mean orbit in a Brouwer-Lyddane sense
      Since:
      11.2

    • computeMeanOrbit

      public static KeplerianOrbit computeMeanOrbit(Orbit osculating, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double m2Value, double epsilon, int maxIterations)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Brouwer-Lyddane sense, corresponding to the osculating SpacecraftState in input.

      Since the osculating orbit is obtained with the computation of short-periodic variation, the resulting output will depend on both the gravity field parameterized in input and the atmospheric drag represented by the m2Value parameter.

      The computation is done through a fixed-point iteration process.

      Parameters:
      osculating - osculating orbit to convert
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not considered
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Returns:
      mean orbit in a Brouwer-Lyddane sense
      Since:
      11.2

    • computeMeanOrbit

      public static KeplerianOrbit computeMeanOrbit(Orbit osculating, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double m2Value, OsculatingToMeanConverter converter)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Brouwer-Lyddane sense, corresponding to the osculating SpacecraftState in input.

      Since the osculating orbit is obtained with the computation of short-periodic variation, the resulting output will depend on both the gravity field parameterized in input and the atmospheric drag represented by the m2Value parameter.

      The computation is done through the given osculating to mean orbit converter.

      Parameters:
      osculating - osculating orbit to convert
      referenceRadius - reference radius of the Earth for the potential model (m)
      mu - central attraction coefficient (m³/s²)
      c20 - un-normalized zonal coefficient (about -1.08e-3 for Earth)
      c30 - un-normalized zonal coefficient (about +2.53e-6 for Earth)
      c40 - un-normalized zonal coefficient (about +1.62e-6 for Earth)
      c50 - un-normalized zonal coefficient (about +2.28e-7 for Earth)
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not considered
      converter - osculating to mean orbit converter
      Returns:
      mean orbit in a Brouwer-Lyddane sense
      Since:
      13.0

    • computeMeanOrbit

      public static KeplerianOrbit computeMeanOrbit(Orbit osculating, UnnormalizedSphericalHarmonicsProvider provider, double m2Value, OsculatingToMeanConverter converter)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Brouwer-Lyddane sense, corresponding to the osculating SpacecraftState in input.

      Since the osculating orbit is obtained with the computation of short-periodic variation, the resulting output will depend on both the gravity field parameterized in input and the atmospheric drag represented by the m2Value parameter.

      The computation is done through the given osculating to mean orbit converter.

      Parameters:
      osculating - osculating orbit to convert
      provider - for un-normalized zonal coefficients
      m2Value - value of empirical drag coefficient in rad/s². If equal to M2 drag is not considered
      converter - osculating to mean orbit converter
      Returns:
      mean orbit in a Brouwer-Lyddane sense
      Since:
      13.0

    • resetInitialState

      public void resetInitialState(SpacecraftState state)
      Reset the propagator initial state.

      The new initial state to consider must be defined with an osculating orbit.

      Specified by:
      resetInitialState in interface Propagator
      Overrides:
      resetInitialState in class AbstractPropagator
      Parameters:
      state - new initial state to consider
      See Also:

    • resetInitialState

      public void resetInitialState(SpacecraftState state, PropagationType stateType)
      Reset the propagator initial state.
      Parameters:
      state - new initial state to consider
      stateType - mean Brouwer-Lyddane orbit or osculating orbit

    • resetInitialState

      public void resetInitialState(SpacecraftState state, PropagationType stateType, double epsilon, int maxIterations)
      Reset the propagator initial state.
      Parameters:
      state - new initial state to consider
      stateType - mean Brouwer-Lyddane orbit or osculating orbit
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Since:
      11.2

    • resetInitialState

      public void resetInitialState(SpacecraftState state, PropagationType stateType, OsculatingToMeanConverter converter)
      Reset the propagator initial state.
      Parameters:
      state - new initial state to consider
      stateType - mean Brouwer-Lyddane orbit or osculating orbit
      converter - osculating to mean orbit converter
      Since:
      13.0

    • resetIntermediateState

      protected void resetIntermediateState(SpacecraftState state, boolean forward)
      Reset an intermediate state.
      Specified by:
      resetIntermediateState in class AbstractAnalyticalPropagator
      Parameters:
      state - new intermediate state to consider
      forward - if true, the intermediate state is valid for propagations after itself

    • resetIntermediateState

      protected void resetIntermediateState(SpacecraftState state, boolean forward, double epsilon, int maxIterations)
      Reset an intermediate state.
      Parameters:
      state - new intermediate state to consider
      forward - if true, the intermediate state is valid for propagations after itself
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Since:
      11.2

    • resetIntermediateState

      protected void resetIntermediateState(SpacecraftState state, boolean forward, OsculatingToMeanConverter converter)
      Reset an intermediate state.
      Parameters:
      state - new intermediate state to consider
      forward - if true, the intermediate state is valid for propagations after itself
      converter - osculating to mean orbit converter
      Since:
      13.0

    • propagateOrbit

      public KeplerianOrbit propagateOrbit(AbsoluteDate date)
      Extrapolate an orbit up to a specific target date.
      Specified by:
      propagateOrbit in class AbstractAnalyticalPropagator
      Parameters:
      date - target date for the orbit
      Returns:
      extrapolated parameters

    • getM2

      public double getM2()
      Get the value of the M2 drag parameter. Beware that M2Driver must have only 1 span on its TimeSpanMap value (that is to say setPeriod method should not be called)
      Returns:
      the value of the M2 drag parameter

    • getMu

      public double getMu()
      Get the central attraction coefficient μ.
      Returns:
      mu central attraction coefficient (m³/s²)

    • getCk0

      public double[] getCk0()
      Get the un-normalized zonal coefficients.
      Returns:
      the un-normalized zonal coefficients

    • getReferenceRadius

      public double getReferenceRadius()
      Get the reference radius of the central body attraction model.
      Returns:
      the reference radius in meters

    • getParametersDrivers

      public List<ParameterDriver> getParametersDrivers()
      Get the parameters driver for propagation model.
      Specified by:
      getParametersDrivers in interface ParameterDriversProvider
      Returns:
      drivers for propagation model

    • createHarvester

      protected AbstractMatricesHarvester createHarvester(String stmName, RealMatrix initialStm, DoubleArrayDictionary initialJacobianColumns)
      Create the harvester suitable for propagator.
      Overrides:
      createHarvester in class AbstractPropagator
      Parameters:
      stmName - State Transition Matrix state name
      initialStm - initial State Transition Matrix ∂Y/∂Y₀, if null (which is the most frequent case), assumed to be 6x6 identity
      initialJacobianColumns - 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

    • getJacobiansColumnsNames

      protected List<String> getJacobiansColumnsNames()
      Get the names of the parameters in the matrix returned by MatricesHarvester.getParametersJacobian(org.orekit.propagation.SpacecraftState).
      Overrides:
      getJacobiansColumnsNames in class AbstractAnalyticalPropagator
      Returns:
      names of the parameters (i.e. columns) of the Jacobian matrix

    • getMass

      protected double getMass(AbsoluteDate date)
      Get the mass.
      Specified by:
      getMass in class AbstractAnalyticalPropagator
      Parameters:
      date - target date for the orbit
      Returns:
      mass mass