Package org.orekit.propagation.analytical

Class EcksteinHechlerPropagator

java.lang.Object
org.orekit.propagation.AbstractPropagator
org.orekit.propagation.analytical.AbstractAnalyticalPropagator
org.orekit.propagation.analytical.EcksteinHechlerPropagator
All Implemented Interfaces:
Propagator, PVCoordinatesProvider
public class EcksteinHechlerPropagator extends AbstractAnalyticalPropagator
This class propagates a SpacecraftState using the analytical Eckstein-Hechler model.

The Eckstein-Hechler model is suited for near circular orbits (e < 0.1, with poor accuracy between 0.005 and 0.1) and inclination neither equatorial (direct or retrograde) nor critical (direct or retrograde).

Note that before version 7.0, there was a large inconsistency in the generated orbits, and it was fixed as of version 7.0 of Orekit, with a visible side effect. The problems is that if the circular parameters produced by the Eckstein-Hechler model are used to build an orbit considered to be osculating, the velocity deduced from this orbit was inconsistent with the position evolution! The reason is that the model includes non-Keplerian effects but it does not include a corresponding circular/Cartesian conversion. As a consequence, all subsequent computation involving velocity were wrong. This includes attitude modes like yaw compensation and Doppler effect. As this effect was considered serious enough and as accurate velocities were considered important, the propagator now generates Cartesian orbits which are built in a special way to ensure consistency throughout propagation. A side effect is that if circular parameters are rebuilt by user from these propagated Cartesian orbit, the circular parameters will generally not match the initial orbit (differences in semi-major axis can exceed 120 m). The position however will match to sub-micrometer level, and this position will be identical to the positions that were generated by previous versions (in other words, the internals of the models have not been changed, only the output parameters have been changed). The correctness of the initialization has been assessed and is good, as it allows the subsequent orbit to remain close to a numerical reference orbit.

If users need a more definitive initialization of an Eckstein-Hechler propagator, they should consider using a propagator converter to initialize their Eckstein-Hechler propagator using a complete sample instead of just a single initial orbit.

Author:
Guylaine Prat
See Also:

  • Constructor Details

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, UnnormalizedSphericalHarmonicsProvider provider)
      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
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitude, double mass, UnnormalizedSphericalHarmonicsProvider provider, UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics harmonics)
      Private helper constructor.

      Using this constructor, an initial osculating orbit is considered.

      Parameters:
      initialOrbit - initial orbit
      attitude - attitude provider
      mass - spacecraft mass
      provider - for un-normalized zonal coefficients
      harmonics - provider.onDate(initialOrbit.getDate())
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60)
      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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, double mass, UnnormalizedSphericalHarmonicsProvider provider)
      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
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60)
      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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, UnnormalizedSphericalHarmonicsProvider provider)
      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

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60)
      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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, UnnormalizedSphericalHarmonicsProvider provider)
      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
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60)
      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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      See Also:

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, UnnormalizedSphericalHarmonicsProvider provider, PropagationType initialType)
      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 Eckstein-Hechler orbit or an osculating one.

      Parameters:
      initialOrbit - initial orbit
      provider - for un-normalized zonal coefficients
      initialType - initial orbit type (mean Eckstein-Hechler orbit or osculating orbit)
      Since:
      10.2

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, UnnormalizedSphericalHarmonicsProvider provider, PropagationType initialType)
      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 Eckstein-Hechler 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 Eckstein-Hechler orbit or osculating orbit)
      Since:
      10.2

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitude, double mass, UnnormalizedSphericalHarmonicsProvider provider, UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics harmonics, PropagationType initialType)
      Private helper constructor.

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

      Parameters:
      initialOrbit - initial orbit
      attitude - attitude provider
      mass - spacecraft mass
      provider - for un-normalized zonal coefficients
      harmonics - provider.onDate(initialOrbit.getDate())
      initialType - initial orbit type (mean Eckstein-Hechler orbit or osculating orbit)
      Since:
      10.2

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60, PropagationType initialType)
      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 Eckstein-Hechler 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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      initialType - initial orbit type (mean Eckstein-Hechler orbit or osculating orbit)
      Since:
      10.2

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60, PropagationType initialType, 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 Eckstein-Hechler 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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      initialType - initial orbit type (mean Eckstein-Hechler orbit or osculating orbit)
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Since:
      11.2

    • EcksteinHechlerPropagator

      public EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double mass, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60, PropagationType initialType, 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 Eckstein-Hechler 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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      initialType - initial orbit type (mean Eckstein-Hechler orbit or osculating orbit)
      converter - osculating to mean orbit converter
      Since:
      13.0

  • Method Details

    • computeMeanOrbit

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

      Compute mean orbit in a Eckstein-Hechler 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 the gravity field parameterized in input.

      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())
      Returns:
      mean orbit in a Eckstein-Hechler sense
      Since:
      11.2

    • computeMeanOrbit

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

      Compute mean orbit in a Eckstein-Hechler 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 the gravity field parameterized in input.

      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())
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Returns:
      mean orbit in a Eckstein-Hechler sense
      Since:
      11.2

    • computeMeanOrbit

      public static CircularOrbit computeMeanOrbit(Orbit osculating, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60, double epsilon, int maxIterations)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Eckstein-Hechler 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 the gravity field parameterized in input.

      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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      epsilon - convergence threshold for mean parameters conversion
      maxIterations - maximum iterations for mean parameters conversion
      Returns:
      mean orbit in a Eckstein-Hechler sense
      Since:
      11.2

    • computeMeanOrbit

      public static CircularOrbit computeMeanOrbit(Orbit osculating, double referenceRadius, double mu, double c20, double c30, double c40, double c50, double c60, OsculatingToMeanConverter converter)
      Conversion from osculating to mean orbit.

      Compute mean orbit in a Eckstein-Hechler 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 the gravity field parameterized in input.

      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)
      c60 - un-normalized zonal coefficient (about -5.41e-7 for Earth)
      converter - osculating to mean orbit converter
      Returns:
      mean orbit in a Eckstein-Hechler 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 Eckstein-Hechler orbit or osculating orbit
      Since:
      10.2

    • 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 Eckstein-Hechler 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 Eckstein-Hechler 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 CartesianOrbit 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

    • getOsculatingCircularOrbit

      public CircularOrbit getOsculatingCircularOrbit(AbsoluteDate date)
      Get the osculating circular orbit from the EH model.

      This method is only relevant for the conversion from osculating to mean orbit.

      Parameters:
      date - target date for the orbit
      Returns:
      the osculating circular orbite

    • getMu

      public double getMu()
      Get the central attraction coefficient μ.
      Returns:
      mu central attraction coefficient (m³/s²)
      Since:
      11.1

    • getCk0

      public double[] getCk0()
      Get the un-normalized zonal coefficients.
      Returns:
      the un-normalized zonal coefficients
      Since:
      11.1

    • getReferenceRadius

      public double getReferenceRadius()
      Get the reference radius of the central body attraction model.
      Returns:
      the reference radius in meters
      Since:
      11.1

    • 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

    • 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