fr.cnes.sirius.patrius.propagation.analytical

Class EcksteinHechlerPropagator

java.lang.Object

fr.cnes.sirius.patrius.propagation.AbstractPropagator

fr.cnes.sirius.patrius.propagation.analytical.EcksteinHechlerPropagator

All Implemented Interfaces:

PVCoordinatesProvider, MeanOsculatingElementsProvider, Propagator, Serializable

public class EcksteinHechlerPropagator
extends AbstractPropagator
implements MeanOsculatingElementsProvider

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).

Author:

Guylaine Prat

See Also:

Orbit, Serialized Form

Field Summary

Fields inherited from class fr.cnes.sirius.patrius.propagation.AbstractPropagator

MASS

Fields inherited from interface fr.cnes.sirius.patrius.propagation.Propagator

EPHEMERIS_GENERATION_MODE, MASTER_MODE, SLAVE_MODE

Constructor Summary

Constructors

Constructor and Description
EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents, double referenceRadiusIn, double muIn, Frame frame, double c20In, double c30In, double c40In, double c50In, double c60In, MassProvider massProvider, ParametersType paramsType) Build a propagator from orbit, attitude provider, mass and potential.
EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents, double referenceRadiusIn, double muIn, Frame frame, double c20In, double c30In, double c40In, double c50In, double c60In, ParametersType paramsType) Build a propagator from orbit, attitude provider for forces and events computation and potential.
EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double referenceRadiusIn, double muIn, Frame frame, double c20In, double c30In, double c40In, double c50In, double c60In, MassProvider massProvider, ParametersType paramsType) Build a propagator from orbit, attitude provider, mass and potential.
EcksteinHechlerPropagator(Orbit initialOrbit, AttitudeProvider attitudeProv, double referenceRadiusIn, double muIn, Frame frame, double c20In, double c30In, double c40In, double c50In, double c60In, ParametersType paramsType) Build a propagator from orbit, attitude provider and potential.
EcksteinHechlerPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, Frame frame, double c20In, double c30In, double c40In, double c50In, double c60In, MassProvider massProvider, ParametersType paramsType) Build a propagator from orbit, mass and potential.
EcksteinHechlerPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, Frame frame, double c20In, double c30In, double c40In, double c50In, double c60In, ParametersType paramsType) Build a propagator from orbit and potential.

Method Summary

Modifier and Type	Method and Description
Orbit	mean2osc(Orbit orbit) Convert provided mean orbit into osculating elements.
Orbit	osc2mean(Orbit orbit) Convert provided osculating orbit into mean elements.
Orbit	propagateMeanOrbit(AbsoluteDate date) Propagate mean orbit until provided date.
Orbit	propagateOrbit(AbsoluteDate date) Extrapolate an orbit up to a specific target date.
void	resetInitialState(SpacecraftState state) Reset the propagator initial state.
static void	setThreshold(double newThreshold) Setter for osculating to mean conversion relative convergence threshold.

Methods inherited from class fr.cnes.sirius.patrius.propagation.AbstractPropagator

acceptStep, addAdditionalStateProvider, addAdditionalStateProvider, addEventDetector, basicPropagate, clearEventsDetectors, getAttitudeProvider, getAttitudeProviderEvents, getAttitudeProviderForces, getEventsDetectors, getFrame, getGeneratedEphemeris, getInitialState, getMode, getPVCoordinates, getPvProvider, manageStateFrame, propagate, propagate, setAttitudeProvider, setAttitudeProviderEvents, setAttitudeProviderForces, setEphemerisMode, setMasterMode, setMasterMode, setOrbitFrame, setSlaveMode, setStartDate

Methods inherited from class java.lang.Object

clone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, wait

Constructor Detail

EcksteinHechlerPropagator

public EcksteinHechlerPropagator(Orbit initialOrbit,
                                 double referenceRadiusIn,
                                 double muIn,
                                 Frame frame,
                                 double c20In,
                                 double c30In,
                                 double c40In,
                                 double c50In,
                                 double c60In,
                                 ParametersType paramsType)
                          throws PropagationException

Build a propagator from orbit and potential.

Mass and attitude provider are set to null value.

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

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

Parameters:

initialOrbit - initial orbit

referenceRadiusIn - reference radius of the Earth for the potential model (m)

muIn - central attraction coefficient (m³/s²) used for propagation

frame - Inertial or quasi-inertial frame in which the model is supposed valid, the Z axis of the frame being the polar axis of the body

c20In - un-normalized zonal coefficient (about -1.08e-3 for Earth)

c30In - un-normalized zonal coefficient (about +2.53e-6 for Earth)

c40In - un-normalized zonal coefficient (about +1.62e-6 for Earth)

c50In - un-normalized zonal coefficient (about +2.28e-7 for Earth)

c60In - un-normalized zonal coefficient (about -5.41e-7 for Earth)

paramsType - parameters type (mean or osculating)

Throws:

PropagationException - if the mean parameters cannot be computed or coefficients frame is not inertial

See Also:

Constants, ParametersType

EcksteinHechlerPropagator

public EcksteinHechlerPropagator(Orbit initialOrbit,
                                 double referenceRadiusIn,
                                 double muIn,
                                 Frame frame,
                                 double c20In,
                                 double c30In,
                                 double c40In,
                                 double c50In,
                                 double c60In,
                                 MassProvider massProvider,
                                 ParametersType paramsType)
                          throws PropagationException

Build a propagator from orbit, mass and potential.

Attitude law is set to null value.

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

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

Parameters:

initialOrbit - initial orbit

referenceRadiusIn - reference radius of the Earth for the potential model (m)

muIn - central attraction coefficient (m³/s²) used for propagation

frame - Inertial or quasi-inertial frame in which the model is supposed valid, the Z axis of the frame being the polar axis of the body

c20In - un-normalized zonal coefficient (about -1.08e-3 for Earth)

c30In - un-normalized zonal coefficient (about +2.53e-6 for Earth)

c40In - un-normalized zonal coefficient (about +1.62e-6 for Earth)

c50In - un-normalized zonal coefficient (about +2.28e-7 for Earth)

c60In - un-normalized zonal coefficient (about -5.41e-7 for Earth)

massProvider - spacecraft mass provider

paramsType - parameters type (mean or osculating)

Throws:

PropagationException - if the mean parameters cannot be computed or coefficients frame is not inertial

EcksteinHechlerPropagator

public EcksteinHechlerPropagator(Orbit initialOrbit,
                                 AttitudeProvider attitudeProv,
                                 double referenceRadiusIn,
                                 double muIn,
                                 Frame frame,
                                 double c20In,
                                 double c30In,
                                 double c40In,
                                 double c50In,
                                 double c60In,
                                 ParametersType paramsType)
                          throws PropagationException

Build a propagator from orbit, attitude provider and potential.

Mass is set to null value.

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

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

Parameters:

initialOrbit - initial orbit

attitudeProv - attitude provider

referenceRadiusIn - reference radius of the Earth for the potential model (m)

muIn - central attraction coefficient (m³/s²) used for propagation

frame - Inertial or quasi-inertial frame in which the model is supposed valid, the Z axis of the frame being the polar axis of the body

c20In - un-normalized zonal coefficient (about -1.08e-3 for Earth)

c30In - un-normalized zonal coefficient (about +2.53e-6 for Earth)

c40In - un-normalized zonal coefficient (about +1.62e-6 for Earth)

c50In - un-normalized zonal coefficient (about +2.28e-7 for Earth)

c60In - un-normalized zonal coefficient (about -5.41e-7 for Earth)

paramsType - parameters type (mean or osculating)

Throws:

PropagationException - if the mean parameters cannot be computed or coefficients frame is not inertial

EcksteinHechlerPropagator

public EcksteinHechlerPropagator(Orbit initialOrbit,
                                 AttitudeProvider attitudeProvForces,
                                 AttitudeProvider attitudeProvEvents,
                                 double referenceRadiusIn,
                                 double muIn,
                                 Frame frame,
                                 double c20In,
                                 double c30In,
                                 double c40In,
                                 double c50In,
                                 double c60In,
                                 ParametersType paramsType)
                          throws PropagationException

Build a propagator from orbit, attitude provider for forces and events computation and potential.

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

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

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

Parameters:

initialOrbit - initial orbit

attitudeProvForces - attitude provider for forces computation

attitudeProvEvents - attitude provider for events computation

referenceRadiusIn - reference radius of the Earth for the potential model (m)

muIn - central attraction coefficient (m³/s²) used for propagation

frame - Inertial or quasi-inertial frame in which the model is supposed valid, the Z axis of the frame being the polar axis of the body

c20In - un-normalized zonal coefficient (about -1.08e-3 for Earth)

c30In - un-normalized zonal coefficient (about +2.53e-6 for Earth)

c40In - un-normalized zonal coefficient (about +1.62e-6 for Earth)

c50In - un-normalized zonal coefficient (about +2.28e-7 for Earth)

c60In - un-normalized zonal coefficient (about -5.41e-7 for Earth)

paramsType - parameters type (mean or osculating)

Throws:

PropagationException - if the mean parameters cannot be computed or coefficients frame is not inertial

EcksteinHechlerPropagator

public EcksteinHechlerPropagator(Orbit initialOrbit,
                                 AttitudeProvider attitudeProv,
                                 double referenceRadiusIn,
                                 double muIn,
                                 Frame frame,
                                 double c20In,
                                 double c30In,
                                 double c40In,
                                 double c50In,
                                 double c60In,
                                 MassProvider massProvider,
                                 ParametersType paramsType)
                          throws PropagationException

Build a propagator from orbit, attitude provider, mass and potential.

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

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

Parameters:

initialOrbit - initial orbit

attitudeProv - attitude provider

referenceRadiusIn - reference radius of the Earth for the potential model (m)

muIn - central attraction coefficient (m³/s²) used for propagation

frame - Inertial or quasi-inertial frame in which the model is supposed valid, the Z axis of the frame being the polar axis of the body

c20In - un-normalized zonal coefficient (about -1.08e-3 for Earth)

c30In - un-normalized zonal coefficient (about +2.53e-6 for Earth)

c40In - un-normalized zonal coefficient (about +1.62e-6 for Earth)

c50In - un-normalized zonal coefficient (about +2.28e-7 for Earth)

c60In - un-normalized zonal coefficient (about -5.41e-7 for Earth)

massProvider - spacecraft mass provider

paramsType - parameters type (mean or osculating)

Throws:

PropagationException - if the mean parameters cannot be computed or coefficients frame is not inertial

EcksteinHechlerPropagator

public EcksteinHechlerPropagator(Orbit initialOrbit,
                                 AttitudeProvider attitudeProvForces,
                                 AttitudeProvider attitudeProvEvents,
                                 double referenceRadiusIn,
                                 double muIn,
                                 Frame frame,
                                 double c20In,
                                 double c30In,
                                 double c40In,
                                 double c50In,
                                 double c60In,
                                 MassProvider massProvider,
                                 ParametersType paramsType)
                          throws PropagationException

Build a propagator from orbit, attitude provider, mass and potential.

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

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

Parameters:

initialOrbit - initial orbit

attitudeProvForces - attitude provider for forces computation

attitudeProvEvents - attitude provider for events computation

referenceRadiusIn - reference radius of the Earth for the potential model (m)

muIn - central attraction coefficient (m³/s²) used for propagation

frame - Inertial or quasi-inertial frame in which the model is supposed valid, the Z axis of the frame being the polar axis of the body

c20In - un-normalized zonal coefficient (about -1.08e-3 for Earth)

c30In - un-normalized zonal coefficient (about +2.53e-6 for Earth)

c40In - un-normalized zonal coefficient (about +1.62e-6 for Earth)

c50In - un-normalized zonal coefficient (about +2.28e-7 for Earth)

c60In - un-normalized zonal coefficient (about -5.41e-7 for Earth)

massProvider - spacecraft mass provider

paramsType - parameters type

Throws:

PropagationException - if the mean parameters cannot be computed or coefficients frame is not inertial

Method Detail

resetInitialState

public void resetInitialState(SpacecraftState state)
                       throws PropagationException

Reset the propagator initial state.

Specified by:

resetInitialState in interface Propagator

Overrides:

resetInitialState in class AbstractPropagator

Parameters:

state - new initial state to consider

Throws:

PropagationException - if initial state cannot be reset

propagateOrbit

public Orbit propagateOrbit(AbsoluteDate date)
                     throws PropagationException

Extrapolate an orbit up to a specific target date.

Specified by:

propagateOrbit in class AbstractPropagator

Parameters:

date - target date for the orbit

Returns:

extrapolated parameters

Throws:

PropagationException - if some parameters are out of bounds

propagateMeanOrbit

public Orbit propagateMeanOrbit(AbsoluteDate date)
                         throws PatriusException

Propagate mean orbit until provided date.

Specified by:

propagateMeanOrbit in interface MeanOsculatingElementsProvider

Parameters:

date - a date

Returns:

mean orbit at provided date

Throws:

PatriusException - thrown if computation failed

osc2mean

public Orbit osc2mean(Orbit orbit)
               throws PatriusException

Convert provided osculating orbit into mean elements.

Warning: Used algorithm often consists in an iterative algorithm with a convergence criterion. As a result convergence is not always ensured, depending on the underlying theory.

Specified by:

osc2mean in interface MeanOsculatingElementsProvider

Parameters:

orbit - an orbit (osculating elements)

Returns:

mean elements of provided orbit

Throws:

PatriusException - if conversion fails

mean2osc

public Orbit mean2osc(Orbit orbit)
               throws PatriusException

Convert provided mean orbit into osculating elements.

Specified by:

mean2osc in interface MeanOsculatingElementsProvider

Parameters:

orbit - an orbit (mean elements)

Returns:

osculating elements of provided orbit

Throws:

PatriusException - if conversion fails

setThreshold

public static void setThreshold(double newThreshold)

Setter for osculating to mean conversion relative convergence threshold. Default value for this threshold is 1E-13.

Parameters:

newThreshold - new threshold to set