Package org.orekit.propagation.analytical.gnss

Class FieldGnssPropagator<T extends CalculusFieldElement<T>>

java.lang.Object

org.orekit.propagation.FieldAbstractPropagator<T>

org.orekit.propagation.analytical.FieldAbstractAnalyticalPropagator<T>

org.orekit.propagation.analytical.gnss.FieldGnssPropagator<T>

Type Parameters:

T - type of the field elements

All Implemented Interfaces:

FieldPropagator<T>, FieldPVCoordinatesProvider<T>, ParameterDriversProvider

public class FieldGnssPropagator<T extends CalculusFieldElement<T>>
extends FieldAbstractAnalyticalPropagator<T>

Common handling of FieldAbstractAnalyticalPropagator methods for GNSS propagators.

This class allows to provide easily a subset of FieldAbstractAnalyticalPropagator methods for specific GNSS propagators.

Since:

13.0

Author:

Pascal Parraud, Luc Maisonobe

Field Summary

Fields inherited from interface org.orekit.propagation.FieldPropagator

DEFAULT_MASS

Constructor Summary

Constructors

Constructor	Description
FieldGnssPropagator(FieldGnssOrbitalElements<T,?> orbitalElements, Frame eci, Frame ecef, AttitudeProvider provider, T mass)	Build a new instance.
FieldGnssPropagator(FieldSpacecraftState<T> initialState, FieldGnssOrbitalElements<T,?> nonKeplerianElements, Frame ecef, AttitudeProvider provider, T mass)	Build a new instance from an initial state.

Method Summary

Modifier and Type	Method	Description
Frame	getECEF()	Gets the Earth Centered Earth Fixed frame used to propagate GNSS orbits according to the Interface Control Document.
Frame	getECI()	Gets the Earth Centered Inertial frame used to propagate the orbit.
Frame	getFrame()	Get the frame in which the orbit is propagated.
protected T	getMass(FieldAbsoluteDate<T> date)	Get the mass.
T	getMU()	Gets the Earth gravity coefficient used for GNSS propagation.
List<ParameterDriver>	getParametersDrivers()	Get the drivers for parameters.
FieldPVCoordinates<T>	propagateInEcef(FieldAbsoluteDate<T> date, T[] parameters)	Gets the PVCoordinates of the GNSS SV in ECEF frame.
FieldOrbit<T>	propagateOrbit(FieldAbsoluteDate<T> date, T[] parameters)	Propagate an orbit up to a specific target date.
void	resetInitialState(FieldSpacecraftState<T> state)	Reset the propagator initial state.
protected void	resetIntermediateState(FieldSpacecraftState<T> state, boolean forward)	Reset an intermediate state.

Methods inherited from class org.orekit.propagation.analytical.FieldAbstractAnalyticalPropagator

acceptStep, addEventDetector, basicPropagate, clearEventsDetectors, getEphemerisGenerator, getEventDetectors, propagate

Methods inherited from class org.orekit.propagation.FieldAbstractPropagator

addAdditionalDataProvider, getAdditionalDataProviders, getAttitudeProvider, getField, getInitialState, getManagedAdditionalData, getMultiplexer, getStartDate, initializeAdditionalData, initializePropagation, isAdditionalDataManaged, propagate, removeAdditionalDataProvider, setAttitudeProvider, setStartDate, stateChanged, updateAdditionalData, updateUnmanagedData

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.FieldPropagator

clearStepHandlers, getPosition, getPVCoordinates, getVelocity, setStepHandler, setStepHandler

Methods inherited from interface org.orekit.utils.ParameterDriversProvider

getNbParametersDriversValue, getParameterDriver, getParameters, getParameters, getParameters, getParameters, getParametersAllValues, getParametersAllValues, isSupported

Constructor Details

FieldGnssPropagator

public FieldGnssPropagator(FieldGnssOrbitalElements<T,?> orbitalElements,
 Frame eci,
 Frame ecef,
 AttitudeProvider provider,
 T mass)

Build a new instance.

Parameters:

orbitalElements - GNSS orbital elements

eci - Earth Centered Inertial frame

ecef - Earth Centered Earth Fixed frame

provider - Attitude provider

mass - Satellite mass (kg)

FieldGnssPropagator

public FieldGnssPropagator(FieldSpacecraftState<T> initialState,
 FieldGnssOrbitalElements<T,?> nonKeplerianElements,
 Frame ecef,
 AttitudeProvider provider,
 T mass)

Build a new instance from an initial state.

The Keplerian elements already present in the nonKeplerianElements argument will be ignored as it is the initialState argument that will be used to build the complete orbital elements of the propagator

Parameters:

initialState - initial state

nonKeplerianElements - non-Keplerian orbital elements (the Keplerian orbital elements will be ignored)

ecef - Earth Centered Earth Fixed frame

provider - attitude provider

mass - spacecraft mass

Method Details

getParametersDrivers

public List<ParameterDriver> getParametersDrivers()

Get the drivers for parameters.

Returns:

drivers for parameters

getECI

public Frame getECI()

Gets the Earth Centered Inertial frame used to propagate the orbit.

Returns:

the ECI frame

getECEF

public Frame getECEF()

Gets the Earth Centered Earth Fixed frame used to propagate GNSS orbits according to the Interface Control Document.

Returns:

the ECEF frame

getMU

public T getMU()

Gets the Earth gravity coefficient used for GNSS propagation.

Returns:

the Earth gravity coefficient.

propagateOrbit

public FieldOrbit<T> propagateOrbit(FieldAbsoluteDate<T> date,
 T[] parameters)

Propagate an orbit up to a specific target date.

Specified by:

propagateOrbit in class FieldAbstractAnalyticalPropagator<T extends CalculusFieldElement<T>>

Parameters:

date - target date for the orbit

parameters - model parameters

Returns:

propagated orbit

propagateInEcef

public FieldPVCoordinates<T> propagateInEcef(FieldAbsoluteDate<T> date,
 T[] parameters)

Gets the PVCoordinates of the GNSS SV in ECEF frame.

The algorithm uses automatic differentiation to compute velocity and acceleration.

Parameters:

date - the computation date

parameters - propagation parameters

Returns:

the GNSS SV PVCoordinates in ECEF frame

getFrame

public Frame getFrame()

Get the frame in which the orbit is propagated.

The propagation frame is the definition frame of the initial state, so this method should be called after this state has been set, otherwise it may return null.

Specified by:

getFrame in interface FieldPropagator<T extends CalculusFieldElement<T>>

Overrides:

getFrame in class FieldAbstractPropagator<T extends CalculusFieldElement<T>>

Returns:

frame in which the orbit is propagated

See Also:

• FieldPropagator.resetInitialState(FieldSpacecraftState)

getMass

protected T getMass(FieldAbsoluteDate<T> date)

Get the mass.

Specified by:

getMass in class FieldAbstractAnalyticalPropagator<T extends CalculusFieldElement<T>>

Parameters:

date - target date for the orbit

Returns:

mass mass

resetInitialState

public void resetInitialState(FieldSpacecraftState<T> state)

Reset the propagator initial state.

Specified by:

resetInitialState in interface FieldPropagator<T extends CalculusFieldElement<T>>

Overrides:

resetInitialState in class FieldAbstractPropagator<T extends CalculusFieldElement<T>>

Parameters:

state - new initial state to consider

resetIntermediateState

protected void resetIntermediateState(FieldSpacecraftState<T> state,
 boolean forward)

Reset an intermediate state.

Specified by:

resetIntermediateState in class FieldAbstractAnalyticalPropagator<T extends CalculusFieldElement<T>>

Parameters:

state - new intermediate state to consider

forward - if true, the intermediate state is valid for propagations after itself