org.orekit.propagation.analytical.gnss

Class AbstractGNSSPropagator

java.lang.Object

org.orekit.propagation.AbstractPropagator

org.orekit.propagation.analytical.AbstractAnalyticalPropagator

org.orekit.propagation.analytical.gnss.AbstractGNSSPropagator

All Implemented Interfaces:

Propagator, PVCoordinatesProvider

Direct Known Subclasses:

BeidouPropagator, GalileoPropagator, GPSPropagator, IRNSSPropagator, QZSSPropagator

public abstract class AbstractGNSSPropagator
extends AbstractAnalyticalPropagator

Common handling of AbstractAnalyticalPropagator methods for GNSS propagators.

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

Author:

Pascal Parraud

Field Summary

Fields inherited from interface org.orekit.propagation.Propagator

DEFAULT_LAW, DEFAULT_MASS, EPHEMERIS_GENERATION_MODE, MASTER_MODE, SLAVE_MODE

Constructor Summary

Constructors

Modifier	Constructor and Description
protected	AbstractGNSSPropagator(GNSSOrbitalElements gnssOrbit, AttitudeProvider attitudeProvider, Frame eci, Frame ecef, double mass, double av, double cycleDuration, double mu) Build a new instance.

Method Summary

Modifier and Type	Method and 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 double	getMass(AbsoluteDate date) Get the mass.
double	getMU() Get the Earth gravity coefficient used for GNSS propagation.
PVCoordinates	propagateInEcef(AbsoluteDate date) Gets the PVCoordinates of the GNSS SV in ECEF frame.
protected Orbit	propagateOrbit(AbsoluteDate date) Extrapolate an orbit up to a specific target date.
void	resetInitialState(SpacecraftState state) Reset the propagator initial state.
protected void	resetIntermediateState(SpacecraftState state, boolean forward) Reset an intermediate state.

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

acceptStep, addEventDetector, basicPropagate, clearEventsDetectors, getEventsDetectors, getGeneratedEphemeris, getPvProvider, propagate

Methods inherited from class org.orekit.propagation.AbstractPropagator

addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFixedStepSize, getInitialState, getManagedAdditionalStates, getMode, getPVCoordinates, getStartDate, getStepHandler, initializePropagation, isAdditionalStateManaged, propagate, setAttitudeProvider, setEphemerisMode, setEphemerisMode, setMasterMode, setMasterMode, setSlaveMode, setStartDate, stateChanged, updateAdditionalStates

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.Propagator

getDefaultLaw

Constructor Detail

AbstractGNSSPropagator

protected AbstractGNSSPropagator(GNSSOrbitalElements gnssOrbit,
                                 AttitudeProvider attitudeProvider,
                                 Frame eci,
                                 Frame ecef,
                                 double mass,
                                 double av,
                                 double cycleDuration,
                                 double mu)

Build a new instance.

Parameters:

gnssOrbit - the common GNSS orbital elements to be used by the Abstract GNSS propagator

attitudeProvider - provider for attitude computation

eci - the ECI frame used for GNSS propagation

ecef - the ECEF frame used for GNSS propagation

mass - the spacecraft mass (kg)

av - mean angular velocity of the Earth (rad/s)

cycleDuration - duration of the GNSS cycle in seconds

mu - the Earth gravity coefficient used for GNSS propagation

Method Detail

propagateInEcef

public PVCoordinates propagateInEcef(AbsoluteDate date)

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

Returns:

the GNSS SV PVCoordinates in ECEF frame

propagateOrbit

protected Orbit 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

getMU

public double getMU()

Get the Earth gravity coefficient used for GNSS propagation.

Returns:

the Earth gravity coefficient.

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 Propagator

Overrides:

getFrame in class AbstractPropagator

Returns:

frame in which the orbit is propagated

See Also:

Propagator.resetInitialState(SpacecraftState)

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

resetInitialState

public void resetInitialState(SpacecraftState state)

Reset the propagator initial state.

Specified by:

resetInitialState in interface Propagator

Overrides:

resetInitialState in class AbstractPropagator

Parameters:

state - new initial state to consider

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

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