org.orekit.propagation.analytical.gnss

Class SBASPropagator

java.lang.Object

org.orekit.propagation.AbstractPropagator

org.orekit.propagation.analytical.AbstractAnalyticalPropagator

org.orekit.propagation.analytical.gnss.SBASPropagator

All Implemented Interfaces:

Propagator, PVCoordinatesProvider

public class SBASPropagator
extends AbstractAnalyticalPropagator

This class aims at propagating a SBAS orbit from SBASOrbitalElements.

Since:

10.1

Author:

Bryan Cazabonne

See Also:

"Tyler Reid, Todd Walker, Per Enge, L1/L5 SBAS MOPS Ephemeris Message to Support Multiple Orbit Classes, ION ITM, 2013"

Field Summary

Fields inherited from interface org.orekit.propagation.Propagator

DEFAULT_MASS

Method Summary

Modifier and Type	Method and Description
Frame	getECEF() Gets the Earth Centered Earth Fixed frame used to propagate GNSS orbits.
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 SBAS propagation.
SBASOrbitalElements	getSBASOrbitalElements() Get the underlying SBAS orbital elements.
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, getEphemerisGenerator, getEventsDetectors, getJacobiansColumnsNames, getPvProvider, propagate

Methods inherited from class org.orekit.propagation.AbstractPropagator

addAdditionalStateProvider, createHarvester, getAdditionalStateProviders, getAttitudeProvider, getHarvester, getInitialState, getManagedAdditionalStates, getMultiplexer, getPVCoordinates, getStartDate, initializePropagation, isAdditionalStateManaged, propagate, setAttitudeProvider, setStartDate, setupMatricesComputation, stateChanged, updateAdditionalStates, updateUnmanagedStates

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

clearStepHandlers, getDefaultLaw, setStepHandler, setStepHandler

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 SBAS propagation.

Returns:

the Earth gravity coefficient.

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.

Returns:

the ECEF frame

getSBASOrbitalElements

public SBASOrbitalElements getSBASOrbitalElements()

Get the underlying SBAS orbital elements.

Returns:

the underlying SBAS orbital elements

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)

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

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

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