org.orekit.propagation.numerical

Class GLONASSNumericalPropagator

java.lang.Object

org.orekit.propagation.AbstractPropagator

org.orekit.propagation.integration.AbstractIntegratedPropagator

org.orekit.propagation.numerical.GLONASSNumericalPropagator

All Implemented Interfaces:

Propagator, PVCoordinatesProvider

public class GLONASSNumericalPropagator
extends AbstractIntegratedPropagator

This class propagates GLONASS orbits using numerical integration.

As recommended by the GLONASS Interface Control Document (ICD), a 4th order Runge-Kutta technique shall be used to integrate the equations.

Classical used of this orbit propagator is to compute GLONASS satellite coordinates from the navigation message.

If the projections of luni-solar accelerations to axes of Greenwich geocentric coordinates X''(tb), Y''(tb) and Z''(tb) are available in the navigation message; a transformation is performed to convert these accelerations into the correct coordinate system. In the case where they are not available into the navigation message, these accelerations are computed.

Author:

Bryan Cazabonne

See Also:

GLONASS Interface Control Document

Nested Class Summary

Nested classes/interfaces inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator

AbstractIntegratedPropagator.MainStateEquations

Field Summary

Fields inherited from interface org.orekit.propagation.Propagator

DEFAULT_MASS

Constructor Summary

Constructors

Constructor and Description
GLONASSNumericalPropagator(ClassicalRungeKuttaIntegrator integrator, GLONASSOrbitalElements glonassOrbit, Frame eci, AttitudeProvider provider, double mass, DataContext context, boolean isAccAvailable) Private constructor.

Method Summary

Modifier and Type	Method and Description
protected StateMapper	createMapper(AbsoluteDate referenceDate, double mu, OrbitType orbitType, PositionAngle positionAngleType, AttitudeProvider attitudeProvider, Frame frame) Create a mapper between raw double components and spacecraft state.
GLONASSOrbitalElements	getGLONASSOrbitalElements() Gets the underlying GLONASS orbital elements.
protected AbstractIntegratedPropagator.MainStateEquations	getMainStateEquations(ODEIntegrator integ) Get the differential equations to integrate (for main state only).
SpacecraftState	propagate(AbsoluteDate date) Propagate towards a target date.

Methods inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator

addAdditionalDerivativesProvider, addAdditionalEquations, addEventDetector, afterIntegration, beforeIntegration, clearEventsDetectors, getAdditionalDerivativesProviders, getBasicDimension, getCalls, getEphemerisGenerator, getEventsDetectors, getInitialIntegrationState, getIntegrator, getManagedAdditionalStates, getMu, getOrbitType, getPositionAngleType, getPropagationType, initMapper, isAdditionalStateManaged, isMeanOrbit, propagate, setAttitudeProvider, setMu, setOrbitType, setPositionAngleType, setResetAtEnd, setUpEventDetector, setUpStmAndJacobianGenerators, setUpUserEventDetectors

Methods inherited from class org.orekit.propagation.AbstractPropagator

addAdditionalStateProvider, createHarvester, getAdditionalStateProviders, getAttitudeProvider, getFrame, getHarvester, getInitialState, getMultiplexer, getPVCoordinates, getStartDate, initializePropagation, resetInitialState, 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

Constructor Detail

GLONASSNumericalPropagator

public GLONASSNumericalPropagator(ClassicalRungeKuttaIntegrator integrator,
                                  GLONASSOrbitalElements glonassOrbit,
                                  Frame eci,
                                  AttitudeProvider provider,
                                  double mass,
                                  DataContext context,
                                  boolean isAccAvailable)

Private constructor.

Parameters:

integrator - Runge-Kutta integrator

glonassOrbit - Glonass orbital elements

eci - Earth Centered Inertial frame

provider - Attitude provider

mass - Satellite mass (kg)

context - Data context

isAccAvailable - true if the acceleration is transmitted within the navigation message

Method Detail

getGLONASSOrbitalElements

public GLONASSOrbitalElements getGLONASSOrbitalElements()

Gets the underlying GLONASS orbital elements.

Returns:

the underlying GLONASS orbital elements

propagate

public SpacecraftState propagate(AbsoluteDate date)

Propagate towards a target date.

Simple propagators use only the target date as the specification for computing the propagated state. More feature rich propagators can consider other information and provide different operating modes or G-stop facilities to stop at pinpointed events occurrences. In these cases, the target date is only a hint, not a mandatory objective.

Specified by:

propagate in interface Propagator

Overrides:

propagate in class AbstractIntegratedPropagator

Parameters:

date - target date towards which orbit state should be propagated

Returns:

propagated state

createMapper

protected StateMapper createMapper(AbsoluteDate referenceDate,
                                   double mu,
                                   OrbitType orbitType,
                                   PositionAngle positionAngleType,
                                   AttitudeProvider attitudeProvider,
                                   Frame frame)

Description copied from class: AbstractIntegratedPropagator

Create a mapper between raw double components and spacecraft state. /** Simple constructor.

The position parameter type is meaningful only if propagation orbit type support it. As an example, it is not meaningful for propagation in Cartesian parameters.

Specified by:

createMapper in class AbstractIntegratedPropagator

Parameters:

referenceDate - reference date

mu - central attraction coefficient (m³/s²)

orbitType - orbit type to use for mapping

positionAngleType - angle type to use for propagation

attitudeProvider - attitude provider

frame - inertial frame

Returns:

new mapper

getMainStateEquations

protected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(ODEIntegrator integ)

Description copied from class: AbstractIntegratedPropagator

Get the differential equations to integrate (for main state only).

Specified by:

getMainStateEquations in class AbstractIntegratedPropagator

Parameters:

integ - numerical integrator to use for propagation.

Returns:

differential equations for main state