Package org.orekit.propagation.analytical.intelsat

Class IntelsatElevenElementsPropagator

java.lang.Object

org.orekit.propagation.AbstractPropagator

org.orekit.propagation.analytical.AbstractAnalyticalPropagator

org.orekit.propagation.analytical.intelsat.IntelsatElevenElementsPropagator

All Implemented Interfaces:

Propagator, PVCoordinatesProvider

public class IntelsatElevenElementsPropagator
extends AbstractAnalyticalPropagator

This class provides elements to propagate Intelsat's 11 elements.

Intelsat's 11 elements propagation is defined in ITU-R S.1525 standard.

Since:

12.1

Author:

Bryan Cazabonne

Field Summary

Fields inherited from interface org.orekit.propagation.Propagator

DEFAULT_MASS

Constructor Summary

Constructors

Constructor	Description
IntelsatElevenElementsPropagator(IntelsatElevenElements elements)	Default constructor.
IntelsatElevenElementsPropagator(IntelsatElevenElements elements, Frame inertialFrame, Frame ecefFrame)	Constructor.
IntelsatElevenElementsPropagator(IntelsatElevenElements elements, Frame inertialFrame, Frame ecefFrame, AttitudeProvider attitudeProvider, double mass)	Constructor.

Method Summary

Modifier and Type	Method	Description
UnivariateDerivative2	getEastLongitudeDegrees()	Get the computed satellite's east longitude.
Frame	getFrame()	Get the frame in which the orbit is propagated.
UnivariateDerivative2	getGeocentricLatitudeDegrees()	Get the computed satellite's geocentric latitude.
protected double	getMass(AbsoluteDate date)	Get the mass.
UnivariateDerivative2	getOrbitRadius()	Get the computed satellite's orbit.
PVCoordinates	propagateInEcef(AbsoluteDate date)	Converts the Intelsat's 11 elements into Position/Velocity coordinates in ECEF.
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, getStartDate, initializeAdditionalStates, 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, getPosition, getPVCoordinates, setStepHandler, setStepHandler

Constructor Detail

IntelsatElevenElementsPropagator

@DefaultDataContext
public IntelsatElevenElementsPropagator(IntelsatElevenElements elements)

Default constructor.

This constructor uses the default data context.

The attitude provider is set by default to be aligned with the inertial frame.
The mass is set by default to the DEFAULT_MASS.
The inertial frame is set by default to the TOD frame in the default data context.
The ECEF frame is set by default to the CIO/2010-based ITRF simple EOP in the default data context.

Parameters:

elements - Intelsat's 11 elements

IntelsatElevenElementsPropagator

public IntelsatElevenElementsPropagator(IntelsatElevenElements elements,
                                        Frame inertialFrame,
                                        Frame ecefFrame)

Constructor.

The attitude provider is set by default to be aligned with the inertial frame.
The mass is set by default to the DEFAULT_MASS.

Parameters:

elements - Intelsat's 11 elements

inertialFrame - inertial frame for the output orbit

ecefFrame - ECEF frame related to the Intelsat's 11 elements

IntelsatElevenElementsPropagator

public IntelsatElevenElementsPropagator(IntelsatElevenElements elements,
                                        Frame inertialFrame,
                                        Frame ecefFrame,
                                        AttitudeProvider attitudeProvider,
                                        double mass)

Constructor.

Parameters:

elements - Intelsat's 11 elements

inertialFrame - inertial frame for the output orbit

ecefFrame - ECEF frame related to the Intelsat's 11 elements

attitudeProvider - attitude provider

mass - spacecraft mass

Method Detail

propagateInEcef

public PVCoordinates propagateInEcef(AbsoluteDate date)

Converts the Intelsat's 11 elements into Position/Velocity coordinates in ECEF.

Parameters:

date - computation epoch

Returns:

Position/Velocity coordinates in ECEF

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

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

getEastLongitudeDegrees

public UnivariateDerivative2 getEastLongitudeDegrees()

Get the computed satellite's east longitude.

Returns:

the satellite's east longitude in degrees

getGeocentricLatitudeDegrees

public UnivariateDerivative2 getGeocentricLatitudeDegrees()

Get the computed satellite's geocentric latitude.

Returns:

the satellite's geocentric latitude in degrees

getOrbitRadius

public UnivariateDerivative2 getOrbitRadius()

Get the computed satellite's orbit.

Returns:

satellite's orbit radius in meters

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)