fr.cnes.sirius.patrius.propagation.analytical.tle

Class TLEPropagator

java.lang.Object

fr.cnes.sirius.patrius.propagation.AbstractPropagator

fr.cnes.sirius.patrius.propagation.analytical.tle.TLEPropagator

All Implemented Interfaces:

PVCoordinatesProvider, Propagator, Serializable

public abstract class TLEPropagator
extends AbstractPropagator
implements Serializable

This class provides elements to propagate TLE's.

The models used are SGP4 and SDP4, initially proposed by NORAD as the unique convenient propagator for TLE's. Inputs and outputs of this propagator are only suited for NORAD two lines elements sets, since it uses estimations and mean values appropriate for TLE's only.

Deep- or near- space propagator is selected internally according to NORAD recommendations so that the user has not to worry about the used computation methods. One instance is created for each TLE (this instance can only be get using selectExtrapolator(TLE) method, and can compute position and velocity coordinates at any time. Maximum accuracy is guaranteed in a 24h range period before and after the provided TLE epoch (of course this accuracy is not really measurable nor predictable: according to CelesTrak, the precision is close to one kilometer and error won't probably rise above 2 km).

This implementation is largely inspired from the paper and source code Revisiting Spacetrack Report #3 and is fully compliant with its results and tests cases.

Author:

Felix R. Hoots, Ronald L. Roehrich, December 1980 (original fortran), David A. Vallado, Paul Crawford, Richard Hujsak, T.S. Kelso (C++ translation and improvements), Fabien Maussion (java translation)

See Also:

TLE, Serialized Form

Field Summary

Fields

Modifier and Type	Field and Description
protected double	a final semi major axis.
protected double	a0dp original recovered semi major axis.
protected double	beta0 sqrt (1 - e2).
protected double	beta02 1 - e2.
protected double	c1 C1 from SPTRCK #3.
protected double	c2 C2 from SPTRCK #3.
protected double	c4 C4 from SPTRCK #3.
protected double	coef coef for SGP C3 computation.
protected double	coef1 coef for SGP C5 computation.
protected double	cosi0 cosinus original inclination.
protected double	e final eccentricity.
protected double	e0sq original eccentricity squared.
protected double	eeta original eccentricity * eta.
protected double	eta eta from SPTRCK #3.
protected double	etasq eta squared.
protected double	i final inclination.
protected double	omega final perigee argument.
protected double	omgdot common parameter for perigee argument (omega) computation.
protected double	perige perigee, expressed in KM and ALTITUDE.
protected double	s4 s* new value for the contant s.
protected double	sini0 sinus original inclination.
protected double	t2cof 3/2 * C1.
protected double	theta2 cos io squared.
protected TLE	tle Initial state.
protected double	tsi tsi from SPTRCK #3.
protected double	xl L from SPTRCK #3.
protected double	xmdot common parameter for mean anomaly (M) computation.
protected double	xn0dp original recovered mean motion.
protected double	xnodcf common parameter for raan (OMEGA) computation.
protected double	xnode final RAAN.
protected double	xnodot common parameter for raan (OMEGA) computation.

Fields inherited from class fr.cnes.sirius.patrius.propagation.AbstractPropagator

MASS

Fields inherited from interface fr.cnes.sirius.patrius.propagation.Propagator

EPHEMERIS_GENERATION_MODE, MASTER_MODE, SLAVE_MODE

Constructor Summary

Constructors

Modifier	Constructor and Description
protected	TLEPropagator(TLE initialTLE, AttitudeProvider attitudeProviderForces, AttitudeProvider attitudeProviderEvents, MassProvider mass) Protected constructor for derived classes.
protected	TLEPropagator(TLE initialTLE, AttitudeProvider attitudeProvider, MassProvider mass) Protected constructor for derived classes.

Method Summary

Modifier and Type	Method and Description
Frame	getFrame() Get the frame in which the orbit is propagated.
PVCoordinates	getPVCoordinates(AbsoluteDate date) Get the extrapolated position and velocity from an initial TLE.
TLE	getTLE() Get the underlying TLE.
protected Orbit	propagateOrbit(AbsoluteDate date) Extrapolate an orbit up to a specific target date.
static TLEPropagator	selectExtrapolator(TLE tle) Selects the extrapolator to use with the selected TLE.
static TLEPropagator	selectExtrapolator(TLE tle, AttitudeProvider attitudeProviderForces, AttitudeProvider attitudeProviderEvents, MassProvider mass) Selects the extrapolator to use with the selected TLE.
static TLEPropagator	selectExtrapolator(TLE tle, AttitudeProvider attitudeProvider, MassProvider mass) Selects the extrapolator to use with the selected TLE.
protected abstract void	sxpInitialize() Initialization proper to each propagator (SGP or SDP).
protected abstract void	sxpPropagate(double t) Propagation proper to each propagator (SGP or SDP).

Methods inherited from class fr.cnes.sirius.patrius.propagation.AbstractPropagator

acceptStep, addAdditionalStateProvider, addAdditionalStateProvider, addEventDetector, basicPropagate, clearEventsDetectors, getAttitudeProvider, getAttitudeProviderEvents, getAttitudeProviderForces, getEventsDetectors, getGeneratedEphemeris, getInitialState, getMode, getPVCoordinates, getPvProvider, manageStateFrame, propagate, propagate, resetInitialState, setAttitudeProvider, setAttitudeProviderEvents, setAttitudeProviderForces, setEphemerisMode, setMasterMode, setMasterMode, setOrbitFrame, setSlaveMode, setStartDate

Methods inherited from class java.lang.Object

clone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, wait

Field Detail

tle

protected final TLE tle

Initial state.

xnode

protected double xnode

final RAAN.

a

protected double a

final semi major axis.

e

protected double e

final eccentricity.

i

protected double i

final inclination.

omega

protected double omega

final perigee argument.

xl

protected double xl

L from SPTRCK #3.

a0dp

protected double a0dp

original recovered semi major axis.

xn0dp

protected double xn0dp

original recovered mean motion.

cosi0

protected double cosi0

cosinus original inclination.

theta2

protected double theta2

cos io squared.

sini0

protected double sini0

sinus original inclination.

xmdot

protected double xmdot

common parameter for mean anomaly (M) computation.

omgdot

protected double omgdot

common parameter for perigee argument (omega) computation.

xnodot

protected double xnodot

common parameter for raan (OMEGA) computation.

e0sq

protected double e0sq

original eccentricity squared.

beta02

protected double beta02

1 - e2.

beta0

protected double beta0

sqrt (1 - e2).

perige

protected double perige

perigee, expressed in KM and ALTITUDE.

etasq

protected double etasq

eta squared.

eeta

protected double eeta

original eccentricity * eta.

s4

protected double s4

s* new value for the contant s.

tsi

protected double tsi

tsi from SPTRCK #3.

eta

protected double eta

eta from SPTRCK #3.

coef

protected double coef

coef for SGP C3 computation.

coef1

protected double coef1

coef for SGP C5 computation.

c1

protected double c1

C1 from SPTRCK #3.

c2

protected double c2

C2 from SPTRCK #3.

c4

protected double c4

C4 from SPTRCK #3.

xnodcf

protected double xnodcf

common parameter for raan (OMEGA) computation.

t2cof

protected double t2cof

3/2 * C1.

Constructor Detail

TLEPropagator

protected TLEPropagator(TLE initialTLE,
                        AttitudeProvider attitudeProvider,
                        MassProvider mass)
                 throws PatriusException

Protected constructor for derived classes.

Parameters:

initialTLE - the unique TLE to propagate

attitudeProvider - provider for attitude computation

mass - spacecraft mass (kg)

Throws:

PatriusException - if some specific error occurs

TLEPropagator

protected TLEPropagator(TLE initialTLE,
                        AttitudeProvider attitudeProviderForces,
                        AttitudeProvider attitudeProviderEvents,
                        MassProvider mass)
                 throws PatriusException

Protected constructor for derived classes.

Parameters:

initialTLE - the unique TLE to propagate

attitudeProviderForces - provider for attitude computation in forces computation case

attitudeProviderEvents - provider for attitude computation in events computation case

mass - spacecraft mass (kg)

Throws:

PatriusException - if some specific error occurs

Method Detail

selectExtrapolator

public static TLEPropagator selectExtrapolator(TLE tle)
                                        throws PatriusException

Selects the extrapolator to use with the selected TLE.

Parameters:

tle - the TLE to propagate.

Returns:

the correct propagator.

Throws:

PatriusException - if the underlying model cannot be initialized

selectExtrapolator

public static TLEPropagator selectExtrapolator(TLE tle,
                                               AttitudeProvider attitudeProvider,
                                               MassProvider mass)
                                        throws PatriusException

Selects the extrapolator to use with the selected TLE.

Parameters:

tle - the TLE to propagate.

attitudeProvider - provider for attitude computation

mass - spacecraft mass (kg)

Returns:

the correct propagator.

Throws:

PatriusException - if the underlying model cannot be initialized

selectExtrapolator

public static TLEPropagator selectExtrapolator(TLE tle,
                                               AttitudeProvider attitudeProviderForces,
                                               AttitudeProvider attitudeProviderEvents,
                                               MassProvider mass)
                                        throws PatriusException

Selects the extrapolator to use with the selected TLE.

Parameters:

tle - the TLE to propagate.

attitudeProviderForces - provider for attitude computation in forces computation case

attitudeProviderEvents - provider for attitude computation in events computation case

mass - spacecraft mass (kg)

Returns:

the correct propagator.

Throws:

PatriusException - if the underlying model cannot be initialized

getPVCoordinates

public PVCoordinates getPVCoordinates(AbsoluteDate date)
                               throws PatriusException

Get the extrapolated position and velocity from an initial TLE.

Parameters:

date - the final date

Returns:

the final PVCoordinates

Throws:

PatriusException - if propagation cannot be performed at given date

sxpInitialize

protected abstract void sxpInitialize()
                               throws PatriusException

Initialization proper to each propagator (SGP or SDP).

Throws:

PatriusException - if some specific error occurs

sxpPropagate

protected abstract void sxpPropagate(double t)
                              throws PatriusException

Propagation proper to each propagator (SGP or SDP).

Parameters:

t - the offset from initial epoch (min)

Throws:

PatriusException - if current state cannot be propagated

propagateOrbit

protected Orbit propagateOrbit(AbsoluteDate date)
                        throws PropagationException

Extrapolate an orbit up to a specific target date.

Specified by:

propagateOrbit in class AbstractPropagator

Parameters:

date - target date for the orbit

Returns:

extrapolated parameters

Throws:

PropagationException - if some parameters are out of bounds

getTLE

public TLE getTLE()

Get the underlying TLE.

Returns:

underlying TLE

getFrame

public Frame getFrame()

Get the frame in which the orbit is propagated.

4 cases are possible:

• The propagation frame has been defined (using Propagator.setOrbitFrame(Frame)): it is returned.
• The propagation frame has not been defined and the initial state has been provided and is expressed in a pseudo-inertial frame: the initial state frame is returned.
• The propagation frame has not been defined and the initial state has been provided and is not expressed in a pseudo-inertial frame: null is returned.
• The propagation frame has not been defined and the initial state has not been provided: null is returned.

Specified by:

getFrame in interface Propagator

Overrides:

getFrame in class AbstractPropagator

Returns:

frame in which the orbit is propagated