Package org.orekit.propagation

Class FieldSpacecraftState<T extends org.hipparchus.RealFieldElement<T>>

java.lang.Object

org.orekit.propagation.FieldSpacecraftState<T>

All Implemented Interfaces:

FieldTimeInterpolable<FieldSpacecraftState<T>,T>, FieldTimeShiftable<FieldSpacecraftState<T>,T>, FieldTimeStamped<T>

public class FieldSpacecraftState<T extends org.hipparchus.RealFieldElement<T>>
extends Object
implements FieldTimeStamped<T>, FieldTimeShiftable<FieldSpacecraftState<T>,T>, FieldTimeInterpolable<FieldSpacecraftState<T>,T>

This class is the representation of a complete state holding orbit, attitude and mass information at a given date.

It contains an orbital state at a current FieldAbsoluteDate both handled by an FieldOrbit, plus the current mass and attitude. FieldOrbitand state are guaranteed to be consistent in terms of date and reference frame. The spacecraft state may also contain additional states, which are simply named double arrays which can hold any user-defined data.

The state can be slightly shifted to close dates. This shift is based on a simple Keplerian model for orbit, a linear extrapolation for attitude taking the spin rate into account and no mass change. It is not intended as a replacement for proper orbit and attitude propagation but should be sufficient for either small time shifts or coarse accuracy.

The instance FieldSpacecraftState is guaranteed to be immutable.

Author:

Fabien Maussion, Véronique Pommier-Maurussane, Luc Maisonobe

See Also:

NumericalPropagator

Constructor Summary

Constructors

Constructor	Description
FieldSpacecraftState(org.hipparchus.Field<T> field, SpacecraftState state)	Convert a SpacecraftState.
FieldSpacecraftState(FieldOrbit<T> orbit)	Build a spacecraft state from orbit only.
FieldSpacecraftState(FieldOrbit<T> orbit, Map<String,T[]> additional)	Build a spacecraft state from orbit only.
FieldSpacecraftState(FieldOrbit<T> orbit, FieldAttitude<T> attitude)	Build a spacecraft state from orbit and attitude provider.
FieldSpacecraftState(FieldOrbit<T> orbit, FieldAttitude<T> attitude, Map<String,T[]> additional)	Build a spacecraft state from orbit and attitude provider.
FieldSpacecraftState(FieldOrbit<T> orbit, FieldAttitude<T> attitude, T mass)	Build a spacecraft state from orbit, attitude provider and mass.
FieldSpacecraftState(FieldOrbit<T> orbit, FieldAttitude<T> attitude, T mass, Map<String,T[]> additional)	Build a spacecraft state from orbit, attitude provider and mass.
FieldSpacecraftState(FieldOrbit<T> orbit, T mass)	Create a new instance from orbit and mass.
FieldSpacecraftState(FieldOrbit<T> orbit, T mass, Map<String,T[]> additional)	Create a new instance from orbit and mass.

Method Summary

Modifier and Type	Method	Description
FieldSpacecraftState<T>	addAdditionalState(String name, T... value)	Add an additional state.
void	ensureCompatibleAdditionalStates(FieldSpacecraftState<T> state)	Check if two instances have the same set of additional states available.
T	getA()	Get the semi-major axis.
T[]	getAdditionalState(String name)	Get an additional state.
Map<String,T[]>	getAdditionalStates()	Get an unmodifiable map of additional states.
FieldAttitude<T>	getAttitude()	Get the attitude.
FieldAbsoluteDate<T>	getDate()	Get the date.
T	getE()	Get the eccentricity.
T	getEquinoctialEx()	Get the first component of the eccentricity vector (as per equinoctial parameters).
T	getEquinoctialEy()	Get the second component of the eccentricity vector (as per equinoctial parameters).
Frame	getFrame()	Get the inertial frame.
T	getHx()	Get the first component of the inclination vector (as per equinoctial parameters).
T	getHy()	Get the second component of the inclination vector (as per equinoctial parameters).
T	getI()	Get the inclination.
T	getKeplerianMeanMotion()	Get the Keplerian mean motion.
T	getKeplerianPeriod()	Get the Keplerian period.
T	getLE()	Get the eccentric latitude argument (as per equinoctial parameters).
T	getLM()	Get the mean latitude argument (as per equinoctial parameters).
T	getLv()	Get the true latitude argument (as per equinoctial parameters).
T	getMass()	Gets the current mass.
double	getMu()	Get the central attraction coefficient.
FieldOrbit<T>	getOrbit()	Gets the current orbit.
TimeStampedFieldPVCoordinates<T>	getPVCoordinates()	Get the TimeStampedFieldPVCoordinates in orbit definition frame.
TimeStampedFieldPVCoordinates<T>	getPVCoordinates(Frame outputFrame)	Get the TimeStampedFieldPVCoordinates in given output frame.
boolean	hasAdditionalState(String name)	Check if an additional state is available.
FieldSpacecraftState<T>	interpolate(FieldAbsoluteDate<T> date, Stream<FieldSpacecraftState<T>> sample)	Get an interpolated instance.
FieldSpacecraftState<T>	shiftedBy(double dt)	Get a time-shifted state.
FieldSpacecraftState<T>	shiftedBy(T dt)	Get a time-shifted state.
SpacecraftState	toSpacecraftState()	To convert a FieldSpacecraftState instance into a SpacecraftState instance.
FieldTransform<T>	toTransform()	Compute the transform from orbite/attitude reference frame to spacecraft frame.

Methods inherited from interface org.orekit.time.FieldTimeInterpolable

interpolate

Methods inherited from class java.lang.Object

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

Constructor Detail

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit)

Build a spacecraft state from orbit only.

FieldAttitude and mass are set to unspecified non-null arbitrary values.

Parameters:

orbit - the orbit

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            FieldAttitude<T> attitude)
                     throws IllegalArgumentException

Build a spacecraft state from orbit and attitude provider.

Mass is set to an unspecified non-null arbitrary value.

Parameters:

orbit - the orbit

attitude - attitude

Throws:

IllegalArgumentException - if orbit and attitude dates or frames are not equal

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            T mass)

Create a new instance from orbit and mass.

FieldAttitude law is set to an unspecified default attitude.

Parameters:

orbit - the orbit

mass - the mass (kg)

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            FieldAttitude<T> attitude,
                            T mass)
                     throws IllegalArgumentException

Build a spacecraft state from orbit, attitude provider and mass.

Parameters:

orbit - the orbit

attitude - attitude

mass - the mass (kg)

Throws:

IllegalArgumentException - if orbit and attitude dates or frames are not equal

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            Map<String,T[]> additional)

Build a spacecraft state from orbit only.

FieldAttitude and mass are set to unspecified non-null arbitrary values.

Parameters:

orbit - the orbit

additional - additional states

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            FieldAttitude<T> attitude,
                            Map<String,T[]> additional)
                     throws IllegalArgumentException

Build a spacecraft state from orbit and attitude provider.

Mass is set to an unspecified non-null arbitrary value.

Parameters:

orbit - the orbit

attitude - attitude

additional - additional states

Throws:

IllegalArgumentException - if orbit and attitude dates or frames are not equal

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            T mass,
                            Map<String,T[]> additional)

Create a new instance from orbit and mass.

FieldAttitude law is set to an unspecified default attitude.

Parameters:

orbit - the orbit

mass - the mass (kg)

additional - additional states

FieldSpacecraftState

public FieldSpacecraftState(FieldOrbit<T> orbit,
                            FieldAttitude<T> attitude,
                            T mass,
                            Map<String,T[]> additional)
                     throws IllegalArgumentException

Build a spacecraft state from orbit, attitude provider and mass.

Parameters:

orbit - the orbit

attitude - attitude

mass - the mass (kg)

additional - additional states (may be null if no additional states are available)

Throws:

IllegalArgumentException - if orbit and attitude dates or frames are not equal

FieldSpacecraftState

public FieldSpacecraftState(org.hipparchus.Field<T> field,
                            SpacecraftState state)

Convert a SpacecraftState.

Parameters:

field - field to which the elements belong

state - state to convert

Method Detail

addAdditionalState

@SafeVarargs
public final FieldSpacecraftState<T> addAdditionalState(String name,
                                                        T... value)

Add an additional state.

SpacecraftState instances are immutable, so this method does not change the instance, but rather creates a new instance, which has the same orbit, attitude, mass and additional states as the original instance, except it also has the specified state. If the original instance already had an additional state with the same name, it will be overridden. If it did not have any additional state with that name, the new instance will have one more additional state than the original instance.

Parameters:

name - name of the additional state

value - value of the additional state

Returns:

a new instance, with the additional state added

See Also:

hasAdditionalState(String), getAdditionalState(String), getAdditionalStates()

shiftedBy

public FieldSpacecraftState<T> shiftedBy(double dt)

Get a time-shifted state.

The state can be slightly shifted to close dates. This shift is based on a simple Keplerian model for orbit, a linear extrapolation for attitude taking the spin rate into account and neither mass nor additional states changes. It is not intended as a replacement for proper orbit and attitude propagation but should be sufficient for small time shifts or coarse accuracy.

As a rough order of magnitude, the following table shows the extrapolation errors obtained between this simple shift method and an numerical propagator for a low Earth Sun Synchronous Orbit, with a 20x20 gravity field, Sun and Moon third bodies attractions, drag and solar radiation pressure. Beware that these results will be different for other orbits.

Extrapolation Error

interpolation time (s)	position error without derivatives (m)	position error with derivatives (m)
60	18	1.1
120	72	9.1
300	447	140
600	1601	1067
900	3141	3307

Specified by:

shiftedBy in interface FieldTimeShiftable<FieldSpacecraftState<T extends org.hipparchus.RealFieldElement<T>>,T extends org.hipparchus.RealFieldElement<T>>

Parameters:

dt - time shift in seconds

Returns:

a new state, shifted with respect to the instance (which is immutable) except for the mass which stay unchanged

shiftedBy

public FieldSpacecraftState<T> shiftedBy(T dt)

Get a time-shifted state.

The state can be slightly shifted to close dates. This shift is based on a simple Keplerian model for orbit, a linear extrapolation for attitude taking the spin rate into account and neither mass nor additional states changes. It is not intended as a replacement for proper orbit and attitude propagation but should be sufficient for small time shifts or coarse accuracy.

As a rough order of magnitude, the following table shows the extrapolation errors obtained between this simple shift method and an numerical propagator for a low Earth Sun Synchronous Orbit, with a 20x20 gravity field, Sun and Moon third bodies attractions, drag and solar radiation pressure. Beware that these results will be different for other orbits.

Extrapolation Error

interpolation time (s)	position error without derivatives (m)	position error with derivatives (m)
60	18	1.1
120	72	9.1
300	447	140
600	1601	1067
900	3141	3307

Specified by:

shiftedBy in interface FieldTimeShiftable<FieldSpacecraftState<T extends org.hipparchus.RealFieldElement<T>>,T extends org.hipparchus.RealFieldElement<T>>

Parameters:

dt - time shift in seconds

Returns:

a new state, shifted with respect to the instance (which is immutable) except for the mass which stay unchanged

interpolate

public FieldSpacecraftState<T> interpolate(FieldAbsoluteDate<T> date,
                                           Stream<FieldSpacecraftState<T>> sample)

Get an interpolated instance.

The additional states that are interpolated are the ones already present in the instance. The sample instances must therefore have at least the same additional states has the instance. They may have more additional states, but the extra ones will be ignored.

As this implementation of interpolation is polynomial, it should be used only with small samples (about 10-20 points) in order to avoid Runge's phenomenon and numerical problems (including NaN appearing).

Specified by:

interpolate in interface FieldTimeInterpolable<FieldSpacecraftState<T extends org.hipparchus.RealFieldElement<T>>,T extends org.hipparchus.RealFieldElement<T>>

Parameters:

date - interpolation date

sample - sample points on which interpolation should be done

Returns:

a new instance, interpolated at specified date

getOrbit

public FieldOrbit<T> getOrbit()

Gets the current orbit.

Returns:

the orbit

getDate

public FieldAbsoluteDate<T> getDate()

Get the date.

Specified by:

getDate in interface FieldTimeStamped<T extends org.hipparchus.RealFieldElement<T>>

Returns:

date

getFrame

public Frame getFrame()

Get the inertial frame.

Returns:

the frame

hasAdditionalState

public boolean hasAdditionalState(String name)

Check if an additional state is available.

Parameters:

name - name of the additional state

Returns:

true if the additional state is available

See Also:

addAdditionalState(String, RealFieldElement...), getAdditionalState(String), getAdditionalStates()

ensureCompatibleAdditionalStates

public void ensureCompatibleAdditionalStates(FieldSpacecraftState<T> state)
                                      throws org.hipparchus.exception.MathIllegalArgumentException

Check if two instances have the same set of additional states available.

Only the names and dimensions of the additional states are compared, not their values.

Parameters:

state - state to compare to instance

Throws:

org.hipparchus.exception.MathIllegalArgumentException - if an additional state does not have the same dimension in both states

getAdditionalState

public T[] getAdditionalState(String name)

Get an additional state.

Parameters:

name - name of the additional state

Returns:

value of the additional state

See Also:

addAdditionalState(String, RealFieldElement...), hasAdditionalState(String), getAdditionalStates()

getAdditionalStates

public Map<String,T[]> getAdditionalStates()

Get an unmodifiable map of additional states.

Returns:

unmodifiable map of additional states

See Also:

addAdditionalState(String, RealFieldElement...), hasAdditionalState(String), getAdditionalState(String)

toTransform

public FieldTransform<T> toTransform()

Compute the transform from orbite/attitude reference frame to spacecraft frame.

The spacecraft frame origin is at the point defined by the orbit, and its orientation is defined by the attitude.

Returns:

transform from specified frame to current spacecraft frame

getMu

public double getMu()

Get the central attraction coefficient.

Returns:

mu central attraction coefficient (m^3/s^2)

getKeplerianPeriod

public T getKeplerianPeriod()

Get the Keplerian period.

The Keplerian period is computed directly from semi major axis and central acceleration constant.

Returns:

Keplerian period in seconds

getKeplerianMeanMotion

public T getKeplerianMeanMotion()

Get the Keplerian mean motion.

The Keplerian mean motion is computed directly from semi major axis and central acceleration constant.

Returns:

Keplerian mean motion in radians per second

getA

public T getA()

Get the semi-major axis.

Returns:

semi-major axis (m)

getEquinoctialEx

public T getEquinoctialEx()

Get the first component of the eccentricity vector (as per equinoctial parameters).

Returns:

e cos(ω + Ω), first component of eccentricity vector

See Also:

getE()

getEquinoctialEy

public T getEquinoctialEy()

Get the second component of the eccentricity vector (as per equinoctial parameters).

Returns:

e sin(ω + Ω), second component of the eccentricity vector

See Also:

getE()

getHx

public T getHx()

Get the first component of the inclination vector (as per equinoctial parameters).

Returns:

tan(i/2) cos(Ω), first component of the inclination vector

See Also:

getI()

getHy

public T getHy()

Get the second component of the inclination vector (as per equinoctial parameters).

Returns:

tan(i/2) sin(Ω), second component of the inclination vector

See Also:

getI()

getLv

public T getLv()

Get the true latitude argument (as per equinoctial parameters).

Returns:

v + ω + Ω true latitude argument (rad)

See Also:

getLE(), getLM()

getLE

public T getLE()

Get the eccentric latitude argument (as per equinoctial parameters).

Returns:

E + ω + Ω eccentric latitude argument (rad)

See Also:

getLv(), getLM()

getLM

public T getLM()

Get the mean latitude argument (as per equinoctial parameters).

Returns:

M + ω + Ω mean latitude argument (rad)

See Also:

getLv(), getLE()

getE

public T getE()

Get the eccentricity.

Returns:

eccentricity

See Also:

getEquinoctialEx(), getEquinoctialEy()

getI

public T getI()

Get the inclination.

Returns:

inclination (rad)

See Also:

getHx(), getHy()

getPVCoordinates

public TimeStampedFieldPVCoordinates<T> getPVCoordinates()

Get the TimeStampedFieldPVCoordinates in orbit definition frame. Compute the position and velocity of the satellite. This method caches its results, and recompute them only when the method is called with a new value for mu. The result is provided as a reference to the internally cached TimeStampedFieldPVCoordinates, so the caller is responsible to copy it in a separate TimeStampedFieldPVCoordinates if it needs to keep the value for a while.

Returns:

pvCoordinates in orbit definition frame

getPVCoordinates

public TimeStampedFieldPVCoordinates<T> getPVCoordinates(Frame outputFrame)

Get the TimeStampedFieldPVCoordinates in given output frame. Compute the position and velocity of the satellite. This method caches its results, and recompute them only when the method is called with a new value for mu. The result is provided as a reference to the internally cached TimeStampedFieldPVCoordinates, so the caller is responsible to copy it in a separate TimeStampedFieldPVCoordinates if it needs to keep the value for a while.

Parameters:

outputFrame - frame in which coordinates should be defined

Returns:

pvCoordinates in orbit definition frame

getAttitude

public FieldAttitude<T> getAttitude()

Get the attitude.

Returns:

the attitude.

getMass

public T getMass()

Gets the current mass.

Returns:

the mass (kg)

toSpacecraftState

public SpacecraftState toSpacecraftState()

To convert a FieldSpacecraftState instance into a SpacecraftState instance.

Returns:

SpacecraftState instance with the same properties