org.orekit.orbits

Class KeplerianOrbit

java.lang.Object

org.orekit.orbits.Orbit

org.orekit.orbits.KeplerianOrbit

All Implemented Interfaces:

Serializable, TimeInterpolable<Orbit>, TimeShiftable<Orbit>, TimeStamped, PVCoordinatesProvider

public class KeplerianOrbit
extends Orbit

This class handles traditional keplerian orbital parameters.

The parameters used internally are the classical keplerian elements:

     a
     e
     i
     ω
     Ω
     v
   

where ω stands for the Perigee Argument, Ω stands for the Right Ascension of the Ascending Node and v stands for the true anomaly.

This class supports hyperbolic orbits, using the convention that semi major axis is negative for such orbits (and of course eccentricity is greater than 1).

When orbit is either equatorial or circular, some keplerian elements (more precisely ω and Ω) become ambiguous so this class should not be used for such orbits. For this reason, equinoctial orbits is the recommended way to represent orbits.

The instance KeplerianOrbit is guaranteed to be immutable.

Author:

Luc Maisonobe, Guylaine Prat, Fabien Maussion, Véronique Pommier-Maurussane

See Also:

Orbit, CircularOrbit, CartesianOrbit, EquinoctialOrbit, Serialized Form

Constructor Summary

Constructors

Constructor and Description
KeplerianOrbit(double a, double e, double i, double pa, double raan, double anomaly, PositionAngle type, Frame frame, AbsoluteDate date, double mu) Creates a new instance.
KeplerianOrbit(Orbit op) Constructor from any kind of orbital parameters.
KeplerianOrbit(PVCoordinates pvCoordinates, Frame frame, AbsoluteDate date, double mu) Constructor from cartesian parameters.
KeplerianOrbit(TimeStampedPVCoordinates pvCoordinates, Frame frame, double mu) Constructor from cartesian parameters.

Method Summary

Modifier and Type	Method and Description
void	addKeplerContribution(PositionAngle type, double gm, double[] pDot) Add the contribution of the Keplerian motion to parameters derivatives
protected double[][]	computeJacobianEccentricWrtCartesian() Compute the Jacobian of the orbital parameters with eccentric angle with respect to the Cartesian parameters.
protected double[][]	computeJacobianMeanWrtCartesian() Compute the Jacobian of the orbital parameters with mean angle with respect to the Cartesian parameters.
protected double[][]	computeJacobianTrueWrtCartesian() Compute the Jacobian of the orbital parameters with true angle with respect to the Cartesian parameters.
double	getA() Get the semi-major axis.
double	getAnomaly(PositionAngle type) Get the anomaly.
double	getE() Get the eccentricity.
double	getEccentricAnomaly() Get the eccentric anomaly.
double	getEquinoctialEx() Get the first component of the equinoctial eccentricity vector.
double	getEquinoctialEy() Get the second component of the equinoctial eccentricity vector.
double	getHx() Get the first component of the inclination vector.
double	getHy() Get the second component of the inclination vector.
double	getI() Get the inclination.
double	getLE() Get the eccentric longitude argument.
double	getLM() Get the mean longitude argument.
double	getLv() Get the true longitude argument.
double	getMeanAnomaly() Get the mean anomaly.
double	getPerigeeArgument() Get the perigee argument.
double	getRightAscensionOfAscendingNode() Get the right ascension of the ascending node.
double	getTrueAnomaly() Get the true anomaly.
OrbitType	getType() Get the orbit type.
protected TimeStampedPVCoordinates	initPVCoordinates() Compute the position/velocity coordinates from the canonical parameters.
KeplerianOrbit	interpolate(AbsoluteDate date, Collection<Orbit> sample) Get an interpolated instance.
KeplerianOrbit	shiftedBy(double dt) Get a time-shifted orbit.
String	toString() Returns a string representation of this keplerian parameters object.

Methods inherited from class org.orekit.orbits.Orbit

fillHalfRow, fillHalfRow, fillHalfRow, fillHalfRow, fillHalfRow, fillHalfRow, getDate, getFrame, getJacobianWrtCartesian, getJacobianWrtParameters, getKeplerianMeanMotion, getKeplerianPeriod, getMu, getPVCoordinates, getPVCoordinates, getPVCoordinates

Methods inherited from class java.lang.Object

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

Constructor Detail

KeplerianOrbit

public KeplerianOrbit(double a,
                      double e,
                      double i,
                      double pa,
                      double raan,
                      double anomaly,
                      PositionAngle type,
                      Frame frame,
                      AbsoluteDate date,
                      double mu)
               throws IllegalArgumentException

Creates a new instance.

Parameters:

a - semi-major axis (m), negative for hyperbolic orbits

e - eccentricity

i - inclination (rad)

pa - perigee argument (ω, rad)

raan - right ascension of ascending node (Ω, rad)

anomaly - mean, eccentric or true anomaly (rad)

type - type of anomaly

frame - the frame in which the parameters are defined (must be a pseudo-inertial frame)

date - date of the orbital parameters

mu - central attraction coefficient (m³/s²)

Throws:

IllegalArgumentException - if frame is not a pseudo-inertial frame or a and e don't match for hyperbolic orbits, or v is out of range for hyperbolic orbits

KeplerianOrbit

public KeplerianOrbit(TimeStampedPVCoordinates pvCoordinates,
                      Frame frame,
                      double mu)
               throws IllegalArgumentException

Constructor from cartesian parameters.

The acceleration provided in pvCoordinates is accessible using Orbit.getPVCoordinates() and Orbit.getPVCoordinates(Frame). All other methods use mu and the position to compute the acceleration, including shiftedBy(double) and Orbit.getPVCoordinates(AbsoluteDate, Frame).

Parameters:

pvCoordinates - the PVCoordinates of the satellite

frame - the frame in which are defined the PVCoordinates (must be a pseudo-inertial frame)

mu - central attraction coefficient (m³/s²)

Throws:

IllegalArgumentException - if frame is not a pseudo-inertial frame

KeplerianOrbit

public KeplerianOrbit(PVCoordinates pvCoordinates,
                      Frame frame,
                      AbsoluteDate date,
                      double mu)
               throws IllegalArgumentException

Constructor from cartesian parameters.

The acceleration provided in pvCoordinates is accessible using Orbit.getPVCoordinates() and Orbit.getPVCoordinates(Frame). All other methods use mu and the position to compute the acceleration, including shiftedBy(double) and Orbit.getPVCoordinates(AbsoluteDate, Frame).

Parameters:

pvCoordinates - the PVCoordinates of the satellite

frame - the frame in which are defined the PVCoordinates (must be a pseudo-inertial frame)

date - date of the orbital parameters

mu - central attraction coefficient (m³/s²)

Throws:

IllegalArgumentException - if frame is not a pseudo-inertial frame

KeplerianOrbit

public KeplerianOrbit(Orbit op)

Constructor from any kind of orbital parameters.

Parameters:

op - orbital parameters to copy

Method Detail

getType

public OrbitType getType()

Get the orbit type.

Specified by:

getType in class Orbit

Returns:

orbit type

getA

public double getA()

Get the semi-major axis.

Note that the semi-major axis is considered negative for hyperbolic orbits.

Specified by:

getA in class Orbit

Returns:

semi-major axis (m)

getE

public double getE()

Get the eccentricity.

Specified by:

getE in class Orbit

Returns:

eccentricity

getI

public double getI()

Get the inclination.

Specified by:

getI in class Orbit

Returns:

inclination (rad)

getPerigeeArgument

public double getPerigeeArgument()

Get the perigee argument.

Returns:

perigee argument (rad)

getRightAscensionOfAscendingNode

public double getRightAscensionOfAscendingNode()

Get the right ascension of the ascending node.

Returns:

right ascension of the ascending node (rad)

getAnomaly

public double getAnomaly(PositionAngle type)

Get the anomaly.

Parameters:

type - type of the angle

Returns:

anomaly (rad)

getTrueAnomaly

public double getTrueAnomaly()

Get the true anomaly.

Returns:

true anomaly (rad)

getEccentricAnomaly

public double getEccentricAnomaly()

Get the eccentric anomaly.

Returns:

eccentric anomaly (rad)

getMeanAnomaly

public double getMeanAnomaly()

Get the mean anomaly.

Returns:

mean anomaly (rad)

getEquinoctialEx

public double getEquinoctialEx()

Get the first component of the equinoctial eccentricity vector.

Specified by:

getEquinoctialEx in class Orbit

Returns:

first component of the equinoctial eccentricity vector

getEquinoctialEy

public double getEquinoctialEy()

Get the second component of the equinoctial eccentricity vector.

Specified by:

getEquinoctialEy in class Orbit

Returns:

second component of the equinoctial eccentricity vector

getHx

public double getHx()

Get the first component of the inclination vector.

Specified by:

getHx in class Orbit

Returns:

first component of the inclination vector

getHy

public double getHy()

Get the second component of the inclination vector.

Specified by:

getHy in class Orbit

Returns:

second component of the inclination vector

getLv

public double getLv()

Get the true longitude argument.

Specified by:

getLv in class Orbit

Returns:

v + ω + Ω true longitude argument (rad)

getLE

public double getLE()

Get the eccentric longitude argument.

Specified by:

getLE in class Orbit

Returns:

E + ω + Ω eccentric longitude argument (rad)

getLM

public double getLM()

Get the mean longitude argument.

Specified by:

getLM in class Orbit

Returns:

M + ω + Ω mean longitude argument (rad)

initPVCoordinates

protected TimeStampedPVCoordinates initPVCoordinates()

Compute the position/velocity coordinates from the canonical parameters.

Specified by:

initPVCoordinates in class Orbit

Returns:

computed position/velocity coordinates

shiftedBy

public KeplerianOrbit shiftedBy(double dt)

Get a time-shifted orbit.

The orbit can be slightly shifted to close dates. This shift is based on a simple keplerian model. It is not intended as a replacement for proper orbit and attitude propagation but should be sufficient for small time shifts or coarse accuracy.

Specified by:

shiftedBy in interface TimeShiftable<Orbit>

Specified by:

shiftedBy in class Orbit

Parameters:

dt - time shift in seconds

Returns:

a new orbit, shifted with respect to the instance (which is immutable)

interpolate

public KeplerianOrbit interpolate(AbsoluteDate date,
                                  Collection<Orbit> sample)

Get an interpolated instance.

Note that the state of the current instance may not be used in the interpolation process, only its type and non interpolable fields are used (for example central attraction coefficient or frame when interpolating orbits). The interpolable fields taken into account are taken only from the states of the sample points. So if the state of the instance must be used, the instance should be included in the sample points.

Note that this method is designed for small samples only (say up to about 10-20 points) so it can be implemented using polynomial interpolation (typically Hermite interpolation). Using too much points may induce Runge's phenomenon and numerical problems (including NaN appearing).

The interpolated instance is created by polynomial Hermite interpolation on Keplerian elements, without derivatives (which means the interpolation falls back to Lagrange interpolation only).

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).

If orbit interpolation on large samples is needed, using the Ephemeris class is a better way than using this low-level interpolation. The Ephemeris class automatically handles selection of a neighboring sub-sample with a predefined number of point from a large global sample in a thread-safe way.

Parameters:

date - interpolation date

sample - sample points on which interpolation should be done

Returns:

a new instance, interpolated at specified date

computeJacobianMeanWrtCartesian

protected double[][] computeJacobianMeanWrtCartesian()

Compute the Jacobian of the orbital parameters with mean angle with respect to the Cartesian parameters.

Element jacobian[i][j] is the derivative of parameter i of the orbit with respect to Cartesian coordinate j. This means each row correspond to one orbital parameter whereas columns 0 to 5 correspond to the Cartesian coordinates x, y, z, xDot, yDot and zDot.

Specified by:

computeJacobianMeanWrtCartesian in class Orbit

Returns:

6x6 Jacobian matrix

See Also:

Orbit.computeJacobianEccentricWrtCartesian(), Orbit.computeJacobianTrueWrtCartesian()

computeJacobianEccentricWrtCartesian

protected double[][] computeJacobianEccentricWrtCartesian()

Compute the Jacobian of the orbital parameters with eccentric angle with respect to the Cartesian parameters.

Element jacobian[i][j] is the derivative of parameter i of the orbit with respect to Cartesian coordinate j. This means each row correspond to one orbital parameter whereas columns 0 to 5 correspond to the Cartesian coordinates x, y, z, xDot, yDot and zDot.

Specified by:

computeJacobianEccentricWrtCartesian in class Orbit

Returns:

6x6 Jacobian matrix

See Also:

Orbit.computeJacobianMeanWrtCartesian(), Orbit.computeJacobianTrueWrtCartesian()

computeJacobianTrueWrtCartesian

protected double[][] computeJacobianTrueWrtCartesian()

Compute the Jacobian of the orbital parameters with true angle with respect to the Cartesian parameters.

Element jacobian[i][j] is the derivative of parameter i of the orbit with respect to Cartesian coordinate j. This means each row correspond to one orbital parameter whereas columns 0 to 5 correspond to the Cartesian coordinates x, y, z, xDot, yDot and zDot.

Specified by:

computeJacobianTrueWrtCartesian in class Orbit

Returns:

6x6 Jacobian matrix

See Also:

Orbit.computeJacobianMeanWrtCartesian(), Orbit.computeJacobianEccentricWrtCartesian()

addKeplerContribution

public void addKeplerContribution(PositionAngle type,
                                  double gm,
                                  double[] pDot)

Add the contribution of the Keplerian motion to parameters derivatives

This method is used by integration-based propagators to evaluate the part of Keplerian motion to evolution of the orbital state.

Specified by:

addKeplerContribution in class Orbit

Parameters:

type - type of the position angle in the state

gm - attraction coefficient to use

pDot - array containing orbital state derivatives to update (the Keplerian part must be added to the array components, as the array may already contain some non-zero elements corresponding to non-Keplerian parts)

toString

public String toString()

Returns a string representation of this keplerian parameters object.

Overrides:

toString in class Object

Returns:

a string representation of this object