Package org.orekit.propagation

Class SpacecraftState

  • All Implemented Interfaces:
    Serializable, TimeInterpolable<SpacecraftState>, TimeShiftable<SpacecraftState>, TimeStamped
    public class SpacecraftState
extends Object
implements TimeStamped, TimeShiftable<SpacecraftState>, TimeInterpolable<SpacecraftState>, Serializable
    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 AbsoluteDate both handled by an Orbit, plus the current mass and attitude. Orbit and 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 SpacecraftState is guaranteed to be immutable.

    Author:
    Fabien Maussion, Véronique Pommier-Maurussane, Luc Maisonobe
    See Also:
    NumericalPropagator, Serialized Form

    • Constructor Detail

      • SpacecraftState

        public SpacecraftState​(Orbit orbit)
        Build a spacecraft state from orbit only.

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

        Parameters:
        orbit - the orbit

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       Attitude 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

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       double mass)
        Create a new instance from orbit and mass.

        Attitude law is set to an unspecified default attitude.

        Parameters:
        orbit - the orbit
        mass - the mass (kg)

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       Attitude attitude,
                       double 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

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       Map<String,​double[]> additional)
        Build a spacecraft state from orbit only.

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

        Parameters:
        orbit - the orbit
        additional - additional states

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       Attitude attitude,
                       Map<String,​double[]> 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

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       double mass,
                       Map<String,​double[]> additional)
        Create a new instance from orbit and mass.

        Attitude law is set to an unspecified default attitude.

        Parameters:
        orbit - the orbit
        mass - the mass (kg)
        additional - additional states

      • SpacecraftState

        public SpacecraftState​(Orbit orbit,
                       Attitude attitude,
                       double mass,
                       Map<String,​double[]> 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

    • Method Detail

      • addAdditionalState

        public SpacecraftState addAdditionalState​(String name,
                                          double... 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 SpacecraftState shiftedBy​(double dt)
        Get a time-shifted state.

        The state can be slightly shifted to close dates. This shift is based on simple models. For orbits, the model is a Keplerian one if no derivatives are available in the orbit, or Keplerian plus quadratic effect of the non-Keplerian acceleration if derivatives are available. For attitude, a polynomial model is used. Neither mass nor additional states change. Shifting 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
        60016011067
        90031413307
        Specified by:
        shiftedBy in interface TimeShiftable<SpacecraftState>
        Parameters:
        dt - time shift in seconds
        Returns:
        a new state, shifted with respect to the instance (which is immutable) except for the mass and additional states which stay unchanged

      • interpolate

        public SpacecraftState interpolate​(AbsoluteDate date,
                                   Stream<SpacecraftState> 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 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 TimeInterpolable<SpacecraftState>
        Parameters:
        date - interpolation date
        sample - sample points on which interpolation should be done
        Returns:
        a new instance, interpolated at specified date

      • getOrbit

        public Orbit getOrbit()
        Gets the current orbit.
        Returns:
        the orbit

      • getFrame

        public Frame getFrame()
        Get the inertial frame.
        Returns:
        the frame

      • ensureCompatibleAdditionalStates

        public void ensureCompatibleAdditionalStates​(SpacecraftState state)
                                      throws org.hipparchus.exception.MathIllegalStateException
        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.MathIllegalStateException - if an additional state does not have the same dimension in both states

      • toTransform

        public Transform 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 double 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 double 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 double getA()
        Get the semi-major axis.
        Returns:
        semi-major axis (m)

      • getEquinoctialEx

        public double 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 double 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 double 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 double 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 double getLv()
        Get the true longitude argument (as per equinoctial parameters).
        Returns:
        v + ω + Ω true longitude argument (rad)
        See Also:
        getLE(), getLM()

      • getLE

        public double getLE()
        Get the eccentric longitude argument (as per equinoctial parameters).
        Returns:
        E + ω + Ω eccentric longitude argument (rad)
        See Also:
        getLv(), getLM()

      • getLM

        public double getLM()
        Get the mean longitude argument (as per equinoctial parameters).
        Returns:
        M + ω + Ω mean longitude argument (rad)
        See Also:
        getLv(), getLE()

      • getI

        public double getI()
        Get the inclination.
        Returns:
        inclination (rad)
        See Also:
        getHx(), getHy()

      • getPVCoordinates

        public TimeStampedPVCoordinates getPVCoordinates()
        Get the TimeStampedPVCoordinates 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 TimeStampedPVCoordinates, so the caller is responsible to copy it in a separate TimeStampedPVCoordinates if it needs to keep the value for a while.
        Returns:
        pvCoordinates in orbit definition frame

      • getPVCoordinates

        public TimeStampedPVCoordinates getPVCoordinates​(Frame outputFrame)
        Get the TimeStampedPVCoordinates 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 TimeStampedPVCoordinates, so the caller is responsible to copy it in a separate TimeStampedPVCoordinates 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 Attitude getAttitude()
        Get the attitude.
        Returns:
        the attitude.

      • getMass

        public double getMass()
        Gets the current mass.
        Returns:
        the mass (kg)