StateCovariance.java
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* this work for additional information regarding copyright ownership.
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*
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* Unless required by applicable law or agreed to in writing, software
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package org.orekit.propagation;
import org.hipparchus.linear.Array2DRowRealMatrix;
import org.hipparchus.linear.MatrixUtils;
import org.hipparchus.linear.RealMatrix;
import org.orekit.errors.OrekitException;
import org.orekit.errors.OrekitMessages;
import org.orekit.frames.Frame;
import org.orekit.frames.LOF;
import org.orekit.frames.Transform;
import org.orekit.orbits.Orbit;
import org.orekit.orbits.OrbitType;
import org.orekit.orbits.PositionAngleType;
import org.orekit.time.AbsoluteDate;
import org.orekit.time.TimeStamped;
import org.orekit.utils.CartesianDerivativesFilter;
/** This class is the representation of a covariance matrix at a given date.
* <p>
* Currently, the covariance only represents the orbital elements.
* <p>
* It is possible to change the covariance frame by using the
* {@link #changeCovarianceFrame(Orbit, Frame)} or {@link #changeCovarianceFrame(Orbit, LOF)} method.
* These methods are based on Equations (18) and (20) of <i>Covariance Transformations for Satellite
* Flight Dynamics Operations</i> by David A. SVallado.
* <p>
* Finally, covariance orbit type can be changed using the
* {@link #changeCovarianceType(Orbit, OrbitType, PositionAngleType)} method.
*
* @author Bryan Cazabonne
* @author Vincent Cucchietti
* @since 11.3
*/
public class StateCovariance implements TimeStamped {
/** State dimension. */
public static final int STATE_DIMENSION = 6;
/** Default position angle for covariance expressed in Cartesian elements. */
private static final PositionAngleType DEFAULT_POSITION_ANGLE = PositionAngleType.TRUE;
/** Orbital covariance [6x6]. */
private final RealMatrix orbitalCovariance;
/** Covariance frame (can be null if LOF is defined). */
private final Frame frame;
/** Covariance LOF type (can be null if frame is defined). */
private final LOF lof;
/** Covariance epoch. */
private final AbsoluteDate epoch;
/** Covariance orbit type. */
private final OrbitType orbitType;
/** Covariance position angle type (not used if orbitType is CARTESIAN). */
private final PositionAngleType angleType;
/**
* Constructor.
* @param orbitalCovariance 6x6 orbital parameters covariance
* @param epoch epoch of the covariance
* @param lof covariance LOF type
*/
public StateCovariance(final RealMatrix orbitalCovariance, final AbsoluteDate epoch, final LOF lof) {
this(orbitalCovariance, epoch, null, lof, OrbitType.CARTESIAN, DEFAULT_POSITION_ANGLE);
}
/**
* Constructor.
* @param orbitalCovariance 6x6 orbital parameters covariance
* @param epoch epoch of the covariance
* @param covarianceFrame covariance frame (inertial or Earth fixed)
* @param orbitType orbit type of the covariance (CARTESIAN if covarianceFrame is not pseudo-inertial)
* @param angleType position angle type of the covariance (not used if orbitType is CARTESIAN)
*/
public StateCovariance(final RealMatrix orbitalCovariance, final AbsoluteDate epoch,
final Frame covarianceFrame,
final OrbitType orbitType, final PositionAngleType angleType) {
this(orbitalCovariance, epoch, covarianceFrame, null, orbitType, angleType);
}
/**
* Private constructor.
* @param orbitalCovariance 6x6 orbital parameters covariance
* @param epoch epoch of the covariance
* @param covarianceFrame covariance frame (inertial or Earth fixed)
* @param lof covariance LOF type
* @param orbitType orbit type of the covariance
* @param angleType position angle type of the covariance (not used if orbitType is CARTESIAN)
*/
private StateCovariance(final RealMatrix orbitalCovariance, final AbsoluteDate epoch,
final Frame covarianceFrame, final LOF lof,
final OrbitType orbitType, final PositionAngleType angleType) {
checkFrameAndOrbitTypeConsistency(covarianceFrame, orbitType);
this.orbitalCovariance = orbitalCovariance;
this.epoch = epoch;
this.frame = covarianceFrame;
this.lof = lof;
this.orbitType = orbitType;
this.angleType = angleType;
}
/**
* Check constructor's inputs consistency.
*
* @param covarianceFrame covariance frame (inertial or Earth fixed)
* @param inputType orbit type of the covariance
*
* @throws OrekitException if input frame is not pseudo-inertial AND the orbit type is not Cartesian
*/
public static void checkFrameAndOrbitTypeConsistency(final Frame covarianceFrame, final OrbitType inputType) {
// State covariance expressed in a celestial body frame
if (covarianceFrame != null) {
// Input frame is not pseudo-inertial
if (!covarianceFrame.isPseudoInertial() && inputType != OrbitType.CARTESIAN) {
throw new OrekitException(OrekitMessages.WRONG_ORBIT_PARAMETERS_TYPE,
inputType.name(),
OrbitType.CARTESIAN.name());
}
}
}
/** {@inheritDoc}. */
@Override
public AbsoluteDate getDate() {
return epoch;
}
/**
* Get the covariance matrix.
* @return the covariance matrix
*/
public RealMatrix getMatrix() {
return orbitalCovariance;
}
/**
* Get the covariance orbit type.
* @return the covariance orbit type
*/
public OrbitType getOrbitType() {
return orbitType;
}
/**
* Get the covariance angle type.
* @return the covariance angle type
*/
public PositionAngleType getPositionAngleType() {
return angleType;
}
/**
* Get the covariance frame.
* @return the covariance frame (can be null)
* @see #getLOF()
*/
public Frame getFrame() {
return frame;
}
/**
* Get the covariance LOF type.
* @return the covariance LOF type (can be null)
* @see #getFrame()
*/
public LOF getLOF() {
return lof;
}
/**
* Get the covariance matrix in another orbit type.
* <p>
* The covariance orbit type <b>cannot</b> be changed if the covariance
* matrix is expressed in a {@link LOF local orbital frame} or a
* non-pseudo inertial frame.
* <p>
* As this type change uses the jacobian matrix of the transformation, it introduces a linear approximation.
* Hence, the current covariance matrix <b>will not exactly match</b> the new linearized case and the
* distribution will not follow a generalized Gaussian distribution anymore.
* <p>
* This is based on equation (1) to (6) from "Vallado, D. A. (2004). Covariance transformations for satellite flight
* dynamics operations."
*
* @param orbit orbit to which the covariance matrix should correspond
* @param outOrbitType target orbit type of the state covariance matrix
* @param outAngleType target position angle type of the state covariance matrix
* @return a new covariance state, expressed in the target orbit type with the target position angle
* @see #changeCovarianceFrame(Orbit, Frame)
*/
public StateCovariance changeCovarianceType(final Orbit orbit, final OrbitType outOrbitType,
final PositionAngleType outAngleType) {
// Handle case where the covariance is already expressed in the output type
if (outOrbitType == orbitType && (outAngleType == angleType || outOrbitType == OrbitType.CARTESIAN)) {
if (lof == null) {
return new StateCovariance(orbitalCovariance, epoch, frame, orbitType, angleType);
}
else {
return new StateCovariance(orbitalCovariance, epoch, lof);
}
}
// Check if the covariance is expressed in a celestial body frame
if (frame != null) {
// Check if the covariance is defined in an inertial frame
if (frame.isPseudoInertial()) {
return changeTypeAndCreate(orbit, epoch, frame, orbitType, angleType, outOrbitType, outAngleType,
orbitalCovariance);
}
// The covariance is not defined in an inertial frame. The orbit type cannot be changed
throw new OrekitException(OrekitMessages.CANNOT_CHANGE_COVARIANCE_TYPE_IF_DEFINED_IN_NON_INERTIAL_FRAME);
}
// The covariance is not expressed in a celestial body frame. The orbit type cannot be changed
throw new OrekitException(OrekitMessages.CANNOT_CHANGE_COVARIANCE_TYPE_IF_DEFINED_IN_LOF);
}
/**
* Get the covariance in a given local orbital frame.
* <p>
* Changing the covariance frame is a linear process, this method does not introduce approximation unless a change
* in covariance orbit type is required.
* <p>
* This is based on equation (18) to (20) "from Vallado, D. A. (2004). Covariance transformations for satellite
* flight dynamics operations."
*
* @param orbit orbit to which the covariance matrix should correspond
* @param lofOut output local orbital frame
* @return a new covariance state, expressed in the output local orbital frame
*/
public StateCovariance changeCovarianceFrame(final Orbit orbit, final LOF lofOut) {
// Verify current covariance frame
if (lof != null) {
// Change the covariance local orbital frame
return changeFrameAndCreate(orbit, epoch, lof, lofOut, orbitalCovariance);
} else {
// Covariance is expressed in celestial body frame
return changeFrameAndCreate(orbit, epoch, frame, lofOut, orbitalCovariance, orbitType, angleType);
}
}
/**
* Get the covariance in the output frame.
* <p>
* Changing the covariance frame is a linear process, this method does not introduce approximation unless a change
* in covariance orbit type is required.
* <p>
* This is based on equation (18) to (20) "from Vallado, D. A. (2004). Covariance transformations for satellite
* flight dynamics operations."
*
* @param orbit orbit to which the covariance matrix should correspond
* @param frameOut output frame
* @return a new covariance state, expressed in the output frame
*/
public StateCovariance changeCovarianceFrame(final Orbit orbit, final Frame frameOut) {
// Verify current covariance frame
if (lof != null) {
// Covariance is expressed in local orbital frame
return changeFrameAndCreate(orbit, epoch, lof, frameOut, orbitalCovariance);
} else {
// Change covariance frame
return changeFrameAndCreate(orbit, epoch, frame, frameOut, orbitalCovariance, orbitType, angleType);
}
}
/**
* Get a time-shifted covariance matrix.
* <p>
* The shifting model is a linearized, Keplerian one. In other words, it is based on a state transition matrix that
* is computed assuming Keplerian motion.
* <p>
* Shifting is <em>not</em> intended as a replacement for proper covariance propagation, but should be sufficient
* for small time shifts or coarse accuracy.
*
* @param orbit orbit to which the covariance matrix should correspond
* @param dt time shift in seconds
* @return a new covariance state, shifted with respect to the instance
*/
public StateCovariance shiftedBy(final Orbit orbit, final double dt) {
// Shifted orbit
final Orbit shifted = orbit.shiftedBy(dt);
// State covariance expressed in celestial body frame
if (frame != null) {
// State covariance expressed in a pseudo-inertial frame
if (frame.isPseudoInertial()) {
// Compute STM
final RealMatrix stm = getStm(orbit, dt);
// Convert covariance in STM type (i.e., Equinoctial elements)
final StateCovariance inStmType = changeTypeAndCreate(orbit, epoch, frame, orbitType, angleType,
OrbitType.EQUINOCTIAL, PositionAngleType.MEAN,
orbitalCovariance);
// Shift covariance by applying the STM
final RealMatrix shiftedCov = stm.multiply(inStmType.getMatrix().multiplyTransposed(stm));
// Restore the initial covariance type
return changeTypeAndCreate(shifted, shifted.getDate(), frame,
OrbitType.EQUINOCTIAL, PositionAngleType.MEAN,
orbitType, angleType, shiftedCov);
}
// State covariance expressed in a non pseudo-inertial frame
else {
// Convert state covariance to orbit pseudo-inertial frame
final StateCovariance inOrbitFrame = changeCovarianceFrame(orbit, orbit.getFrame());
// Shift the state covariance
final StateCovariance shiftedCovariance = inOrbitFrame.shiftedBy(orbit, dt);
// Restore the initial covariance frame
return shiftedCovariance.changeCovarianceFrame(shifted, frame);
}
}
// State covariance expressed in a commonly used local orbital frame (LOF)
else {
// Convert state covariance to orbit pseudo-inertial frame
final StateCovariance inOrbitFrame = changeCovarianceFrame(orbit, orbit.getFrame());
// Shift the state covariance
final StateCovariance shiftedCovariance = inOrbitFrame.shiftedBy(orbit, dt);
// Restore the initial covariance frame
return shiftedCovariance.changeCovarianceFrame(shifted, lof);
}
}
/**
* Create a covariance matrix in another orbit type.
* <p>
* As this type change uses the jacobian matrix of the transformation, it introduces a linear approximation.
* Hence, the input covariance matrix <b>will not exactly match</b> the new linearized case and the
* distribution will not follow a generalized Gaussian distribution anymore.
* <p>
* This is based on equation (1) to (6) from "Vallado, D. A. (2004). Covariance transformations for satellite flight
* dynamics operations."
*
* @param orbit orbit to which the covariance matrix should correspond
* @param date covariance epoch
* @param covFrame covariance frame
* @param inOrbitType initial orbit type of the state covariance matrix
* @param inAngleType initial position angle type of the state covariance matrix
* @param outOrbitType target orbit type of the state covariance matrix
* @param outAngleType target position angle type of the state covariance matrix
* @param inputCov input covariance
* @return the covariance expressed in the target orbit type with the target position angle
*/
private static StateCovariance changeTypeAndCreate(final Orbit orbit, final AbsoluteDate date,
final Frame covFrame,
final OrbitType inOrbitType, final PositionAngleType inAngleType,
final OrbitType outOrbitType, final PositionAngleType outAngleType,
final RealMatrix inputCov) {
// Notations:
// I: Input orbit type
// O: Output orbit type
// C: Cartesian parameters
// Compute dOutputdCartesian
final double[][] aOC = new double[STATE_DIMENSION][STATE_DIMENSION];
final Orbit orbitInOutputType = outOrbitType.convertType(orbit);
orbitInOutputType.getJacobianWrtCartesian(outAngleType, aOC);
final RealMatrix dOdC = new Array2DRowRealMatrix(aOC, false);
// Compute dCartesiandInput
final double[][] aCI = new double[STATE_DIMENSION][STATE_DIMENSION];
final Orbit orbitInInputType = inOrbitType.convertType(orbit);
orbitInInputType.getJacobianWrtParameters(inAngleType, aCI);
final RealMatrix dCdI = new Array2DRowRealMatrix(aCI, false);
// Compute dOutputdInput
final RealMatrix dOdI = dOdC.multiply(dCdI);
// Conversion of the state covariance matrix in target orbit type
final RealMatrix outputCov = dOdI.multiply(inputCov.multiplyTransposed(dOdI));
// Return the converted covariance
return new StateCovariance(outputCov, date, covFrame, outOrbitType, outAngleType);
}
/**
* Create a covariance matrix from a {@link LOF local orbital frame} to another
* {@link LOF local orbital frame}.
* <p>
* Changing the covariance frame is a linear process, this method does not introduce approximation.
* <p>
* The transformation is based on Equation (18) to (20) of "Covariance Transformations for Satellite Flight Dynamics
* Operations" by David A. Vallado.
*
* @param orbit orbit to which the covariance matrix should correspond
* @param date covariance epoch
* @param lofIn the local orbital frame in which the input covariance matrix is expressed
* @param lofOut the target local orbital frame
* @param inputCartesianCov input covariance {@code CARTESIAN})
* @return the covariance matrix expressed in the target commonly used local orbital frame in Cartesian elements
*/
private static StateCovariance changeFrameAndCreate(final Orbit orbit, final AbsoluteDate date,
final LOF lofIn, final LOF lofOut,
final RealMatrix inputCartesianCov) {
// Builds the matrix to perform covariance transformation
final RealMatrix jacobianFromLofInToLofOut =
getJacobian(LOF.transformFromLOFInToLOFOut(lofIn, lofOut, date, orbit.getPVCoordinates()));
// Get the Cartesian covariance matrix converted to frameOut
final RealMatrix cartesianCovarianceOut =
jacobianFromLofInToLofOut.multiply(inputCartesianCov.multiplyTransposed(jacobianFromLofInToLofOut));
// Output converted covariance
return new StateCovariance(cartesianCovarianceOut, date, lofOut);
}
/**
* Convert the covariance matrix from a {@link Frame frame} to a {@link LOF local orbital frame}.
* <p>
* Changing the covariance frame is a linear process, this method does not introduce approximation unless a change
* in covariance orbit type is required.
* <p>
* The transformation is based on Equation (18) to (20) of "Covariance Transformations for Satellite Flight Dynamics
* Operations" by David A. Vallado.
* <p>
* As the frame transformation must be performed with the covariance expressed in Cartesian elements, both the orbit
* and position angle types of the input covariance must be provided.
* <p>
* <b>The output covariance matrix will provided in Cartesian orbit type and not converted back to
* its original expression (if input different from Cartesian elements).</b>
*
* @param orbit orbit to which the covariance matrix should correspond
* @param date covariance epoch
* @param frameIn the frame in which the input covariance matrix is expressed. In case the frame is <b>not</b>
* pseudo-inertial, the input covariance matrix is expected to be expressed in <b>Cartesian elements</b>.
* @param lofOut the target local orbital frame
* @param inputCov input covariance
* @param covOrbitType orbit type of the covariance matrix (used if frameIn is pseudo-inertial)
* @param covAngleType position angle type of the covariance matrix (used if frameIn is pseudo-inertial) (not used
* if covOrbitType equals {@code CARTESIAN})
* @return the covariance matrix expressed in the target local orbital frame in Cartesian elements
* @throws OrekitException if input frame is <b>not</b> pseudo-inertial <b>and</b> the input covariance is
* <b>not</b> expressed in Cartesian elements.
*/
private static StateCovariance changeFrameAndCreate(final Orbit orbit, final AbsoluteDate date,
final Frame frameIn, final LOF lofOut,
final RealMatrix inputCov,
final OrbitType covOrbitType,
final PositionAngleType covAngleType) {
// Input frame is inertial
if (frameIn.isPseudoInertial()) {
// Convert input matrix to Cartesian parameters in input frame
final RealMatrix cartesianCovarianceIn =
changeTypeAndCreate(orbit, date, frameIn, covOrbitType, covAngleType,
OrbitType.CARTESIAN, PositionAngleType.MEAN,
inputCov).getMatrix();
// Builds the matrix to perform covariance transformation
final RealMatrix jacobianFromFrameInToLofOut =
getJacobian(lofOut.transformFromInertial(date, orbit.getPVCoordinates(frameIn)));
// Get the Cartesian covariance matrix converted to frameOut
final RealMatrix cartesianCovarianceOut =
jacobianFromFrameInToLofOut.multiply(cartesianCovarianceIn.multiplyTransposed(jacobianFromFrameInToLofOut));
// Return converted covariance matrix expressed in Cartesian elements
return new StateCovariance(cartesianCovarianceOut, date, lofOut);
}
// Input frame is not inertial so the covariance matrix is expected to be in Cartesian elements
else {
// Method checkInputConsistency already verify that input covariance is defined in CARTESIAN type
final Frame orbitInertialFrame = orbit.getFrame();
// Compute rotation matrix from frameIn to orbit inertial frame
final RealMatrix cartesianCovarianceInOrbitFrame =
changeFrameAndCreate(orbit, date, frameIn, orbitInertialFrame, inputCov,
OrbitType.CARTESIAN, PositionAngleType.MEAN).getMatrix();
// Convert from orbit inertial frame to lofOut
return changeFrameAndCreate(orbit, date, orbitInertialFrame, lofOut, cartesianCovarianceInOrbitFrame,
OrbitType.CARTESIAN, PositionAngleType.MEAN);
}
}
/**
* Convert the covariance matrix from a {@link LOF local orbital frame} to a {@link Frame frame}.
* <p>
* Changing the covariance frame is a linear process, this method does not introduce approximation.
* <p>
* The transformation is based on Equation (18) to (20) of "Covariance Transformations for Satellite Flight Dynamics
* Operations" by David A. Vallado.
* <p>
* The <u>input</u> covariance matrix is necessarily provided in <b>Cartesian orbit type</b>.
* <p>
* The <u>output</u> covariance matrix will be expressed in <b>Cartesian elements</b>.
*
* @param orbit orbit to which the covariance matrix should correspond
* @param date covariance epoch
* @param lofIn the local orbital frame in which the input covariance matrix is expressed
* @param frameOut the target frame
* @param inputCartesianCov input covariance ({@code CARTESIAN})
* @return the covariance matrix expressed in the target frame in Cartesian elements
*/
private static StateCovariance changeFrameAndCreate(final Orbit orbit, final AbsoluteDate date,
final LOF lofIn, final Frame frameOut,
final RealMatrix inputCartesianCov) {
// Output frame is pseudo-inertial
if (frameOut.isPseudoInertial()) {
// Builds the matrix to perform covariance transformation
final RealMatrix jacobianFromLofInToFrameOut =
getJacobian(lofIn.transformFromInertial(date, orbit.getPVCoordinates(frameOut)).getInverse());
// Transform covariance
final RealMatrix transformedCovariance =
jacobianFromLofInToFrameOut.multiply(inputCartesianCov.multiplyTransposed(jacobianFromLofInToFrameOut));
// Get the Cartesian covariance matrix converted to frameOut
return new StateCovariance(transformedCovariance, date, frameOut, OrbitType.CARTESIAN,
DEFAULT_POSITION_ANGLE);
}
// Output frame is not pseudo-inertial
else {
// Builds the matrix to perform covariance transformation
final RealMatrix jacobianFromLofInToOrbitFrame =
getJacobian(lofIn.transformFromInertial(date, orbit.getPVCoordinates()).getInverse());
// Get the Cartesian covariance matrix converted to orbit inertial frame
final RealMatrix cartesianCovarianceInOrbitFrame = jacobianFromLofInToOrbitFrame.multiply(
inputCartesianCov.multiplyTransposed(jacobianFromLofInToOrbitFrame));
// Get the Cartesian covariance matrix converted to frameOut
return changeFrameAndCreate(orbit, date, orbit.getFrame(), frameOut, cartesianCovarianceInOrbitFrame,
OrbitType.CARTESIAN, PositionAngleType.MEAN);
}
}
/**
* Get the covariance matrix in another frame.
* <p>
* The transformation is based on Equation (18) of "Covariance Transformations for Satellite Flight Dynamics
* Operations" by David A. Vallado.
* <p>
* Changing the covariance frame is a linear process, this method does not introduce approximation unless a change
* in covariance orbit type is required.
* <p>
* As the frame transformation must be performed with the covariance expressed in Cartesian elements, both the orbit
* and position angle types of the input covariance must be provided.
* <p>
* In case the <u>input</u> frame is <b>not</b> pseudo-inertial, the <u>input</u> covariance matrix is necessarily
* expressed in <b>Cartesian elements</b>.
* <p>
* In case the <u>output</u> frame is <b>not</b> pseudo-inertial, the <u>output</u> covariance matrix will be
* expressed in <b>Cartesian elements</b>.
*
* @param orbit orbit to which the covariance matrix should correspond
* @param date covariance epoch
* @param frameIn the frame in which the input covariance matrix is expressed
* @param frameOut the target frame
* @param inputCov input covariance
* @param covOrbitType orbit type of the covariance matrix (used if frameIn is pseudo-inertial)
* @param covAngleType position angle type of the covariance matrix (used if frameIn is pseudo-inertial) (<b>not</b>
* used if covOrbitType equals {@code CARTESIAN})
* @return the covariance matrix expressed in the target frame
* @throws OrekitException if input frame is <b>not</b> pseudo-inertial <b>and</b> the input covariance is
* <b>not</b> expressed in Cartesian elements.
*/
private static StateCovariance changeFrameAndCreate(final Orbit orbit, final AbsoluteDate date,
final Frame frameIn, final Frame frameOut,
final RealMatrix inputCov,
final OrbitType covOrbitType,
final PositionAngleType covAngleType) {
// Get the transform from the covariance frame to the output frame
final Transform inToOut = frameIn.getTransformTo(frameOut, orbit.getDate());
// Matrix to perform the covariance transformation
final RealMatrix j = getJacobian(inToOut);
// Input frame pseudo-inertial
if (frameIn.isPseudoInertial()) {
// Convert input matrix to Cartesian parameters in input frame
final RealMatrix cartesianCovarianceIn =
changeTypeAndCreate(orbit, date, frameIn, covOrbitType, covAngleType,
OrbitType.CARTESIAN, PositionAngleType.MEAN,
inputCov).getMatrix();
// Get the Cartesian covariance matrix converted to frameOut
final RealMatrix cartesianCovarianceOut = j.multiply(cartesianCovarianceIn.multiplyTransposed(j));
// Output frame is pseudo-inertial
if (frameOut.isPseudoInertial()) {
// Convert output Cartesian matrix to initial orbit type and position angle
return changeTypeAndCreate(orbit, date, frameOut, OrbitType.CARTESIAN, PositionAngleType.MEAN,
covOrbitType, covAngleType, cartesianCovarianceOut);
}
// Output frame is not pseudo-inertial
else {
// Output Cartesian matrix
return new StateCovariance(cartesianCovarianceOut, date, frameOut, OrbitType.CARTESIAN,
DEFAULT_POSITION_ANGLE);
}
}
// Input frame is not pseudo-inertial
else {
// Method checkInputConsistency already verify that input covariance is defined in CARTESIAN type
// Convert covariance matrix to frameOut
final RealMatrix covInFrameOut = j.multiply(inputCov.multiplyTransposed(j));
// Output the Cartesian covariance matrix converted to frameOut
return new StateCovariance(covInFrameOut, date, frameOut, OrbitType.CARTESIAN, DEFAULT_POSITION_ANGLE);
}
}
/**
* Builds the matrix to perform covariance frame transformation.
* @param transform input transformation
* @return the matrix to perform the covariance frame transformation
*/
private static RealMatrix getJacobian(final Transform transform) {
// Get the Jacobian of the transform
final double[][] jacobian = new double[STATE_DIMENSION][STATE_DIMENSION];
transform.getJacobian(CartesianDerivativesFilter.USE_PV, jacobian);
// Return
return new Array2DRowRealMatrix(jacobian, false);
}
/**
* Get the state transition matrix considering Keplerian contribution only.
*
* @param initialOrbit orbit to which the initial covariance matrix should correspond
* @param dt time difference between the two orbits
* @return the state transition matrix used to shift the covariance matrix
*/
public static RealMatrix getStm(final Orbit initialOrbit, final double dt) {
// initialize the STM
final RealMatrix stm = MatrixUtils.createRealIdentityMatrix(STATE_DIMENSION);
// State transition matrix using Keplerian contribution only
final double contribution = initialOrbit.getMeanAnomalyDotWrtA() * dt;
stm.setEntry(5, 0, contribution);
// Return
return stm;
}
}