AngularRaDec.java
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* contributor license agreements. See the NOTICE file distributed with
* this work for additional information regarding copyright ownership.
* CS licenses this file to You under the Apache License, Version 2.0
* (the "License"); you may not use this file except in compliance with
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*
* http://www.apache.org/licenses/LICENSE-2.0
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* Unless required by applicable law or agreed to in writing, software
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package org.orekit.estimation.measurements;
import java.util.Arrays;
import java.util.Collections;
import java.util.HashMap;
import java.util.Map;
import org.hipparchus.analysis.differentiation.Gradient;
import org.hipparchus.analysis.differentiation.GradientField;
import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
import org.hipparchus.util.MathUtils;
import org.orekit.frames.FieldTransform;
import org.orekit.frames.Frame;
import org.orekit.propagation.SpacecraftState;
import org.orekit.time.AbsoluteDate;
import org.orekit.time.FieldAbsoluteDate;
import org.orekit.utils.ParameterDriver;
import org.orekit.utils.TimeStampedFieldPVCoordinates;
import org.orekit.utils.TimeStampedPVCoordinates;
/** Class modeling an Right Ascension - Declination measurement from a ground point (station, telescope).
* The angles are given in an inertial reference frame.
* The motion of the spacecraft during the signal flight time is taken into
* account. The date of the measurement corresponds to the reception on
* ground of the reflected signal.
*
* @author Thierry Ceolin
* @author Maxime Journot
* @since 9.0
*/
public class AngularRaDec extends AbstractMeasurement<AngularRaDec> {
/** Type of the measurement. */
public static final String MEASUREMENT_TYPE = "AngularRaDec";
/** Ground station from which measurement is performed. */
private final GroundStation station;
/** Reference frame in which the right ascension - declination angles are given. */
private final Frame referenceFrame;
/** Simple constructor.
* @param station ground station from which measurement is performed
* @param referenceFrame Reference frame in which the right ascension - declination angles are given
* @param date date of the measurement
* @param angular observed value
* @param sigma theoretical standard deviation
* @param baseWeight base weight
* @param satellite satellite related to this measurement
* @since 9.3
*/
public AngularRaDec(final GroundStation station, final Frame referenceFrame, final AbsoluteDate date,
final double[] angular, final double[] sigma, final double[] baseWeight,
final ObservableSatellite satellite) {
super(date, angular, sigma, baseWeight, Collections.singletonList(satellite));
addParameterDriver(station.getClockOffsetDriver());
addParameterDriver(station.getEastOffsetDriver());
addParameterDriver(station.getNorthOffsetDriver());
addParameterDriver(station.getZenithOffsetDriver());
addParameterDriver(station.getPrimeMeridianOffsetDriver());
addParameterDriver(station.getPrimeMeridianDriftDriver());
addParameterDriver(station.getPolarOffsetXDriver());
addParameterDriver(station.getPolarDriftXDriver());
addParameterDriver(station.getPolarOffsetYDriver());
addParameterDriver(station.getPolarDriftYDriver());
this.station = station;
this.referenceFrame = referenceFrame;
}
/** Get the ground station from which measurement is performed.
* @return ground station from which measurement is performed
*/
public GroundStation getStation() {
return station;
}
/** Get the reference frame in which the right ascension - declination angles are given.
* @return reference frame in which the right ascension - declination angles are given
*/
public Frame getReferenceFrame() {
return referenceFrame;
}
/** {@inheritDoc} */
@Override
protected EstimatedMeasurement<AngularRaDec> theoreticalEvaluation(final int iteration, final int evaluation,
final SpacecraftState[] states) {
final SpacecraftState state = states[0];
// Right Ascension/elevation (in reference frame )derivatives are computed with respect to spacecraft state in inertial frame
// and station parameters
// ----------------------
//
// Parameters:
// - 0..2 - Position of the spacecraft in inertial frame
// - 3..5 - Velocity of the spacecraft in inertial frame
// - 6..n - station parameters (clock offset, station offsets, pole, prime meridian...)
// Get the number of parameters used for derivation
// Place the selected drivers into a map
int nbParams = 6;
final Map<String, Integer> indices = new HashMap<>();
for (ParameterDriver driver : getParametersDrivers()) {
if (driver.isSelected()) {
indices.put(driver.getName(), nbParams++);
}
}
final FieldVector3D<Gradient> zero = FieldVector3D.getZero(GradientField.getField(nbParams));
// Coordinates of the spacecraft expressed as a gradient
final TimeStampedFieldPVCoordinates<Gradient> pvaDS = getCoordinates(state, 0, nbParams);
// Transform between station and inertial frame, expressed as a gradient
// The components of station's position in offset frame are the 3 last derivative parameters
final FieldTransform<Gradient> offsetToInertialDownlink =
station.getOffsetToInertial(state.getFrame(), getDate(), nbParams, indices);
final FieldAbsoluteDate<Gradient> downlinkDateDS =
offsetToInertialDownlink.getFieldDate();
// Station position/velocity in inertial frame at end of the downlink leg
final TimeStampedFieldPVCoordinates<Gradient> stationDownlink =
offsetToInertialDownlink.transformPVCoordinates(new TimeStampedFieldPVCoordinates<>(downlinkDateDS,
zero, zero, zero));
// Compute propagation times
// (if state has already been set up to pre-compensate propagation delay,
// we will have delta == tauD and transitState will be the same as state)
// Downlink delay
final Gradient tauD = signalTimeOfFlight(pvaDS, stationDownlink.getPosition(), downlinkDateDS);
// Transit state
final Gradient delta = downlinkDateDS.durationFrom(state.getDate());
final Gradient deltaMTauD = tauD.negate().add(delta);
final SpacecraftState transitState = state.shiftedBy(deltaMTauD.getValue());
// Transit state (re)computed with gradients
final TimeStampedFieldPVCoordinates<Gradient> transitStateDS = pvaDS.shiftedBy(deltaMTauD);
// Station-satellite vector expressed in inertial frame
final FieldVector3D<Gradient> staSatInertial = transitStateDS.getPosition().subtract(stationDownlink.getPosition());
// Field transform from inertial to reference frame at station's reception date
final FieldTransform<Gradient> inertialToReferenceDownlink =
state.getFrame().getTransformTo(referenceFrame, downlinkDateDS);
// Station-satellite vector in reference frame
final FieldVector3D<Gradient> staSatReference = inertialToReferenceDownlink.transformVector(staSatInertial);
// Compute right ascension and declination
final Gradient baseRightAscension = staSatReference.getAlpha();
final double twoPiWrap = MathUtils.normalizeAngle(baseRightAscension.getReal(),
getObservedValue()[0]) - baseRightAscension.getReal();
final Gradient rightAscension = baseRightAscension.add(twoPiWrap);
final Gradient declination = staSatReference.getDelta();
// Prepare the estimation
final EstimatedMeasurement<AngularRaDec> estimated =
new EstimatedMeasurement<>(this, iteration, evaluation,
new SpacecraftState[] {
transitState
}, new TimeStampedPVCoordinates[] {
transitStateDS.toTimeStampedPVCoordinates(),
stationDownlink.toTimeStampedPVCoordinates()
});
// azimuth - elevation values
estimated.setEstimatedValue(rightAscension.getValue(), declination.getValue());
// Partial derivatives of right ascension/declination in reference frame with respect to state
// (beware element at index 0 is the value, not a derivative)
final double[] raDerivatives = rightAscension.getGradient();
final double[] decDerivatives = declination.getGradient();
estimated.setStateDerivatives(0,
Arrays.copyOfRange(raDerivatives, 0, 6), Arrays.copyOfRange(decDerivatives, 0, 6));
// Partial derivatives with respect to parameters
// (beware element at index 0 is the value, not a derivative)
for (final ParameterDriver driver : getParametersDrivers()) {
final Integer index = indices.get(driver.getName());
if (index != null) {
estimated.setParameterDerivatives(driver, raDerivatives[index], decDerivatives[index]);
}
}
return estimated;
}
}