AngularRaDec.java
/* Copyright 2002-2018 CS Systèmes d'Information
* Licensed to CS Systèmes d'Information (CS) under one or more
* 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
* the License. You may obtain a copy of the License at
*
* http://www.apache.org/licenses/LICENSE-2.0
*
* Unless required by applicable law or agreed to in writing, software
* distributed under the License is distributed on an "AS IS" BASIS,
* WITHOUT WARRANTIES OR CONDITIONS OF ANY KIND, either express or implied.
* See the License for the specific language governing permissions and
* limitations under the License.
*/
package org.orekit.estimation.measurements;
import java.util.Arrays;
import java.util.HashMap;
import java.util.Map;
import org.hipparchus.Field;
import org.hipparchus.analysis.differentiation.DSFactory;
import org.hipparchus.analysis.differentiation.DerivativeStructure;
import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
import org.hipparchus.util.MathUtils;
import org.orekit.errors.OrekitException;
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> {
/** 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.
* <p>
* This constructor uses 0 as the index of the propagator related
* to this measurement, thus being well suited for mono-satellite
* orbit determination.
* </p>
* @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
* @exception OrekitException if a {@link org.orekit.utils.ParameterDriver}
* name conflict occurs
*/
public AngularRaDec(final GroundStation station, final Frame referenceFrame, final AbsoluteDate date,
final double[] angular, final double[] sigma, final double[] baseWeight)
throws OrekitException {
this(station, referenceFrame, date, angular, sigma, baseWeight, 0);
}
/** 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 propagatorIndex index of the propagator related to this measurement
* @exception OrekitException if a {@link org.orekit.utils.ParameterDriver}
* name conflict occurs
*/
public AngularRaDec(final GroundStation station, final Frame referenceFrame, final AbsoluteDate date,
final double[] angular, final double[] sigma, final double[] baseWeight,
final int propagatorIndex)
throws OrekitException {
super(date, angular, sigma, baseWeight, Arrays.asList(propagatorIndex),
station.getEastOffsetDriver(),
station.getNorthOffsetDriver(),
station.getZenithOffsetDriver(),
station.getPrimeMeridianOffsetDriver(),
station.getPrimeMeridianDriftDriver(),
station.getPolarOffsetXDriver(),
station.getPolarDriftXDriver(),
station.getPolarOffsetYDriver(),
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)
throws OrekitException {
final SpacecraftState state = states[getPropagatorsIndices().get(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 (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 DSFactory factory = new DSFactory(nbParams, 1);
final Field<DerivativeStructure> field = factory.getDerivativeField();
final FieldVector3D<DerivativeStructure> zero = FieldVector3D.getZero(field);
// Coordinates of the spacecraft expressed as a derivative structure
final TimeStampedFieldPVCoordinates<DerivativeStructure> pvaDS = getCoordinates(state, 0, factory);
// Transform between station and inertial frame, expressed as a derivative structure
// The components of station's position in offset frame are the 3 last derivative parameters
final AbsoluteDate downlinkDate = getDate();
final FieldAbsoluteDate<DerivativeStructure> downlinkDateDS =
new FieldAbsoluteDate<>(field, downlinkDate);
final FieldTransform<DerivativeStructure> offsetToInertialDownlink =
station.getOffsetToInertial(state.getFrame(), downlinkDateDS, factory, indices);
// Station position/velocity in inertial frame at end of the downlink leg
final TimeStampedFieldPVCoordinates<DerivativeStructure> 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 DerivativeStructure tauD = signalTimeOfFlight(pvaDS, stationDownlink.getPosition(), downlinkDateDS);
// Transit state
final double delta = downlinkDate.durationFrom(state.getDate());
final DerivativeStructure deltaMTauD = tauD.negate().add(delta);
final SpacecraftState transitState = state.shiftedBy(deltaMTauD.getValue());
// Transit state (re)computed with derivative structures
final TimeStampedFieldPVCoordinates<DerivativeStructure> transitStateDS = pvaDS.shiftedBy(deltaMTauD);
// Station-satellite vector expressed in inertial frame
final FieldVector3D<DerivativeStructure> staSatInertial = transitStateDS.getPosition().subtract(stationDownlink.getPosition());
// Field transform from inertial to reference frame at station's reception date
final FieldTransform<DerivativeStructure> inertialToReferenceDownlink =
state.getFrame().getTransformTo(referenceFrame, downlinkDateDS);
// Station-satellite vector in reference frame
final FieldVector3D<DerivativeStructure> staSatReference = inertialToReferenceDownlink.transformPosition(staSatInertial);
// Compute right ascension and declination
final DerivativeStructure baseRightAscension = staSatReference.getAlpha();
final double twoPiWrap = MathUtils.normalizeAngle(baseRightAscension.getReal(),
getObservedValue()[0]) - baseRightAscension.getReal();
final DerivativeStructure rightAscension = baseRightAscension.add(twoPiWrap);
final DerivativeStructure 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.getAllDerivatives();
final double[] decDerivatives = declination.getAllDerivatives();
estimated.setStateDerivatives(0,
Arrays.copyOfRange(raDerivatives, 1, 7), Arrays.copyOfRange(decDerivatives, 1, 7));
// 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 + 1], decDerivatives[index + 1]);
}
}
return estimated;
}
}