Range.java
/* Copyright 2002-2021 CS GROUP
* Licensed to CS GROUP (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,
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* 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.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.orekit.frames.FieldTransform;
import org.orekit.propagation.SpacecraftState;
import org.orekit.time.AbsoluteDate;
import org.orekit.time.FieldAbsoluteDate;
import org.orekit.utils.Constants;
import org.orekit.utils.ParameterDriver;
import org.orekit.utils.TimeStampedFieldPVCoordinates;
import org.orekit.utils.TimeStampedPVCoordinates;
/** Class modeling a range measurement from a ground station.
* <p>
* For one-way measurements, a signal is emitted by the satellite
* and received by the ground station. The measurement value is the
* elapsed time between emission and reception multiplied by c where
* c is the speed of light.
* </p>
* <p>
* For two-way measurements, the measurement is considered to be a signal
* emitted from a ground station, reflected on spacecraft, and received
* on the same ground station. Its value is the elapsed time between
* emission and reception multiplied by c/2 where c is the speed of light.
* </p>
* <p>
* The motion of both the station and the spacecraft during the signal
* flight time are taken into account. The date of the measurement
* corresponds to the reception on ground of the emitted or reflected signal.
* </p>
* <p>
* The clock offsets of both the ground station and the satellite are taken
* into account. These offsets correspond to the values that must be subtracted
* from station (resp. satellite) reading of time to compute the real physical
* date. These offsets have two effects:
* </p>
* <ul>
* <li>as measurement date is evaluated at reception time, the real physical date
* of the measurement is the observed date to which the receiving ground station
* clock offset is subtracted</li>
* <li>as range is evaluated using the total signal time of flight, for one-way
* measurements the observed range is the real physical signal time of flight to
* which (Δtg - Δts) ⨉ c is added, where Δtg (resp. Δts) is the clock offset for the
* receiving ground station (resp. emitting satellite). A similar effect exists in
* two-way measurements but it is computed as (Δtg - Δtg) ⨉ c / 2 as the same ground
* station clock is used for initial emission and final reception and therefore it evaluates
* to zero.</li>
* </ul>
* <p>
* @author Thierry Ceolin
* @author Luc Maisonobe
* @author Maxime Journot
* @since 8.0
*/
public class Range extends AbstractMeasurement<Range> {
/** Ground station from which measurement is performed. */
private final GroundStation station;
/** Flag indicating whether it is a two-way measurement. */
private final boolean twoway;
/** Simple constructor.
* @param station ground station from which measurement is performed
* @param twoWay flag indicating whether it is a two-way measurement
* @param date date of the measurement
* @param range observed value
* @param sigma theoretical standard deviation
* @param baseWeight base weight
* @param satellite satellite related to this measurement
* @since 9.3
*/
public Range(final GroundStation station, final boolean twoWay, final AbsoluteDate date,
final double range, final double sigma, final double baseWeight,
final ObservableSatellite satellite) {
super(date, range, 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());
if (!twoWay) {
// for one way measurements, the satellite clock offset affects the measurement
addParameterDriver(satellite.getClockOffsetDriver());
}
this.station = station;
this.twoway = twoWay;
}
/** Get the ground station from which measurement is performed.
* @return ground station from which measurement is performed
*/
public GroundStation getStation() {
return station;
}
/** Check if the instance represents a two-way measurement.
* @return true if the instance represents a two-way measurement
*/
public boolean isTwoWay() {
return twoway;
}
/** {@inheritDoc} */
@Override
protected EstimatedMeasurement<Range> theoreticalEvaluation(final int iteration,
final int evaluation,
final SpacecraftState[] states) {
final SpacecraftState state = states[0];
// Range 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 - measurements parameters (clock offset, station offsets, pole, prime meridian, sat clock offset...)
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 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 & Transit state (re)computed with gradients
final Gradient delta = downlinkDateDS.durationFrom(state.getDate());
final Gradient deltaMTauD = tauD.negate().add(delta);
final SpacecraftState transitState = state.shiftedBy(deltaMTauD.getValue());
final TimeStampedFieldPVCoordinates<Gradient> transitStateDS = pvaDS.shiftedBy(deltaMTauD);
// prepare the evaluation
final EstimatedMeasurement<Range> estimated;
final Gradient range;
if (twoway) {
// Station at transit state date (derivatives of tauD taken into account)
final TimeStampedFieldPVCoordinates<Gradient> stationAtTransitDate =
stationDownlink.shiftedBy(tauD.negate());
// Uplink delay
final Gradient tauU =
signalTimeOfFlight(stationAtTransitDate, transitStateDS.getPosition(), transitStateDS.getDate());
final TimeStampedFieldPVCoordinates<Gradient> stationUplink =
stationDownlink.shiftedBy(-tauD.getValue() - tauU.getValue());
// Prepare the evaluation
estimated = new EstimatedMeasurement<Range>(this, iteration, evaluation,
new SpacecraftState[] {
transitState
}, new TimeStampedPVCoordinates[] {
stationUplink.toTimeStampedPVCoordinates(),
transitStateDS.toTimeStampedPVCoordinates(),
stationDownlink.toTimeStampedPVCoordinates()
});
// Range value
final double cOver2 = 0.5 * Constants.SPEED_OF_LIGHT;
final Gradient tau = tauD.add(tauU);
range = tau.multiply(cOver2);
} else {
estimated = new EstimatedMeasurement<Range>(this, iteration, evaluation,
new SpacecraftState[] {
transitState
}, new TimeStampedPVCoordinates[] {
transitStateDS.toTimeStampedPVCoordinates(),
stationDownlink.toTimeStampedPVCoordinates()
});
// Clock offsets
final ObservableSatellite satellite = getSatellites().get(0);
final Gradient dts = satellite.getClockOffsetDriver().getValue(nbParams, indices);
final Gradient dtg = station.getClockOffsetDriver().getValue(nbParams, indices);
// Range value
range = tauD.add(dtg).subtract(dts).multiply(Constants.SPEED_OF_LIGHT);
}
estimated.setEstimatedValue(range.getValue());
// Range partial derivatives with respect to state
final double[] derivatives = range.getGradient();
estimated.setStateDerivatives(0, Arrays.copyOfRange(derivatives, 0, 6));
// set 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, derivatives[index]);
}
}
return estimated;
}
}