ClockCorrectionsProvider.java

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package org.orekit.propagation.analytical.gnss;

import org.hipparchus.geometry.euclidean.threed.Vector3D;
import org.orekit.estimation.measurements.EstimationModifier;
import org.orekit.propagation.AdditionalStateProvider;
import org.orekit.propagation.SpacecraftState;
import org.orekit.propagation.analytical.gnss.data.GNSSClockElements;
import org.orekit.time.AbsoluteDate;
import org.orekit.utils.Constants;
import org.orekit.utils.PVCoordinates;

/** Provider for clock corrections as additional states.
 * <p>
 * The value of this additional state is a three elements array containing
 * </p>
 * <ul>
 *   <li>at index 0, the polynomial satellite clock model
 *       Δtₛₐₜ = {@link GNSSClockElements#getAf0() a₀} +
 *               {@link GNSSClockElements#getAf1() a₁} (t - {@link GNSSClockElements#getToc() toc}) +
 *               {@link GNSSClockElements#getAf1() a₂} (t - {@link GNSSClockElements#getToc() toc})²
 *   </li>
 *   <li>at index 1 the relativistic clock correction due to eccentricity</li>
 *   <li>at index 2 the estimated group delay differential {@link GNSSClockElements#getTGD() TGD} for L1-L2 correction</li>
 * </ul>
 * <p>
 * Since Orekit 10.3 the relativistic clock correction can be used as an {@link EstimationModifier}
 * in orbit determination applications to take into consideration this effect
 * in measurement modeling.
 * </p>
 *
 * @author Luc Maisonobe
 * @since 9.3
 */
public class ClockCorrectionsProvider implements AdditionalStateProvider {

    /** Name of the additional state for satellite clock corrections.
     * @since 9.3
     */
    public static final String CLOCK_CORRECTIONS = "";

    /** The GPS clock elements. */
    private final GNSSClockElements gnssClk;

    /** Clock reference epoch. */
    private final AbsoluteDate clockRef;

    /** Simple constructor.
     * @param gnssClk GNSS clock elements
     */
    public ClockCorrectionsProvider(final GNSSClockElements gnssClk) {
        this.gnssClk  = gnssClk;
        this.clockRef = gnssClk.getDate();
    }

    /** {@inheritDoc} */
    @Override
    public String getName() {
        return CLOCK_CORRECTIONS;
    }

    /**
     * Get the duration from clock Reference epoch.
     * <p>This takes the GNSS week roll-over into account.</p>
     *
     * @param date the considered date
     * @return the duration from clock Reference epoch (s)
     */
    private double getDT(final AbsoluteDate date) {
        final double cycleDuration = gnssClk.getCycleDuration();
        // Time from ephemeris reference epoch
        double dt = date.durationFrom(clockRef);
        // Adjusts the time to take roll over week into account
        while (dt > 0.5 * cycleDuration) {
            dt -= cycleDuration;
        }
        while (dt < -0.5 * cycleDuration) {
            dt += cycleDuration;
        }
        // Returns the time from ephemeris reference epoch
        return dt;
    }

    /** {@inheritDoc} */
    @Override
    public double[] getAdditionalState(final SpacecraftState state) {

        // polynomial clock model
        final double  dt    = getDT(state.getDate());
        final double  dtSat = gnssClk.getAf0() + dt * (gnssClk.getAf1() + dt * gnssClk.getAf2());

        // relativistic effect due to eccentricity
        final PVCoordinates pv    = state.getPVCoordinates();
        final double        dtRel = -2 * Vector3D.dotProduct(pv.getPosition(), pv.getVelocity()) /
                        (Constants.SPEED_OF_LIGHT * Constants.SPEED_OF_LIGHT);

        // estimated group delay differential
        final double tg = gnssClk.getTGD();

        return new double[] {
            dtSat, dtRel, tg
        };
    }

}