SBASPropagator.java
- /* Copyright 2002-2020 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,
- * 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.propagation.analytical.gnss;
- import org.hipparchus.analysis.differentiation.UnivariateDerivative2;
- import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
- import org.hipparchus.geometry.euclidean.threed.Vector3D;
- import org.orekit.attitudes.AttitudeProvider;
- import org.orekit.attitudes.InertialProvider;
- import org.orekit.errors.OrekitException;
- import org.orekit.errors.OrekitMessages;
- import org.orekit.frames.Frame;
- import org.orekit.frames.Frames;
- import org.orekit.orbits.CartesianOrbit;
- import org.orekit.orbits.Orbit;
- import org.orekit.propagation.SpacecraftState;
- import org.orekit.propagation.analytical.AbstractAnalyticalPropagator;
- import org.orekit.time.AbsoluteDate;
- import org.orekit.utils.IERSConventions;
- import org.orekit.utils.PVCoordinates;
- /**
- * This class aims at propagating a SBAS orbit from {@link SBASOrbitalElements}.
- *
- * @see "Tyler Reid, Todd Walker, Per Enge, L1/L5 SBAS MOPS Ephemeris Message to
- * Support Multiple Orbit Classes, ION ITM, 2013"
- *
- * @author Bryan Cazabonne
- * @since 10.1
- *
- */
- public class SBASPropagator extends AbstractAnalyticalPropagator {
- /** The SBAS orbital elements used. */
- private final SBASOrbitalElements sbasOrbit;
- /** The spacecraft mass (kg). */
- private final double mass;
- /** The Earth gravity coefficient used for SBAS propagation. */
- private final double mu;
- /** The ECI frame used for SBAS propagation. */
- private final Frame eci;
- /** The ECEF frame used for SBAS propagation. */
- private final Frame ecef;
- /**
- * This nested class aims at building a SBASPropagator.
- * <p>It implements the classical builder pattern.</p>
- *
- */
- public static class Builder {
- /** The SBAS orbital elements. */
- private final SBASOrbitalElements orbit;
- /** The Earth gravity coefficient used for SBAS propagation. */
- private double mu;
- /** The attitude provider. */
- private AttitudeProvider attitudeProvider;
- /** The mass. */
- private double mass = DEFAULT_MASS;
- /** The ECI frame. */
- private Frame eci = null;
- /** The ECEF frame. */
- private Frame ecef = null;
- /** Initializes the builder.
- * <p>The SBAS orbital elements is the only requested parameter to build a SBASPropagator.</p>
- * <p>The attitude provider is set by default to the
- * {@link org.orekit.propagation.Propagator#DEFAULT_LAW DEFAULT_LAW}.<br>
- * The Earth gravity coefficient is set by default to the
- * {@link org.orekit.propagation.analytical.gnss.SBASOrbitalElements#SBAS_MU SBAS_MU}.<br>
- * The mass is set by default to the
- * {@link org.orekit.propagation.Propagator#DEFAULT_MASS DEFAULT_MASS}.<br>
- * The ECI frame is set by default to the
- * {@link org.orekit.frames.Predefined#EME2000 EME2000 frame}.<br>
- * The ECEF frame is set by default to the
- * {@link org.orekit.frames.Predefined#ITRF_CIO_CONV_2010_SIMPLE_EOP CIO/2010-based ITRF simple EOP}.
- * </p>
- *
- * @param sbasOrbElt the SBAS orbital elements to be used by the SBAS propagator.
- * @param frames set of reference frames to use to initialize {@link
- * #ecef(Frame)}, {@link #eci(Frame)}, and {@link
- * #attitudeProvider(AttitudeProvider)}.
- * @see #attitudeProvider(AttitudeProvider provider)
- * @see #mu(double coefficient)
- * @see #mass(double mass)
- * @see #eci(Frame inertial)
- * @see #ecef(Frame bodyFixed)
- */
- public Builder(final SBASOrbitalElements sbasOrbElt, final Frames frames) {
- this.orbit = sbasOrbElt;
- this.eci = frames.getEME2000();
- this.ecef = frames.getITRF(IERSConventions.IERS_2010, true);
- this.mu = SBASOrbitalElements.SBAS_MU;
- this.attitudeProvider = new InertialProvider(eci);
- }
- /** Sets the attitude provider.
- *
- * @param userProvider the attitude provider
- * @return the updated builder
- */
- public Builder attitudeProvider(final AttitudeProvider userProvider) {
- this.attitudeProvider = userProvider;
- return this;
- }
- /** Sets the Earth gravity coefficient.
- *
- * @param coefficient the Earth gravity coefficient
- * @return the updated builder
- */
- public Builder mu(final double coefficient) {
- this.mu = coefficient;
- return this;
- }
- /** Sets the mass.
- *
- * @param userMass the mass (in kg)
- * @return the updated builder
- */
- public Builder mass(final double userMass) {
- this.mass = userMass;
- return this;
- }
- /** Sets the Earth Centered Inertial frame used for propagation.
- *
- * @param inertial the ECI frame
- * @return the updated builder
- */
- public Builder eci(final Frame inertial) {
- this.eci = inertial;
- return this;
- }
- /** Sets the Earth Centered Earth Fixed frame assimilated to the WGS84 ECEF.
- *
- * @param bodyFixed the ECEF frame
- * @return the updated builder
- */
- public Builder ecef(final Frame bodyFixed) {
- this.ecef = bodyFixed;
- return this;
- }
- /** Finalizes the build.
- *
- * @return the built SBASPropagator
- */
- public SBASPropagator build() {
- return new SBASPropagator(this);
- }
- }
- /**
- * Private constructor.
- * @param builder the builder
- */
- private SBASPropagator(final Builder builder) {
- super(builder.attitudeProvider);
- // Stores the SBAS orbital elements
- this.sbasOrbit = builder.orbit;
- // Sets the start date as the date of the orbital elements
- setStartDate(sbasOrbit.getDate());
- // Sets the mu
- this.mu = builder.mu;
- // Sets the mass
- this.mass = builder.mass;
- // Sets the Earth Centered Inertial frame
- this.eci = builder.eci;
- // Sets the Earth Centered Earth Fixed frame
- this.ecef = builder.ecef;
- }
- /**
- * Gets the PVCoordinates of the GNSS SV in {@link #getECEF() ECEF frame}.
- *
- * <p>The algorithm uses automatic differentiation to compute velocity and
- * acceleration.</p>
- *
- * @param date the computation date
- * @return the GNSS SV PVCoordinates in {@link #getECEF() ECEF frame}
- */
- public PVCoordinates propagateInEcef(final AbsoluteDate date) {
- // Duration from SBAS ephemeris Reference date
- final UnivariateDerivative2 dt = new UnivariateDerivative2( getDT(date), 1.0, 0.0);
- // Satellite coordinates
- final UnivariateDerivative2 x = dt.multiply(dt.multiply(0.5 * sbasOrbit.getXDotDot()).add(sbasOrbit.getXDot())).add(sbasOrbit.getX());
- final UnivariateDerivative2 y = dt.multiply(dt.multiply(0.5 * sbasOrbit.getYDotDot()).add(sbasOrbit.getYDot())).add(sbasOrbit.getY());
- final UnivariateDerivative2 z = dt.multiply(dt.multiply(0.5 * sbasOrbit.getZDotDot()).add(sbasOrbit.getZDot())).add(sbasOrbit.getZ());
- // Returns the Earth-fixed coordinates
- final FieldVector3D<UnivariateDerivative2> positionwithDerivatives =
- new FieldVector3D<>(x, y, z);
- return new PVCoordinates(new Vector3D(positionwithDerivatives.getX().getValue(),
- positionwithDerivatives.getY().getValue(),
- positionwithDerivatives.getZ().getValue()),
- new Vector3D(positionwithDerivatives.getX().getFirstDerivative(),
- positionwithDerivatives.getY().getFirstDerivative(),
- positionwithDerivatives.getZ().getFirstDerivative()),
- new Vector3D(positionwithDerivatives.getX().getSecondDerivative(),
- positionwithDerivatives.getY().getSecondDerivative(),
- positionwithDerivatives.getZ().getSecondDerivative()));
- }
- /** {@inheritDoc} */
- protected Orbit propagateOrbit(final AbsoluteDate date) {
- // Gets the PVCoordinates in ECEF frame
- final PVCoordinates pvaInECEF = propagateInEcef(date);
- // Transforms the PVCoordinates to ECI frame
- final PVCoordinates pvaInECI = ecef.getTransformTo(eci, date).transformPVCoordinates(pvaInECEF);
- // Returns the Cartesian orbit
- return new CartesianOrbit(pvaInECI, eci, date, mu);
- }
- /**
- * Get the Earth gravity coefficient used for SBAS propagation.
- * @return the Earth gravity coefficient.
- */
- public double getMU() {
- return mu;
- }
- /**
- * Gets the Earth Centered Inertial frame used to propagate the orbit.
- *
- * @return the ECI frame
- */
- public Frame getECI() {
- return eci;
- }
- /**
- * Gets the Earth Centered Earth Fixed frame used to propagate GNSS orbits.
- *
- * @return the ECEF frame
- */
- public Frame getECEF() {
- return ecef;
- }
- /**
- * Get the underlying SBAS orbital elements.
- *
- * @return the underlying SBAS orbital elements
- */
- public SBASOrbitalElements getSBASOrbitalElements() {
- return sbasOrbit;
- }
- /** {@inheritDoc} */
- public Frame getFrame() {
- return eci;
- }
- /** {@inheritDoc} */
- public void resetInitialState(final SpacecraftState state) {
- throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
- }
- /** {@inheritDoc} */
- protected double getMass(final AbsoluteDate date) {
- return mass;
- }
- /** {@inheritDoc} */
- protected void resetIntermediateState(final SpacecraftState state, final boolean forward) {
- throw new OrekitException(OrekitMessages.NON_RESETABLE_STATE);
- }
- /** Get the duration from SBAS Reference epoch.
- * @param date the considered date
- * @return the duration from SBAS orbit Reference epoch (s)
- */
- private double getDT(final AbsoluteDate date) {
- // Time from ephemeris reference epoch
- return date.durationFrom(sbasOrbit.getDate());
- }
- }