SBASPropagator.java
/* Copyright 2002-2020 CS GROUP
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* this work for additional information regarding copyright ownership.
* CS licenses this file to You under the Apache License, Version 2.0
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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.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());
}
}