ECOM2.java
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package org.orekit.forces.radiation;
import java.util.ArrayList;
import java.util.Collections;
import java.util.List;
import org.hipparchus.CalculusFieldElement;
import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
import org.hipparchus.geometry.euclidean.threed.Vector3D;
import org.hipparchus.util.FastMath;
import org.hipparchus.util.FieldSinCos;
import org.hipparchus.util.SinCos;
import org.orekit.annotation.DefaultDataContext;
import org.orekit.bodies.OneAxisEllipsoid;
import org.orekit.frames.FramesFactory;
import org.orekit.propagation.FieldSpacecraftState;
import org.orekit.propagation.SpacecraftState;
import org.orekit.utils.ExtendedPositionProvider;
import org.orekit.utils.ParameterDriver;
/**
* The Empirical CODE Orbit Model 2 (ECOM2) of the Center for Orbit Determination in Europe (CODE).
* <p>
* The drag acceleration is computed as follows :
* γ = γ<sub>0</sub> + D(u)e<sub>D</sub> + Y(u)e<sub>Y</sub> + B(u)e<sub>B</sub>
* </p> <p>
* In the above equation, γ<sub>0</sub> is a selectable a priori model. Since 2013, no
* a priori model is used for CODE IGS contribution (i.e. γ<sub>0</sub> = 0). Moreover,
* u denotes the satellite's argument of latitude.
* </p> <p>
* D(u), Y(u) and B(u) are three functions of the ECOM2 model that can be represented
* as Fourier series. The coefficients of the Fourier series are estimated during the
* estimation process. he ECOM2 model has user-defines upper limits <i>nD</i> and
* <i>nB</i> for the Fourier series (i.e. <i>nD</i> for D(u) and <i>nB</i> for
* B(u). Y(u) is defined as a constant value).
* </p> <p>
* It exists several configurations to initialize <i>nD</i> and <i>nB</i> values. However,
* Arnold et al recommend to use <b>D2B1</b> (i.e. <i>nD</i> = 1 and <i>nB</i> = 1) and
* <b>D4B1</b> (i.e. <i>nD</i> = 2 an <i>nB</i> = 1) configurations. At the opposite, in Arnold paper, it
* is recommend to not use <b>D2B0</b> (i.e. <i>nD</i> = 1 and <i>nB</i> = 0) configuration.
* </p> <p>
* Since Orekit 11.0, it is possible to take into account
* the eclipses generated by Moon in the solar radiation
* pressure force model using the
* {@link #addOccultingBody(ExtendedPositionProvider, double)}
* method.<br>
* <code> ECOM2 srp =</code>
* <code> new ECOM2(1, 1, 0.0, CelestialBodyFactory.getSun(), Constants.EIGEN5C_EARTH_EQUATORIAL_RADIUS);</code><br>
* <code> srp.addOccultingBody(CelestialBodyFactory.getMoon(), Constants.MOON_EQUATORIAL_RADIUS);</code><br>
*
* @see "Arnold, Daniel, et al, CODE’s new solar radiation pressure model for GNSS orbit determination,
* Journal of geodesy 89.8 (2015): 775-791."
*
* @see "Tzu-Pang tseng and Michael Moore, Impact of solar radiation pressure mis-modeling on
* GNSS satellite orbit determination, IGS Worshop, Wuhan, China, 2018."
*
* @author David Soulard
* @since 10.2
*/
public class ECOM2 extends AbstractRadiationForceModel {
/** Parameter name for ECOM model coefficients enabling Jacobian processing. */
public static final String ECOM_COEFFICIENT = "ECOM coefficient";
/** Minimum value for ECOM2 estimated parameters. */
private static final double MIN_VALUE = Double.NEGATIVE_INFINITY;
/** Maximum value for ECOM2 estimated parameters. */
private static final double MAX_VALUE = Double.POSITIVE_INFINITY;
/** Parameters scaling factor.
* <p>
* We use a power of 2 to avoid numeric noise introduction
* in the multiplications/divisions sequences.
* </p>
*/
private final double SCALE = FastMath.scalb(1.0, -22);
/** Highest order for parameter along eD axis (satellite --> sun direction). */
private final int nD;
/** Highest order for parameter along eB axis. */
private final int nB;
/** Estimated acceleration coefficients.
* <p>
* The 2 * nD first driver are Fourier driver along eD, axis,
* then along eY, then 2*nB following are along eB axis.
* </p>
*/
private final List<ParameterDriver> coefficients;
/** Sun model. */
private final ExtendedPositionProvider sun;
/**
* Constructor.
* @param nD truncation rank of Fourier series in D term.
* @param nB truncation rank of Fourier series in B term.
* @param value parameters initial value
* @param sun provide for Sun parameter
* @param equatorialRadius spherical shape model (for umbra/penumbra computation)
*/
@DefaultDataContext
public ECOM2(final int nD, final int nB, final double value,
final ExtendedPositionProvider sun, final double equatorialRadius) {
super(sun, new OneAxisEllipsoid(equatorialRadius, 0.0, FramesFactory.getGCRF()));
this.nB = nB;
this.nD = nD;
this.coefficients = new ArrayList<>(2 * (nD + nB) + 3);
// Add parameter along eB axis in alphabetical order
coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " B0", value, SCALE, MIN_VALUE, MAX_VALUE));
for (int i = 1; i < nB + 1; i++) {
coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Bcos" + Integer.toString(i - 1), value, SCALE, MIN_VALUE, MAX_VALUE));
}
for (int i = nB + 1; i < 2 * nB + 1; i++) {
coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Bsin" + Integer.toString(i - (nB + 1)), value, SCALE, MIN_VALUE, MAX_VALUE));
}
// Add driver along eD axis in alphabetical order
coefficients.add(2 * nB + 1, new ParameterDriver(ECOM_COEFFICIENT + " D0", value, SCALE, MIN_VALUE, MAX_VALUE));
for (int i = 2 * nB + 2; i < 2 * nB + 2 + nD; i++) {
coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Dcos" + Integer.toString(i - (2 * nB + 2)), value, SCALE, MIN_VALUE, MAX_VALUE));
}
for (int i = 2 * nB + 2 + nD; i < 2 * (nB + nD) + 2; i++) {
coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Dsin" + Integer.toString(i - (2 * nB + nD + 2)), value, SCALE, MIN_VALUE, MAX_VALUE));
}
// Add Y0
coefficients.add(new ParameterDriver(ECOM_COEFFICIENT + " Y0", value, SCALE, MIN_VALUE, MAX_VALUE));
// For ECOM2 model, all parameters are estimated
coefficients.forEach(parameter -> parameter.setSelected(true));
this.sun = sun;
}
/** {@inheritDoc} */
@Override
public Vector3D acceleration(final SpacecraftState s, final double[] parameters) {
// Spacecraft and Sun position vectors (expressed in the spacecraft's frame)
final Vector3D satPos = s.getPosition();
final Vector3D sunPos = sun.getPosition(s.getDate(), s.getFrame());
// Build the coordinate system
final Vector3D Z = s.getPVCoordinates().getMomentum();
final Vector3D Y = Z.crossProduct(sunPos).normalize();
final Vector3D X = Y.crossProduct(Z).normalize();
// Build eD, eY, eB vectors
final Vector3D eD = sunPos.subtract(satPos).normalize();
final Vector3D eY = eD.crossProduct(satPos).normalize();
final Vector3D eB = eD.crossProduct(eY);
// Angular argument difference u_s - u
final double delta_u = FastMath.atan2(satPos.dotProduct(Y), satPos.dotProduct(X));
// Compute B(u)
double b_u = parameters[0];
for (int i = 1; i < nB + 1; i++) {
final SinCos sc = FastMath.sinCos((2 * i - 1) * delta_u);
b_u += parameters[i] * sc.cos() + parameters[i + nB] * sc.sin();
}
// Compute D(u)
double d_u = parameters[2 * nB + 1];
for (int i = 1; i < nD + 1; i++) {
final SinCos sc = FastMath.sinCos(2 * i * delta_u);
d_u += parameters[2 * nB + 1 + i] * sc.cos() + parameters[2 * nB + 1 + i + nD] * sc.sin();
}
// Return acceleration
return new Vector3D(d_u, eD, parameters[2 * (nD + nB) + 2], eY, b_u, eB);
}
/** {@inheritDoc} */
@Override
public <T extends CalculusFieldElement<T>> FieldVector3D<T> acceleration(final FieldSpacecraftState<T> s, final T[] parameters) {
// Spacecraft and Sun position vectors (expressed in the spacecraft's frame)
final FieldVector3D<T> satPos = s.getPosition();
final FieldVector3D<T> sunPos = sun.getPosition(s.getDate(), s.getFrame());
// Build the coordinate system
final FieldVector3D<T> Z = s.getPVCoordinates().getMomentum();
final FieldVector3D<T> Y = Z.crossProduct(sunPos).normalize();
final FieldVector3D<T> X = Y.crossProduct(Z).normalize();
// Build eD, eY, eB vectors
final FieldVector3D<T> eD = sunPos.subtract(satPos).normalize();
final FieldVector3D<T> eY = eD.crossProduct(satPos).normalize();
final FieldVector3D<T> eB = eD.crossProduct(eY);
// Angular argument difference u_s - u
final T delta_u = FastMath.atan2(satPos.dotProduct(Y), satPos.dotProduct(X));
// Compute B(u)
T b_u = parameters[0];
for (int i = 1; i < nB + 1; i++) {
final FieldSinCos<T> sc = FastMath.sinCos(delta_u.multiply(2 * i - 1));
b_u = b_u.add(sc.cos().multiply(parameters[i])).add(sc.sin().multiply(parameters[i + nB]));
}
// Compute D(u)
T d_u = parameters[2 * nB + 1];
for (int i = 1; i < nD + 1; i++) {
final FieldSinCos<T> sc = FastMath.sinCos(delta_u.multiply(2 * i));
d_u = d_u.add(sc.cos().multiply(parameters[2 * nB + 1 + i])).add(sc.sin().multiply(parameters[2 * nB + 1 + i + nD]));
}
// Return the acceleration
return new FieldVector3D<>(d_u, eD, parameters[2 * (nD + nB) + 2], eY, b_u, eB);
}
/** {@inheritDoc} */
@Override
public List<ParameterDriver> getParametersDrivers() {
return Collections.unmodifiableList(coefficients);
}
}