ThirdBodyAttractionEpoch.java
- /* Copyright 2002-2022 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.forces.gravity;
- import org.hipparchus.analysis.differentiation.Gradient;
- import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
- import org.hipparchus.geometry.euclidean.threed.Vector3D;
- import org.orekit.bodies.CelestialBody;
- import org.orekit.propagation.SpacecraftState;
- /** Third body attraction force model.
- * This class is a copy of {@link ThirdBodyAttraction} class.
- * The computation of derivatives of the acceleration
- * w.r.t. the Epoch has been added.
- *
- * @author Fabien Maussion
- * @author Véronique Pommier-Maurussane
- * @since 10.2
- */
- public class ThirdBodyAttractionEpoch extends ThirdBodyAttraction {
- /** The body to consider. */
- private final CelestialBody body;
- /** Simple constructor.
- * @param body the third body to consider
- * (ex: {@link org.orekit.bodies.CelestialBodyFactory#getSun()} or
- * {@link org.orekit.bodies.CelestialBodyFactory#getMoon()})
- */
- public ThirdBodyAttractionEpoch(final CelestialBody body) {
- super(body);
- this.body = body;
- }
- /** Compute acceleration.
- * @param s current state information: date, kinematics, attitude
- * @param parameters values of the force model parameters
- * @return acceleration in same frame as state
- */
- private FieldVector3D<Gradient> accelerationToEpoch(final SpacecraftState s, final double[] parameters) {
- final double gm = parameters[0];
- // compute bodies separation vectors and squared norm
- final Vector3D centralToBody = body.getPVCoordinates(s.getDate(), s.getFrame()).getPosition();
- // Spacecraft Position
- final double rx = centralToBody.getX();
- final double ry = centralToBody.getY();
- final double rz = centralToBody.getZ();
- final int freeParameters = 3;
- final Gradient fpx = Gradient.variable(freeParameters, 0, rx);
- final Gradient fpy = Gradient.variable(freeParameters, 1, ry);
- final Gradient fpz = Gradient.variable(freeParameters, 2, rz);
- final FieldVector3D<Gradient> centralToBodyFV = new FieldVector3D<>(new Gradient[] {fpx, fpy, fpz});
- final Gradient r2Central = centralToBodyFV.getNormSq();
- final FieldVector3D<Gradient> satToBody = centralToBodyFV.subtract(s.getPVCoordinates().getPosition());
- final Gradient r2Sat = satToBody.getNormSq();
- return new FieldVector3D<>(gm, satToBody.scalarMultiply(r2Sat.multiply(r2Sat.sqrt()).reciprocal()),
- -gm, centralToBodyFV.scalarMultiply(r2Central.multiply(r2Central.sqrt()).reciprocal()));
- }
- /** Compute derivatives of the state w.r.t epoch.
- * @param s current state information: date, kinematics, attitude
- * @param parameters values of the force model parameters
- * @return derivatives
- */
- public double[] getDerivativesToEpoch(final SpacecraftState s, final double[] parameters) {
- final FieldVector3D<Gradient> acc = accelerationToEpoch(s, parameters);
- final Vector3D centralToBodyVelocity = body.getPVCoordinates(s.getDate(), s.getFrame()).getVelocity();
- final double[] dAccxdR1i = acc.getX().getGradient();
- final double[] dAccydR1i = acc.getY().getGradient();
- final double[] dAcczdR1i = acc.getZ().getGradient();
- final double[] v = centralToBodyVelocity.toArray();
- return new double[] {
- dAccxdR1i[0] * v[0] + dAccxdR1i[1] * v[1] + dAccxdR1i[2] * v[2],
- dAccydR1i[0] * v[0] + dAccydR1i[1] * v[1] + dAccydR1i[2] * v[2],
- dAcczdR1i[0] * v[0] + dAcczdR1i[1] * v[1] + dAcczdR1i[2] * v[2]
- };
- }
- }