/* Copyright 2002-2017 CS Systèmes d'Information
* Licensed to CS Systèmes d'Information (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.Field;
import org.hipparchus.analysis.differentiation.DSFactory;
import org.hipparchus.analysis.differentiation.DerivativeStructure;
import org.hipparchus.geometry.euclidean.threed.FieldVector3D;
import org.hipparchus.geometry.euclidean.threed.Vector3D;
import org.hipparchus.ode.nonstiff.AdaptiveStepsizeFieldIntegrator;
import org.hipparchus.ode.nonstiff.AdaptiveStepsizeIntegrator;
import org.hipparchus.ode.nonstiff.DormandPrince853FieldIntegrator;
import org.hipparchus.ode.nonstiff.DormandPrince853Integrator;
import org.hipparchus.ode.nonstiff.GraggBulirschStoerIntegrator;
import org.hipparchus.random.GaussianRandomGenerator;
import org.hipparchus.random.RandomGenerator;
import org.hipparchus.random.UncorrelatedRandomVectorGenerator;
import org.hipparchus.random.Well19937a;
import org.hipparchus.util.FastMath;
import org.junit.Assert;
import org.junit.Before;
import org.junit.Test;
import org.orekit.Utils;
import org.orekit.bodies.CelestialBody;
import org.orekit.bodies.CelestialBodyFactory;
import org.orekit.errors.OrekitException;
import org.orekit.forces.AbstractForceModelTest;
import org.orekit.frames.Frame;
import org.orekit.frames.FramesFactory;
import org.orekit.orbits.CartesianOrbit;
import org.orekit.orbits.EquinoctialOrbit;
import org.orekit.orbits.FieldKeplerianOrbit;
import org.orekit.orbits.KeplerianOrbit;
import org.orekit.orbits.Orbit;
import org.orekit.orbits.OrbitType;
import org.orekit.orbits.PositionAngle;
import org.orekit.propagation.FieldSpacecraftState;
import org.orekit.propagation.SpacecraftState;
import org.orekit.propagation.numerical.FieldNumericalPropagator;
import org.orekit.propagation.numerical.NumericalPropagator;
import org.orekit.propagation.sampling.OrekitFixedStepHandler;
import org.orekit.time.AbsoluteDate;
import org.orekit.time.DateComponents;
import org.orekit.time.FieldAbsoluteDate;
import org.orekit.time.TimeComponents;
import org.orekit.time.TimeScalesFactory;
import org.orekit.utils.Constants;
import org.orekit.utils.FieldPVCoordinates;
import org.orekit.utils.PVCoordinates;
public class ThirdBodyAttractionTest extends AbstractForceModelTest {
private double mu;
@Test(expected= OrekitException.class)
public void testSunContrib() throws OrekitException {
// initialization
AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 07, 01),
new TimeComponents(13, 59, 27.816),
TimeScalesFactory.getUTC());
Orbit orbit = new EquinoctialOrbit(42164000, 10e-3, 10e-3,
FastMath.tan(0.001745329) * FastMath.cos(2 * FastMath.PI / 3),
FastMath.tan(0.001745329) * FastMath.sin(2 * FastMath.PI / 3),
0.1, PositionAngle.TRUE, FramesFactory.getEME2000(), date, mu);
double period = 2 * FastMath.PI * orbit.getA() * FastMath.sqrt(orbit.getA() / orbit.getMu());
// set up propagator
NumericalPropagator calc =
new NumericalPropagator(new GraggBulirschStoerIntegrator(10.0, period, 0, 1.0e-5));
calc.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getSun()));
// set up step handler to perform checks
calc.setMasterMode(FastMath.floor(period), new ReferenceChecker(date) {
protected double hXRef(double t) {
return -1.06757e-3 + 0.221415e-11 * t + 18.9421e-5 *
FastMath.cos(3.9820426e-7*t) - 7.59983e-5 * FastMath.sin(3.9820426e-7*t);
}
protected double hYRef(double t) {
return 1.43526e-3 + 7.49765e-11 * t + 6.9448e-5 *
FastMath.cos(3.9820426e-7*t) + 17.6083e-5 * FastMath.sin(3.9820426e-7*t);
}
});
AbsoluteDate finalDate = date.shiftedBy(365 * period);
calc.setInitialState(new SpacecraftState(orbit));
calc.propagate(finalDate);
}
/**Testing if the propagation between the FieldPropagation and the propagation
* is equivalent.
* Also testing if propagating X+dX with the propagation is equivalent to
* propagation X with the FieldPropagation and then applying the taylor
* expansion of dX to the result.*/
@Test
public void RealFieldTest() throws OrekitException {
DSFactory factory = new DSFactory(6, 5);
DerivativeStructure a_0 = factory.variable(0, 7e7);
DerivativeStructure e_0 = factory.variable(1, 0.4);
DerivativeStructure i_0 = factory.variable(2, 85 * FastMath.PI / 180);
DerivativeStructure R_0 = factory.variable(3, 0.7);
DerivativeStructure O_0 = factory.variable(4, 0.5);
DerivativeStructure n_0 = factory.variable(5, 0.1);
Field<DerivativeStructure> field = a_0.getField();
DerivativeStructure zero = field.getZero();
FieldAbsoluteDate<DerivativeStructure> J2000 = new FieldAbsoluteDate<DerivativeStructure>(field);
Frame EME = FramesFactory.getEME2000();
FieldKeplerianOrbit<DerivativeStructure> FKO = new FieldKeplerianOrbit<DerivativeStructure>(a_0, e_0, i_0, R_0, O_0, n_0,
PositionAngle.MEAN,
EME,
J2000,
Constants.EIGEN5C_EARTH_MU);
FieldSpacecraftState<DerivativeStructure> initialState = new FieldSpacecraftState<DerivativeStructure>(FKO);
SpacecraftState iSR = initialState.toSpacecraftState();
OrbitType type = OrbitType.KEPLERIAN;
double[][] tolerance = NumericalPropagator.tolerances(10.0, FKO.toOrbit(), type);
AdaptiveStepsizeFieldIntegrator<DerivativeStructure> integrator =
new DormandPrince853FieldIntegrator<DerivativeStructure>(field, 0.001, 200, tolerance[0], tolerance[1]);
integrator.setInitialStepSize(zero.add(60));
AdaptiveStepsizeIntegrator RIntegrator =
new DormandPrince853Integrator(0.001, 200, tolerance[0], tolerance[1]);
RIntegrator.setInitialStepSize(60);
FieldNumericalPropagator<DerivativeStructure> FNP = new FieldNumericalPropagator<>(field, integrator);
FNP.setOrbitType(type);
FNP.setInitialState(initialState);
NumericalPropagator NP = new NumericalPropagator(RIntegrator);
NP.setOrbitType(type);
NP.setInitialState(iSR);
final ThirdBodyAttraction forceModel = new ThirdBodyAttraction(CelestialBodyFactory.getSun());
FNP.addForceModel(forceModel);
NP.addForceModel(forceModel);
FieldAbsoluteDate<DerivativeStructure> target = J2000.shiftedBy(10000.);
FieldSpacecraftState<DerivativeStructure> finalState_DS = FNP.propagate(target);
SpacecraftState finalState_R = NP.propagate(target.toAbsoluteDate());
FieldPVCoordinates<DerivativeStructure> finPVC_DS = finalState_DS.getPVCoordinates();
PVCoordinates finPVC_R = finalState_R.getPVCoordinates();
Assert.assertEquals(finPVC_DS.toPVCoordinates().getPosition().getX(), finPVC_R.getPosition().getX(), FastMath.abs(finPVC_R.getPosition().getX()) * 1e-11);
Assert.assertEquals(finPVC_DS.toPVCoordinates().getPosition().getY(), finPVC_R.getPosition().getY(), FastMath.abs(finPVC_R.getPosition().getY()) * 1e-11);
Assert.assertEquals(finPVC_DS.toPVCoordinates().getPosition().getZ(), finPVC_R.getPosition().getZ(), FastMath.abs(finPVC_R.getPosition().getZ()) * 1e-11);
long number = 23091991;
RandomGenerator RG = new Well19937a(number);
GaussianRandomGenerator NGG = new GaussianRandomGenerator(RG);
UncorrelatedRandomVectorGenerator URVG = new UncorrelatedRandomVectorGenerator(new double[] {0.0 , 0.0 , 0.0 , 0.0 , 0.0 , 0.0 },
new double[] {1e3, 0.01, 0.01, 0.01, 0.01, 0.01},
NGG);
double a_R = a_0.getReal();
double e_R = e_0.getReal();
double i_R = i_0.getReal();
double R_R = R_0.getReal();
double O_R = O_0.getReal();
double n_R = n_0.getReal();
for (int ii = 0; ii < 1; ii++){
double[] rand_next = URVG.nextVector();
double a_shift = a_R + rand_next[0];
double e_shift = e_R + rand_next[1];
double i_shift = i_R + rand_next[2];
double R_shift = R_R + rand_next[3];
double O_shift = O_R + rand_next[4];
double n_shift = n_R + rand_next[5];
KeplerianOrbit shiftedOrb = new KeplerianOrbit(a_shift, e_shift, i_shift, R_shift, O_shift, n_shift,
PositionAngle.MEAN,
EME,
J2000.toAbsoluteDate(),
Constants.EIGEN5C_EARTH_MU
);
SpacecraftState shift_iSR = new SpacecraftState(shiftedOrb);
NumericalPropagator shift_NP = new NumericalPropagator(RIntegrator);
shift_NP.setInitialState(shift_iSR);
shift_NP.addForceModel(forceModel);
SpacecraftState finalState_shift = shift_NP.propagate(target.toAbsoluteDate());
PVCoordinates finPVC_shift = finalState_shift.getPVCoordinates();
//position check
FieldVector3D<DerivativeStructure> pos_DS = finPVC_DS.getPosition();
double x_DS = pos_DS.getX().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double y_DS = pos_DS.getY().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double z_DS = pos_DS.getZ().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
//System.out.println(pos_DS.getX().getPartialDerivative(1));
double x = finPVC_shift.getPosition().getX();
double y = finPVC_shift.getPosition().getY();
double z = finPVC_shift.getPosition().getZ();
Assert.assertEquals(x_DS, x, FastMath.abs(x - pos_DS.getX().getReal()) * 1e-8);
Assert.assertEquals(y_DS, y, FastMath.abs(y - pos_DS.getY().getReal()) * 1e-8);
Assert.assertEquals(z_DS, z, FastMath.abs(z - pos_DS.getZ().getReal()) * 1e-8);
//velocity check
FieldVector3D<DerivativeStructure> vel_DS = finPVC_DS.getVelocity();
double vx_DS = vel_DS.getX().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double vy_DS = vel_DS.getY().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double vz_DS = vel_DS.getZ().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double vx = finPVC_shift.getVelocity().getX();
double vy = finPVC_shift.getVelocity().getY();
double vz = finPVC_shift.getVelocity().getZ();
Assert.assertEquals(vx_DS, vx, FastMath.abs(vx) * 1e-9);
Assert.assertEquals(vy_DS, vy, FastMath.abs(vy) * 1e-9);
Assert.assertEquals(vz_DS, vz, FastMath.abs(vz) * 1e-9);
//acceleration check
FieldVector3D<DerivativeStructure> acc_DS = finPVC_DS.getAcceleration();
double ax_DS = acc_DS.getX().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double ay_DS = acc_DS.getY().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double az_DS = acc_DS.getZ().taylor(rand_next[0],rand_next[1],rand_next[2],rand_next[3],rand_next[4],rand_next[5]);
double ax = finPVC_shift.getAcceleration().getX();
double ay = finPVC_shift.getAcceleration().getY();
double az = finPVC_shift.getAcceleration().getZ();
Assert.assertEquals(ax_DS, ax, FastMath.abs(ax) * 1e-8);
Assert.assertEquals(ay_DS, ay, FastMath.abs(ay) * 1e-8);
Assert.assertEquals(az_DS, az, FastMath.abs(az) * 1e-8);
}
}
/**Same test as the previous one but not adding the ForceModel to the NumericalPropagator
it is a test to validate the previous test.
(to test if the ForceModel it's actually
doing something in the Propagator and the FieldPropagator)*/
@Test
public void RealFieldExpectErrorTest() throws OrekitException {
DSFactory factory = new DSFactory(6, 5);
DerivativeStructure a_0 = factory.variable(0, 7e7);
DerivativeStructure e_0 = factory.variable(1, 0.4);
DerivativeStructure i_0 = factory.variable(2, 85 * FastMath.PI / 180);
DerivativeStructure R_0 = factory.variable(3, 0.7);
DerivativeStructure O_0 = factory.variable(4, 0.5);
DerivativeStructure n_0 = factory.variable(5, 0.1);
Field<DerivativeStructure> field = a_0.getField();
DerivativeStructure zero = field.getZero();
FieldAbsoluteDate<DerivativeStructure> J2000 = new FieldAbsoluteDate<DerivativeStructure>(field);
Frame EME = FramesFactory.getEME2000();
FieldKeplerianOrbit<DerivativeStructure> FKO = new FieldKeplerianOrbit<DerivativeStructure>(a_0, e_0, i_0, R_0, O_0, n_0,
PositionAngle.MEAN,
EME,
J2000,
Constants.EIGEN5C_EARTH_MU);
FieldSpacecraftState<DerivativeStructure> initialState = new FieldSpacecraftState<DerivativeStructure>(FKO);
SpacecraftState iSR = initialState.toSpacecraftState();
OrbitType type = OrbitType.KEPLERIAN;
double[][] tolerance = NumericalPropagator.tolerances(0.001, FKO.toOrbit(), type);
AdaptiveStepsizeFieldIntegrator<DerivativeStructure> integrator =
new DormandPrince853FieldIntegrator<DerivativeStructure>(field, 0.001, 200, tolerance[0], tolerance[1]);
integrator.setInitialStepSize(zero.add(60));
AdaptiveStepsizeIntegrator RIntegrator =
new DormandPrince853Integrator(0.001, 200, tolerance[0], tolerance[1]);
RIntegrator.setInitialStepSize(60);
FieldNumericalPropagator<DerivativeStructure> FNP = new FieldNumericalPropagator<>(field, integrator);
FNP.setOrbitType(type);
FNP.setInitialState(initialState);
NumericalPropagator NP = new NumericalPropagator(RIntegrator);
NP.setOrbitType(type);
NP.setInitialState(iSR);
final ThirdBodyAttraction forceModel = new ThirdBodyAttraction(CelestialBodyFactory.getSun());
FNP.addForceModel(forceModel);
//NOT ADDING THE FORCE MODEL TO THE NUMERICAL PROPAGATOR NP.addForceModel(forceModel);
FieldAbsoluteDate<DerivativeStructure> target = J2000.shiftedBy(10000.);
FieldSpacecraftState<DerivativeStructure> finalState_DS = FNP.propagate(target);
SpacecraftState finalState_R = NP.propagate(target.toAbsoluteDate());
FieldPVCoordinates<DerivativeStructure> finPVC_DS = finalState_DS.getPVCoordinates();
PVCoordinates finPVC_R = finalState_R.getPVCoordinates();
Assert.assertFalse(FastMath.abs(finPVC_DS.toPVCoordinates().getPosition().getX() - finPVC_R.getPosition().getX()) < FastMath.abs(finPVC_R.getPosition().getX()) * 1e-11);
Assert.assertFalse(FastMath.abs(finPVC_DS.toPVCoordinates().getPosition().getY() - finPVC_R.getPosition().getY()) < FastMath.abs(finPVC_R.getPosition().getY()) * 1e-11);
Assert.assertFalse(FastMath.abs(finPVC_DS.toPVCoordinates().getPosition().getZ() - finPVC_R.getPosition().getZ()) < FastMath.abs(finPVC_R.getPosition().getZ()) * 1e-11);
}
@Test
public void testMoonContrib() throws OrekitException {
// initialization
AbsoluteDate date = new AbsoluteDate(new DateComponents(1970, 07, 01),
new TimeComponents(13, 59, 27.816),
TimeScalesFactory.getUTC());
Orbit orbit =
new EquinoctialOrbit(42164000,10e-3,10e-3,
FastMath.tan(0.001745329) * FastMath.cos(2 * FastMath.PI / 3),
FastMath.tan(0.001745329) * FastMath.sin(2 * FastMath.PI / 3),
0.1, PositionAngle.TRUE, FramesFactory.getEME2000(), date, mu);
double period = 2 * FastMath.PI * orbit.getA() * FastMath.sqrt(orbit.getA() / orbit.getMu());
// set up propagator
NumericalPropagator calc =
new NumericalPropagator(new GraggBulirschStoerIntegrator(10.0, period, 0, 1.0e-5));
calc.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getMoon()));
// set up step handler to perform checks
calc.setMasterMode(FastMath.floor(period), new ReferenceChecker(date) {
protected double hXRef(double t) {
return -0.000906173 + 1.93933e-11 * t +
1.0856e-06 * FastMath.cos(5.30637e-05 * t) -
1.22574e-06 * FastMath.sin(5.30637e-05 * t);
}
protected double hYRef(double t) {
return 0.00151973 + 1.88991e-10 * t -
1.25972e-06 * FastMath.cos(5.30637e-05 * t) -
1.00581e-06 * FastMath.sin(5.30637e-05 * t);
}
});
AbsoluteDate finalDate = date.shiftedBy(31 * period);
calc.setInitialState(new SpacecraftState(orbit));
calc.propagate(finalDate);
}
private static abstract class ReferenceChecker implements OrekitFixedStepHandler {
private final AbsoluteDate reference;
protected ReferenceChecker(AbsoluteDate reference) {
this.reference = reference;
}
public void handleStep(SpacecraftState currentState, boolean isLast) {
double t = currentState.getDate().durationFrom(reference);
Assert.assertEquals(hXRef(t), currentState.getHx(), 1e-4);
Assert.assertEquals(hYRef(t), currentState.getHy(), 1e-4);
}
protected abstract double hXRef(double t);
protected abstract double hYRef(double t);
}
@Test
public void testParameterDerivative() throws OrekitException {
final Vector3D pos = new Vector3D(6.46885878304673824e+06, -1.88050918456274318e+06, -1.32931592294715829e+04);
final Vector3D vel = new Vector3D(2.14718074509906819e+03, 7.38239351251748485e+03, -1.14097953925384523e+01);
final SpacecraftState state =
new SpacecraftState(new CartesianOrbit(new PVCoordinates(pos, vel),
FramesFactory.getGCRF(),
new AbsoluteDate(2003, 3, 5, 0, 24, 0.0, TimeScalesFactory.getTAI()),
Constants.EIGEN5C_EARTH_MU));
final CelestialBody moon = CelestialBodyFactory.getMoon();
final ThirdBodyAttraction forceModel = new ThirdBodyAttraction(moon);
final String name = moon.getName() + ThirdBodyAttraction.ATTRACTION_COEFFICIENT_SUFFIX;
checkParameterDerivative(state, forceModel, name, 1.0, 7.0e-15);
}
@Test
public void testStateJacobian()
throws OrekitException {
// initialization
AbsoluteDate date = new AbsoluteDate(new DateComponents(2003, 03, 01),
new TimeComponents(13, 59, 27.816),
TimeScalesFactory.getUTC());
double i = FastMath.toRadians(98.7);
double omega = FastMath.toRadians(93.0);
double OMEGA = FastMath.toRadians(15.0 * 22.5);
Orbit orbit = new KeplerianOrbit(7201009.7124401, 1e-3, i , omega, OMEGA,
0, PositionAngle.MEAN, FramesFactory.getEME2000(), date,
Constants.EIGEN5C_EARTH_MU);
OrbitType integrationType = OrbitType.CARTESIAN;
double[][] tolerances = NumericalPropagator.tolerances(0.01, orbit, integrationType);
NumericalPropagator propagator =
new NumericalPropagator(new DormandPrince853Integrator(1.0e-3, 120,
tolerances[0], tolerances[1]));
propagator.setOrbitType(integrationType);
final CelestialBody moon = CelestialBodyFactory.getMoon();
final ThirdBodyAttraction forceModel = new ThirdBodyAttraction(moon);
propagator.addForceModel(forceModel);
SpacecraftState state0 = new SpacecraftState(orbit);
checkStateJacobian(propagator, state0, date.shiftedBy(3.5 * 3600.0),
1e4, tolerances[0], 2.0e-9);
}
@Before
public void setUp() {
mu = 3.986e14;
Utils.setDataRoot("regular-data");
}
}