Examples with InertialProvider org.orekit.attitudes.InertialProvider used on opensource projects

Example 6 with InertialProvider

use of org.orekit.attitudes.InertialProvider in project Orekit by CS-SI.

the class ConstantThrustManeuverTest method testInertialManeuver.

@Test
public void testInertialManeuver() throws OrekitException {
    final double isp = 318;
    final double mass = 2500;
    final double a = 24396159;
    final double e = 0.72831215;
    final double i = FastMath.toRadians(7);
    final double omega = FastMath.toRadians(180);
    final double OMEGA = FastMath.toRadians(261);
    final double lv = 0;
    final AbsoluteDate initDate = new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC());
    final Orbit orbit = new KeplerianOrbit(a, e, i, omega, OMEGA, lv, PositionAngle.TRUE, FramesFactory.getEME2000(), initDate, mu);
    final double duration = 3653.99;
    final double f = 420;
    final double delta = FastMath.toRadians(-7.4978);
    final double alpha = FastMath.toRadians(351);
    final AttitudeProvider inertialLaw = new InertialProvider(new Rotation(new Vector3D(alpha, delta), Vector3D.PLUS_I));
    final AttitudeProvider lofLaw = new LofOffset(orbit.getFrame(), LOFType.VNC);
    final SpacecraftState initialState = new SpacecraftState(orbit, inertialLaw.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass);
    final AbsoluteDate fireDate = new AbsoluteDate(new DateComponents(2004, 01, 02), new TimeComponents(04, 15, 34.080), TimeScalesFactory.getUTC());
    final ConstantThrustManeuver maneuverWithoutOverride = new ConstantThrustManeuver(fireDate, duration, f, isp, Vector3D.PLUS_I);
    Assert.assertEquals(f, maneuverWithoutOverride.getThrust(), 1.0e-10);
    Assert.assertEquals(isp, maneuverWithoutOverride.getISP(), 1.0e-10);
    // reference propagation:
    // propagator already uses inertial law
    // maneuver does not need to override it to get an inertial maneuver
    double[][] tol = NumericalPropagator.tolerances(1.0, orbit, OrbitType.KEPLERIAN);
    AdaptiveStepsizeIntegrator integrator1 = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator1.setInitialStepSize(60);
    final NumericalPropagator propagator1 = new NumericalPropagator(integrator1);
    propagator1.setInitialState(initialState);
    propagator1.setAttitudeProvider(inertialLaw);
    propagator1.addForceModel(maneuverWithoutOverride);
    final SpacecraftState finalState1 = propagator1.propagate(fireDate.shiftedBy(3800));
    // test propagation:
    // propagator uses a LOF-aligned law
    // maneuver needs to override it to get an inertial maneuver
    final ConstantThrustManeuver maneuverWithOverride = new ConstantThrustManeuver(fireDate, duration, f, isp, inertialLaw, Vector3D.PLUS_I);
    Assert.assertEquals(f, maneuverWithoutOverride.getThrust(), 1.0e-10);
    Assert.assertEquals(isp, maneuverWithoutOverride.getISP(), 1.0e-10);
    AdaptiveStepsizeIntegrator integrator2 = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator2.setInitialStepSize(60);
    final NumericalPropagator propagator2 = new NumericalPropagator(integrator2);
    propagator2.setInitialState(initialState);
    propagator2.setAttitudeProvider(lofLaw);
    propagator2.addForceModel(maneuverWithOverride);
    final SpacecraftState finalState2 = propagator2.propagate(finalState1.getDate());
    Assert.assertThat(finalState2.getPVCoordinates(), OrekitMatchers.pvCloseTo(finalState1.getPVCoordinates(), 1.0e-10));
    // intentionally wrong propagation, that will produce a very different state
    // propagator uses LOF attitude,
    // maneuver forget to override it, so maneuver will be LOF-aligned in this case
    AdaptiveStepsizeIntegrator integrator3 = new DormandPrince853Integrator(0.001, 1000, tol[0], tol[1]);
    integrator3.setInitialStepSize(60);
    final NumericalPropagator propagator3 = new NumericalPropagator(integrator3);
    propagator3.setInitialState(initialState);
    propagator3.setAttitudeProvider(lofLaw);
    propagator3.addForceModel(maneuverWithoutOverride);
    final SpacecraftState finalState3 = propagator3.propagate(finalState1.getDate());
    Assert.assertEquals(345859.0, Vector3D.distance(finalState1.getPVCoordinates().getPosition(), finalState3.getPVCoordinates().getPosition()), 1.0);
}

Example 7 with InertialProvider

use of org.orekit.attitudes.InertialProvider in project Orekit by CS-SI.

the class ImpulseManeuverTest method testAdditionalStateNumerical.

@Test
public void testAdditionalStateNumerical() throws OrekitException {
    final double mu = CelestialBodyFactory.getEarth().getGM();
    final double initialX = 7100e3;
    final double initialY = 0.0;
    final double initialZ = 1300e3;
    final double initialVx = 0;
    final double initialVy = 8000;
    final double initialVz = 1000;
    final Vector3D position = new Vector3D(initialX, initialY, initialZ);
    final Vector3D velocity = new Vector3D(initialVx, initialVy, initialVz);
    final AbsoluteDate epoch = new AbsoluteDate(2010, 1, 1, 0, 0, 0, TimeScalesFactory.getUTC());
    final TimeStampedPVCoordinates pv = new TimeStampedPVCoordinates(epoch, position, velocity, Vector3D.ZERO);
    final Orbit initialOrbit = new CartesianOrbit(pv, FramesFactory.getEME2000(), mu);
    final double totalPropagationTime = 10.0;
    final double deltaX = 0.01;
    final double deltaY = 0.02;
    final double deltaZ = 0.03;
    final double isp = 300;
    final Vector3D deltaV = new Vector3D(deltaX, deltaY, deltaZ);
    final AttitudeProvider attitudeProvider = new LofOffset(initialOrbit.getFrame(), LOFType.VNC);
    final Attitude initialAttitude = attitudeProvider.getAttitude(initialOrbit, initialOrbit.getDate(), initialOrbit.getFrame());
    double[][] tolerances = NumericalPropagator.tolerances(10.0, initialOrbit, initialOrbit.getType());
    DormandPrince853Integrator integrator = new DormandPrince853Integrator(1.0e-3, 60, tolerances[0], tolerances[1]);
    NumericalPropagator propagator = new NumericalPropagator(integrator);
    propagator.setOrbitType(initialOrbit.getType());
    PartialDerivativesEquations pde = new PartialDerivativesEquations("derivatives", propagator);
    final SpacecraftState initialState = pde.setInitialJacobians(new SpacecraftState(initialOrbit, initialAttitude));
    propagator.resetInitialState(initialState);
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(0.5 * totalPropagationTime));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, RotationConvention.VECTOR_OPERATOR, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector, attitudeOverride, deltaV, isp).withThreshold(1.0e-3);
    propagator.addEventDetector(burnAtEpoch);
    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));
    Assert.assertEquals(1, finalState.getAdditionalStates().size());
    Assert.assertEquals(36, finalState.getAdditionalState("derivatives").length);
    double[][] stateTransitionMatrix = new double[6][6];
    pde.getMapper().getStateJacobian(finalState, stateTransitionMatrix);
    for (int i = 0; i < 6; ++i) {
        for (int j = 0; j < 6; ++j) {
            double sIJ = stateTransitionMatrix[i][j];
            if (j == i) {
                // dPi/dPj and dVi/dVj are roughly 1 for small propagation times
                Assert.assertEquals(1.0, sIJ, 2.0e-4);
            } else if (j == i + 3) {
                // dVi/dPi is roughly the propagation time for small propagation times
                Assert.assertEquals(totalPropagationTime, sIJ, 4.0e-5 * totalPropagationTime);
            } else {
                // other derivatives are almost zero for small propagation times
                Assert.assertEquals(0, sIJ, 1.0e-4);
            }
        }
    }
}

Example 8 with InertialProvider

use of org.orekit.attitudes.InertialProvider in project Orekit by CS-SI.

the class ImpulseManeuverTest method testInertialManeuver.

@Test
public void testInertialManeuver() throws OrekitException {
    final double mu = CelestialBodyFactory.getEarth().getGM();
    final double initialX = 7100e3;
    final double initialY = 0.0;
    final double initialZ = 1300e3;
    final double initialVx = 0;
    final double initialVy = 8000;
    final double initialVz = 1000;
    final Vector3D position = new Vector3D(initialX, initialY, initialZ);
    final Vector3D velocity = new Vector3D(initialVx, initialVy, initialVz);
    final AbsoluteDate epoch = new AbsoluteDate(2010, 1, 1, 0, 0, 0, TimeScalesFactory.getUTC());
    final TimeStampedPVCoordinates state = new TimeStampedPVCoordinates(epoch, position, velocity, Vector3D.ZERO);
    final Orbit initialOrbit = new CartesianOrbit(state, FramesFactory.getEME2000(), mu);
    final double totalPropagationTime = 0.00001;
    final double driftTimeInSec = totalPropagationTime / 2.0;
    final double deltaX = 0.01;
    final double deltaY = 0.02;
    final double deltaZ = 0.03;
    final double isp = 300;
    final Vector3D deltaV = new Vector3D(deltaX, deltaY, deltaZ);
    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit, new LofOffset(initialOrbit.getFrame(), LOFType.VNC));
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(driftTimeInSec));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, RotationConvention.VECTOR_OPERATOR, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector, attitudeOverride, deltaV, isp).withThreshold(driftTimeInSec / 4);
    propagator.addEventDetector(burnAtEpoch);
    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));
    final double finalVxExpected = initialVx + deltaX;
    final double finalVyExpected = initialVy + deltaY;
    final double finalVzExpected = initialVz + deltaZ;
    final double maneuverTolerance = 1e-4;
    final Vector3D finalVelocity = finalState.getPVCoordinates().getVelocity();
    Assert.assertEquals(finalVxExpected, finalVelocity.getX(), maneuverTolerance);
    Assert.assertEquals(finalVyExpected, finalVelocity.getY(), maneuverTolerance);
    Assert.assertEquals(finalVzExpected, finalVelocity.getZ(), maneuverTolerance);
}

Example 9 with InertialProvider

use of org.orekit.attitudes.InertialProvider in project Orekit by CS-SI.

the class ImpulseManeuverTest method testBackAndForth.

@Test
public void testBackAndForth() throws OrekitException {
    final AttitudeProvider lof = new LofOffset(FramesFactory.getEME2000(), LOFType.VNC);
    final double mu = Constants.EIGEN5C_EARTH_MU;
    final AbsoluteDate iniDate = new AbsoluteDate(2003, 5, 1, 17, 30, 0.0, TimeScalesFactory.getUTC());
    final Orbit pastOrbit = new KeplerianOrbit(7e6, 1.0e-4, FastMath.toRadians(98.5), FastMath.toRadians(87.0), FastMath.toRadians(216.1807), FastMath.toRadians(319.779), PositionAngle.MEAN, FramesFactory.getEME2000(), iniDate, mu);
    final double pastMass = 2500.0;
    DateDetector dateDetector = new DateDetector(iniDate.shiftedBy(600));
    Vector3D deltaV = new Vector3D(12.0, 1.0, -4.0);
    final double isp = 300;
    ImpulseManeuver<DateDetector> maneuver = new ImpulseManeuver<DateDetector>(dateDetector, new InertialProvider(Rotation.IDENTITY), deltaV, isp).withMaxCheck(3600.0).withThreshold(1.0e-6);
    double span = 900.0;
    KeplerianPropagator forwardPropagator = new KeplerianPropagator(pastOrbit, lof, mu, pastMass);
    forwardPropagator.addEventDetector(maneuver);
    SpacecraftState futureState = forwardPropagator.propagate(pastOrbit.getDate().shiftedBy(span));
    KeplerianPropagator backwardPropagator = new KeplerianPropagator(futureState.getOrbit(), lof, mu, futureState.getMass());
    backwardPropagator.addEventDetector(maneuver);
    SpacecraftState rebuiltPast = backwardPropagator.propagate(pastOrbit.getDate());
    Assert.assertEquals(0.0, Vector3D.distance(pastOrbit.getPVCoordinates().getPosition(), rebuiltPast.getPVCoordinates().getPosition()), 2.0e-8);
    Assert.assertEquals(0.0, Vector3D.distance(pastOrbit.getPVCoordinates().getVelocity(), rebuiltPast.getPVCoordinates().getVelocity()), 2.0e-11);
    Assert.assertEquals(pastMass, rebuiltPast.getMass(), 5.0e-13);
}

Example 10 with InertialProvider

use of org.orekit.attitudes.InertialProvider in project Orekit by CS-SI.

the class ImpulseManeuverTest method testAdditionalStateKeplerian.

@Test
public void testAdditionalStateKeplerian() throws OrekitException {
    final double mu = CelestialBodyFactory.getEarth().getGM();
    final double initialX = 7100e3;
    final double initialY = 0.0;
    final double initialZ = 1300e3;
    final double initialVx = 0;
    final double initialVy = 8000;
    final double initialVz = 1000;
    final Vector3D position = new Vector3D(initialX, initialY, initialZ);
    final Vector3D velocity = new Vector3D(initialVx, initialVy, initialVz);
    final AbsoluteDate epoch = new AbsoluteDate(2010, 1, 1, 0, 0, 0, TimeScalesFactory.getUTC());
    final TimeStampedPVCoordinates pv = new TimeStampedPVCoordinates(epoch, position, velocity, Vector3D.ZERO);
    final Orbit initialOrbit = new CartesianOrbit(pv, FramesFactory.getEME2000(), mu);
    final double totalPropagationTime = 10;
    final double deltaX = 0.01;
    final double deltaY = 0.02;
    final double deltaZ = 0.03;
    final double isp = 300;
    final Vector3D deltaV = new Vector3D(deltaX, deltaY, deltaZ);
    final AttitudeProvider attitudeProvider = new LofOffset(initialOrbit.getFrame(), LOFType.VNC);
    final Attitude initialAttitude = attitudeProvider.getAttitude(initialOrbit, initialOrbit.getDate(), initialOrbit.getFrame());
    final SpacecraftState initialState = new SpacecraftState(initialOrbit, initialAttitude);
    KeplerianPropagator propagator = new KeplerianPropagator(initialOrbit);
    propagator.resetInitialState(initialState.addAdditionalState("testOnly", -1.0));
    DateDetector dateDetector = new DateDetector(epoch.shiftedBy(0.5 * totalPropagationTime));
    InertialProvider attitudeOverride = new InertialProvider(new Rotation(RotationOrder.XYX, RotationConvention.VECTOR_OPERATOR, 0, 0, 0));
    ImpulseManeuver<DateDetector> burnAtEpoch = new ImpulseManeuver<DateDetector>(dateDetector, attitudeOverride, deltaV, isp).withThreshold(1.0e-3);
    propagator.addEventDetector(burnAtEpoch);
    SpacecraftState finalState = propagator.propagate(epoch.shiftedBy(totalPropagationTime));
    Assert.assertEquals(1, finalState.getAdditionalStates().size());
    Assert.assertEquals(-1.0, finalState.getAdditionalState("testOnly")[0], 1.0e-15);
}