List of usage examples for org.apache.commons.math3.util FastMath expm1
public static double expm1(double x)
From source file:de.tuberlin.uebb.jbop.example.DSCompilerOnlyCompose.java
@Override public void expm1(final double[] operand, final double[] result) { // create the function value and derivatives final double[] function = new double[1 + order]; function[0] = FastMath.expm1(operand[0]); Arrays.fill(function, 1, 1 + order, FastMath.exp(operand[0])); // apply function composition compose(operand, function, result);/*from ww w.j a va 2s . c o m*/ }
From source file:de.tuberlin.uebb.jbop.example.DSCompiler.java
@Override @Optimizable//w w w. j a va 2 s . c o m @StrictLoops public void expm1(final double[] operand, final double[] result) { // create the function value and derivatives final double[] function = new double[1 + order]; function[0] = FastMath.expm1(operand[0]); Arrays.fill(function, 1, 1 + order, FastMath.exp(operand[0])); // apply function composition compose(operand, function, result); }
From source file:org.esa.beam.util.math.FastMathPerformance.java
public void testExpm1() { System.gc();// w w w. ja va2 s .c om double x = 0; long time = System.nanoTime(); for (int i = 0; i < RUNS; i++) x += StrictMath.expm1(-i * F1); long strictTime = System.nanoTime() - time; System.gc(); double y = 0; time = System.nanoTime(); for (int i = 0; i < RUNS; i++) y += FastMath.expm1(-i * F1); long fastTime = System.nanoTime() - time; System.gc(); double z = 0; time = System.nanoTime(); for (int i = 0; i < RUNS; i++) z += Math.expm1(-i * F1); long mathTime = System.nanoTime() - time; report("expm1", x + y + z, strictTime, fastTime, mathTime); }
From source file:org.orekit.forces.maneuvers.SmallManeuverAnalyticalModelTest.java
private BoundedPropagator getEphemeris(final Orbit orbit, final double mass, final AbsoluteDate t0, final Vector3D dV, final double f, final double isp) throws OrekitException { AttitudeProvider law = new LofOffset(orbit.getFrame(), LOFType.LVLH); final SpacecraftState initialState = new SpacecraftState(orbit, law.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass); // set up numerical propagator final double dP = 1.0; double[][] tolerances = NumericalPropagator.tolerances(dP, orbit, orbit.getType()); AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 1000, tolerances[0], tolerances[1]);//from w w w . ja v a 2 s . co m integrator.setInitialStepSize(orbit.getKeplerianPeriod() / 100.0); final NumericalPropagator propagator = new NumericalPropagator(integrator); propagator.setOrbitType(orbit.getType()); propagator.setInitialState(initialState); propagator.setAttitudeProvider(law); if (dV.getNorm() > 1.0e-6) { // set up maneuver final double vExhaust = Constants.G0_STANDARD_GRAVITY * isp; final double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust); final ConstantThrustManeuver maneuver = new ConstantThrustManeuver(t0, dt, f, isp, dV.normalize()); propagator.addForceModel(maneuver); } propagator.setEphemerisMode(); propagator.propagate(t0.shiftedBy(5 * orbit.getKeplerianPeriod())); return propagator.getGeneratedEphemeris(); }
From source file:org.orekit.propagation.analytical.AdapterPropagatorTest.java
@Test public void testLowEarthOrbit() throws OrekitException, ParseException, IOException { Orbit leo = new CircularOrbit( 7200000.0, -1.0e-5, 2.0e-4, FastMath.toRadians(98.0), FastMath.toRadians(123.456), 0.0, PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU); double mass = 5600.0; AbsoluteDate t0 = leo.getDate().shiftedBy(1000.0); Vector3D dV = new Vector3D(-0.1, 0.2, 0.3); double f = 20.0; double isp = 315.0; double vExhaust = Constants.G0_STANDARD_GRAVITY * isp; double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust); BoundedPropagator withoutManeuver = getEphemeris(leo, mass, 5, new LofOffset(leo.getFrame(), LOFType.LVLH), t0, Vector3D.ZERO, f, isp, false, false, null); BoundedPropagator withManeuver = getEphemeris(leo, mass, 5, new LofOffset(leo.getFrame(), LOFType.LVLH), t0, dV, f, isp, false, false, null); // we set up a model that reverts the maneuvers AdapterPropagator adapterPropagator = new AdapterPropagator(withManeuver); AdapterPropagator.DifferentialEffect effect = new SmallManeuverAnalyticalModel( adapterPropagator.propagate(t0), dV.negate(), isp); adapterPropagator.addEffect(effect); adapterPropagator.addAdditionalStateProvider(new AdditionalStateProvider() { public String getName() { return "dummy 3"; }/*from w w w . j a va 2 s .c o m*/ public double[] getAdditionalState(SpacecraftState state) { return new double[3]; } }); // the adapted propagators do not manage the additional states from the reference, // they simply forward them Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 1")); Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 2")); Assert.assertTrue(adapterPropagator.isAdditionalStateManaged("dummy 3")); for (AbsoluteDate t = t0.shiftedBy(0.5 * dt); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t .shiftedBy(60.0)) { PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, leo.getFrame()); PVCoordinates pvReverted = adapterPropagator.getPVCoordinates(t, leo.getFrame()); double revertError = new PVCoordinates(pvWithout, pvReverted).getPosition().getNorm(); Assert.assertEquals(0, revertError, 0.45); Assert.assertEquals(2, adapterPropagator.propagate(t).getAdditionalState("dummy 1").length); Assert.assertEquals(1, adapterPropagator.propagate(t).getAdditionalState("dummy 2").length); Assert.assertEquals(3, adapterPropagator.propagate(t).getAdditionalState("dummy 3").length); } }
From source file:org.orekit.propagation.analytical.AdapterPropagatorTest.java
@Test public void testEccentricOrbit() throws OrekitException, ParseException, IOException { Orbit heo = new KeplerianOrbit(90000000.0, 0.92, FastMath.toRadians(98.0), FastMath.toRadians(12.3456), FastMath.toRadians(123.456), FastMath.toRadians(1.23456), PositionAngle.MEAN, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2004, 01, 01), new TimeComponents(23, 30, 00.000), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU); double mass = 5600.0; AbsoluteDate t0 = heo.getDate().shiftedBy(1000.0); Vector3D dV = new Vector3D(-0.01, 0.02, 0.03); double f = 20.0; double isp = 315.0; double vExhaust = Constants.G0_STANDARD_GRAVITY * isp; double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust); BoundedPropagator withoutManeuver = getEphemeris(heo, mass, 5, new LofOffset(heo.getFrame(), LOFType.LVLH), t0, Vector3D.ZERO, f, isp, false, false, null); BoundedPropagator withManeuver = getEphemeris(heo, mass, 5, new LofOffset(heo.getFrame(), LOFType.LVLH), t0, dV, f, isp, false, false, null); // we set up a model that reverts the maneuvers AdapterPropagator adapterPropagator = new AdapterPropagator(withManeuver); AdapterPropagator.DifferentialEffect effect = new SmallManeuverAnalyticalModel( adapterPropagator.propagate(t0), dV.negate(), isp); adapterPropagator.addEffect(effect); adapterPropagator.addAdditionalStateProvider(new AdditionalStateProvider() { public String getName() { return "dummy 3"; }/*w ww .j av a2 s.c om*/ public double[] getAdditionalState(SpacecraftState state) { return new double[3]; } }); // the adapted propagators do not manage the additional states from the reference, // they simply forward them Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 1")); Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 2")); Assert.assertTrue(adapterPropagator.isAdditionalStateManaged("dummy 3")); for (AbsoluteDate t = t0.shiftedBy(0.5 * dt); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t .shiftedBy(300.0)) { PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, heo.getFrame()); PVCoordinates pvReverted = adapterPropagator.getPVCoordinates(t, heo.getFrame()); double revertError = new PVCoordinates(pvWithout, pvReverted).getPosition().getNorm(); Assert.assertEquals(0, revertError, 180.0); Assert.assertEquals(2, adapterPropagator.propagate(t).getAdditionalState("dummy 1").length); Assert.assertEquals(1, adapterPropagator.propagate(t).getAdditionalState("dummy 2").length); Assert.assertEquals(3, adapterPropagator.propagate(t).getAdditionalState("dummy 3").length); } }
From source file:org.orekit.propagation.analytical.AdapterPropagatorTest.java
@Test public void testNonKeplerian() throws OrekitException, ParseException, IOException { Orbit leo = new CircularOrbit(7204319.233600575, 4.434564637450575E-4, 0.0011736728299091088, 1.7211611441767323, 5.5552084166959474, 24950.321259193086, PositionAngle.TRUE, FramesFactory.getEME2000(), new AbsoluteDate(new DateComponents(2003, 9, 16), new TimeComponents(23, 11, 20.264), TimeScalesFactory.getUTC()), Constants.EIGEN5C_EARTH_MU); double mass = 4093.0; AbsoluteDate t0 = new AbsoluteDate(new DateComponents(2003, 9, 16), new TimeComponents(23, 14, 40.264), TimeScalesFactory.getUTC()); Vector3D dV = new Vector3D(0.0, 3.0, 0.0); double f = 40.0; double isp = 300.0; double vExhaust = Constants.G0_STANDARD_GRAVITY * isp; double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust); // setup a specific coefficient file for gravity potential as it will also // try to read a corrupted one otherwise GravityFieldFactory.addPotentialCoefficientsReader(new ICGEMFormatReader("g007_eigen_05c_coef", false)); NormalizedSphericalHarmonicsProvider gravityField = GravityFieldFactory.getNormalizedProvider(8, 8); BoundedPropagator withoutManeuver = getEphemeris(leo, mass, 10, new LofOffset(leo.getFrame(), LOFType.VNC), t0, Vector3D.ZERO, f, isp, true, true, gravityField); BoundedPropagator withManeuver = getEphemeris(leo, mass, 10, new LofOffset(leo.getFrame(), LOFType.VNC), t0, dV, f, isp, true, true, gravityField); // we set up a model that reverts the maneuvers AdapterPropagator adapterPropagator = new AdapterPropagator(withManeuver); SpacecraftState state0 = adapterPropagator.propagate(t0); AdapterPropagator.DifferentialEffect directEffect = new SmallManeuverAnalyticalModel(state0, dV.negate(), isp);//from w ww. j ava 2 s.c o m AdapterPropagator.DifferentialEffect derivedEffect = new J2DifferentialEffect(state0, directEffect, false, GravityFieldFactory.getUnnormalizedProvider(gravityField)); adapterPropagator.addEffect(directEffect); adapterPropagator.addEffect(derivedEffect); adapterPropagator.addAdditionalStateProvider(new AdditionalStateProvider() { public String getName() { return "dummy 3"; } public double[] getAdditionalState(SpacecraftState state) { return new double[3]; } }); // the adapted propagators do not manage the additional states from the reference, // they simply forward them Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 1")); Assert.assertFalse(adapterPropagator.isAdditionalStateManaged("dummy 2")); Assert.assertTrue(adapterPropagator.isAdditionalStateManaged("dummy 3")); double maxDelta = 0; double maxNominal = 0; for (AbsoluteDate t = t0.shiftedBy(0.5 * dt); t.compareTo(withoutManeuver.getMaxDate()) < 0; t = t .shiftedBy(600.0)) { PVCoordinates pvWithout = withoutManeuver.getPVCoordinates(t, leo.getFrame()); PVCoordinates pvWith = withManeuver.getPVCoordinates(t, leo.getFrame()); PVCoordinates pvReverted = adapterPropagator.getPVCoordinates(t, leo.getFrame()); double nominal = new PVCoordinates(pvWithout, pvWith).getPosition().getNorm(); double revertError = new PVCoordinates(pvWithout, pvReverted).getPosition().getNorm(); maxDelta = FastMath.max(maxDelta, revertError); maxNominal = FastMath.max(maxNominal, nominal); Assert.assertEquals(2, adapterPropagator.propagate(t).getAdditionalState("dummy 1").length); Assert.assertEquals(1, adapterPropagator.propagate(t).getAdditionalState("dummy 2").length); Assert.assertEquals(3, adapterPropagator.propagate(t).getAdditionalState("dummy 3").length); } Assert.assertTrue(maxDelta < 120); Assert.assertTrue(maxNominal > 2800); }
From source file:org.orekit.propagation.analytical.AdapterPropagatorTest.java
private BoundedPropagator getEphemeris(final Orbit orbit, final double mass, final int nbOrbits, final AttitudeProvider law, final AbsoluteDate t0, final Vector3D dV, final double f, final double isp, final boolean sunAttraction, final boolean moonAttraction, final NormalizedSphericalHarmonicsProvider gravityField) throws OrekitException, ParseException, IOException { SpacecraftState initialState = new SpacecraftState(orbit, law.getAttitude(orbit, orbit.getDate(), orbit.getFrame()), mass); // add some dummy additional states initialState = initialState.addAdditionalState("dummy 1", 1.25, 2.5); initialState = initialState.addAdditionalState("dummy 2", 5.0); // set up numerical propagator final double dP = 1.0; double[][] tolerances = NumericalPropagator.tolerances(dP, orbit, orbit.getType()); AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(0.001, 1000, tolerances[0], tolerances[1]);// w w w . ja va 2 s. c om integrator.setInitialStepSize(orbit.getKeplerianPeriod() / 100.0); final NumericalPropagator propagator = new NumericalPropagator(integrator); propagator.addAdditionalStateProvider(new AdditionalStateProvider() { public String getName() { return "dummy 2"; } public double[] getAdditionalState(SpacecraftState state) { return new double[] { 5.0 }; } }); propagator.setInitialState(initialState); propagator.setAttitudeProvider(law); if (dV.getNorm() > 1.0e-6) { // set up maneuver final double vExhaust = Constants.G0_STANDARD_GRAVITY * isp; final double dt = -(mass * vExhaust / f) * FastMath.expm1(-dV.getNorm() / vExhaust); final ConstantThrustManeuver maneuver = new ConstantThrustManeuver(t0.shiftedBy(-0.5 * dt), dt, f, isp, dV.normalize()); propagator.addForceModel(maneuver); } if (sunAttraction) { propagator.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getSun())); } if (moonAttraction) { propagator.addForceModel(new ThirdBodyAttraction(CelestialBodyFactory.getMoon())); } if (gravityField != null) { propagator.addForceModel( new HolmesFeatherstoneAttractionModel(FramesFactory.getGTOD(false), gravityField)); } propagator.setEphemerisMode(); propagator.propagate(t0.shiftedBy(nbOrbits * orbit.getKeplerianPeriod())); final BoundedPropagator ephemeris = propagator.getGeneratedEphemeris(); // both the initial propagator and generated ephemeris manage one of the two // additional states, but they also contain unmanaged copies of the other one Assert.assertFalse(propagator.isAdditionalStateManaged("dummy 1")); Assert.assertTrue(propagator.isAdditionalStateManaged("dummy 2")); Assert.assertFalse(ephemeris.isAdditionalStateManaged("dummy 1")); Assert.assertTrue(ephemeris.isAdditionalStateManaged("dummy 2")); Assert.assertEquals(2, ephemeris.getInitialState().getAdditionalState("dummy 1").length); Assert.assertEquals(1, ephemeris.getInitialState().getAdditionalState("dummy 2").length); return ephemeris; }