| Package | Description |
|---|---|
| org.orekit.attitudes |
This package provides classes to represent simple attitudes.
|
| org.orekit.bodies |
This package provides interface to represent the position and geometry of
space objects such as stars, planets or asteroids.
|
| org.orekit.data |
This package provide base classes for exploring the configured data
directory tree and read external data that can be used by the library.
|
| org.orekit.errors |
This package provides classes to generate and handle exceptions.
|
| org.orekit.estimation.leastsquares |
The leastsquares package provides an implementation of a batch least
squares estimator engine to perform an orbit determination.
|
| org.orekit.estimation.measurements |
The measurements package defines everything that is related to orbit
determination measurements.
|
| org.orekit.estimation.measurements.modifiers | |
| org.orekit.files.ccsds |
This package provides a parser for orbit data stored in CCSDS Orbit Data Message format.
|
| org.orekit.files.general |
This package provides interfaces for orbit file representations and corresponding
parsers.
|
| org.orekit.files.sp3 |
This package provides a parser for orbit data stored in SP3 format.
|
| org.orekit.forces |
This package provides the interface for force models that will be used by the
NumericalPropagator, as well as
some classical spacecraft models for surface forces (spherical, box and solar array ...). |
| org.orekit.forces.drag |
This package provides all drag-related forces.
|
| org.orekit.forces.drag.atmosphere |
This package provides the atmosphere model interface and several implementations.
|
| org.orekit.forces.drag.atmosphere.data |
This package provides classes to get atmospheric data,
including solar flux and planetary geomagnetic indices.
|
| org.orekit.forces.gravity |
This package provides all gravity-related forces.
|
| org.orekit.forces.gravity.potential |
This package provides classes to read gravity field files and supports several
different formats.
|
| org.orekit.forces.maneuvers |
This package provides models of simple maneuvers.
|
| org.orekit.forces.radiation |
This package provides all radiation pressure related forces.
|
| org.orekit.frames |
This package provides classes to handle frames and transforms between them.
|
| org.orekit.gnss |
This package provides classes related to GNSS applications.
|
| org.orekit.models.earth |
This package provides models that simulate certain physical phenomena
of Earth and the near-Earth environment.
|
| org.orekit.models.earth.tessellation |
This package provides ways to do tessellation and sampling of zones of
interest over an ellipsoid surface.
|
| org.orekit.orbits |
This package provides classes to represent orbits.
|
| org.orekit.propagation |
Propagation
|
| org.orekit.propagation.analytical |
Top level package for analytical propagators.
|
| org.orekit.propagation.analytical.gnss |
This package provides classes to propagate GNSS orbits.
|
| org.orekit.propagation.analytical.tle |
This package provides classes to read and extrapolate tle's.
|
| org.orekit.propagation.conversion |
This package provides tools to convert a given propagator or a set of
SpacecraftState into another propagator. |
| org.orekit.propagation.events |
This package provides interfaces and classes dealing with events occurring during propagation.
|
| org.orekit.propagation.events.handlers |
This package provides an interface and classes dealing with events occurrence only.
|
| org.orekit.propagation.integration |
Utilities for integration-based propagators (both numerical and semi-analytical).
|
| org.orekit.propagation.numerical |
Top level package for numerical propagators.
|
| org.orekit.propagation.sampling |
This package provides interfaces and classes dealing with step handling during propagation.
|
| org.orekit.propagation.semianalytical.dsst |
This package provides an implementation of the Draper Semi-analytical
Satellite Theory (DSST).
|
| org.orekit.propagation.semianalytical.dsst.forces |
This package provides force models for Draper Semi-analytical Satellite Theory (DSST).
|
| org.orekit.propagation.semianalytical.dsst.utilities |
This package provides utilities for Draper Semi-analytical Satellite Theory (DSST).
|
| org.orekit.time |
This independent package provides classes to handle epochs, time scales,
and to compare instants together.
|
| org.orekit.utils |
This package provides useful objects.
|
| Modifier and Type | Method and Description |
|---|---|
<T extends EventDetector> |
AttitudesSequence.addSwitchingCondition(AttitudeProvider past,
AttitudeProvider future,
T switchEvent,
boolean switchOnIncrease,
boolean switchOnDecrease,
double transitionTime,
AngularDerivativesFilter transitionFilter,
AttitudesSequence.SwitchHandler handler)
Add a switching condition between two attitude providers.
|
<T extends RealFieldElement<T>> |
YawCompensation.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
LofOffsetPointing.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
YawSteering.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
FixedRate.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
InertialProvider.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
CelestialBodyPointed.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
AttitudesSequence.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
LofOffset.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
SpinStabilized.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
AttitudeProvider.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
GroundPointing.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
TabulatedLofOffset.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
TabulatedProvider.getAttitude(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
YawCompensation.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
LofOffsetPointing.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
YawSteering.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
FixedRate.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
InertialProvider.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
CelestialBodyPointed.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AttitudesSequence.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
LofOffset.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
SpinStabilized.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AttitudeProvider.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
GroundPointing.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
TabulatedLofOffset.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
TabulatedProvider.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
<T extends RealFieldElement<T>> |
YawCompensation.getBaseState(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the base system state at given date, without compensation.
|
<T extends RealFieldElement<T>> |
YawSteering.getBaseState(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the base system state at given date, without compensation.
|
Attitude |
YawCompensation.getBaseState(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the base system state at given date, without compensation.
|
Attitude |
YawSteering.getBaseState(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the base system state at given date, without compensation.
|
<T extends RealFieldElement<T>> |
TargetPointing.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
<T extends RealFieldElement<T>> |
YawCompensation.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
<T extends RealFieldElement<T>> |
LofOffsetPointing.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
<T extends RealFieldElement<T>> |
YawSteering.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
<T extends RealFieldElement<T>> |
BodyCenterPointing.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
abstract <T extends RealFieldElement<T>> |
GroundPointing.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
<T extends RealFieldElement<T>> |
NadirPointing.getTargetPV(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
TimeStampedPVCoordinates |
TargetPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
TimeStampedPVCoordinates |
YawCompensation.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
TimeStampedPVCoordinates |
LofOffsetPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
TimeStampedPVCoordinates |
YawSteering.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
TimeStampedPVCoordinates |
BodyCenterPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
abstract TimeStampedPVCoordinates |
GroundPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
TimeStampedPVCoordinates |
NadirPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
<T extends RealFieldElement<T>> |
YawCompensation.getYawAngle(FieldPVCoordinatesProvider<T> pvProv,
FieldAbsoluteDate<T> date,
Frame frame)
Compute the yaw compensation angle at date.
|
double |
YawCompensation.getYawAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the yaw compensation angle at date.
|
Attitude |
Attitude.interpolate(AbsoluteDate interpolationDate,
Stream<Attitude> sample)
Get an interpolated instance.
|
FieldAttitude<T> |
FieldAttitude.interpolate(FieldAbsoluteDate<T> interpolationDate,
Stream<FieldAttitude<T>> sample)
Get an interpolated instance.
|
void |
AttitudesSequence.SwitchHandler.switchOccurred(AttitudeProvider preceding,
AttitudeProvider following,
SpacecraftState state)
Method called when attitude is switched from one law to another law.
|
Attitude |
Attitude.withReferenceFrame(Frame newReferenceFrame)
Get a similar attitude with a specific reference frame.
|
FieldAttitude<T> |
FieldAttitude.withReferenceFrame(Frame newReferenceFrame)
Get a similar attitude with a specific reference frame.
|
| Constructor and Description |
|---|
BodyCenterPointing(Frame inertialFrame,
Ellipsoid shape)
Creates new instance.
|
GroundPointing(Frame inertialFrame,
Frame bodyFrame)
Default constructor.
|
LofOffset(Frame inertialFrame,
LOFType type)
Create a LOF-aligned attitude.
|
LofOffset(Frame inertialFrame,
LOFType type,
RotationOrder order,
double alpha1,
double alpha2,
double alpha3)
Creates new instance.
|
LofOffsetPointing(Frame inertialFrame,
BodyShape shape,
AttitudeProvider attLaw,
Vector3D satPointingVector)
Creates new instance.
|
NadirPointing(Frame inertialFrame,
BodyShape shape)
Creates new instance.
|
TabulatedLofOffset(Frame inertialFrame,
LOFType type,
List<TimeStampedAngularCoordinates> table,
int n,
AngularDerivativesFilter filter)
Creates new instance.
|
TargetPointing(Frame inertialFrame,
Frame bodyFrame,
Vector3D target)
Creates a new instance from body frame and target expressed in Cartesian coordinates.
|
TargetPointing(Frame inertialFrame,
GeodeticPoint targetGeo,
BodyShape shape)
Creates a new instance from body shape and target expressed in geodetic coordinates.
|
YawCompensation(Frame inertialFrame,
GroundPointing groundPointingLaw)
Creates a new instance.
|
YawSteering(Frame inertialFrame,
GroundPointing groundPointingLaw,
PVCoordinatesProvider sun,
Vector3D phasingAxis)
Creates a new instance.
|
| Modifier and Type | Method and Description |
|---|---|
static void |
CelestialBodyFactory.addDefaultCelestialBodyLoader(String supportedNames)
Add the default loaders for all predefined celestial bodies.
|
static void |
CelestialBodyFactory.addDefaultCelestialBodyLoader(String name,
String supportedNames)
Add the default loaders for celestial bodies.
|
static CelestialBody |
CelestialBodyFactory.getBody(String name)
Get a celestial body.
|
Frame |
CelestialBody.getBodyOrientedFrame()
Get a body oriented, body centered frame.
|
static CelestialBody |
CelestialBodyFactory.getEarth()
Get the Earth singleton body.
|
static CelestialBody |
CelestialBodyFactory.getEarthMoonBarycenter()
Get the Earth-Moon barycenter singleton bodies pair.
|
Frame |
CelestialBody.getInertiallyOrientedFrame()
Get an inertially oriented, body centered frame.
|
<T extends RealFieldElement<T>> |
BodyShape.getIntersectionPoint(FieldLine<T> line,
FieldVector3D<T> close,
Frame frame,
FieldAbsoluteDate<T> date)
Get the intersection point of a line with the surface of the body.
|
<T extends RealFieldElement<T>> |
OneAxisEllipsoid.getIntersectionPoint(FieldLine<T> line,
FieldVector3D<T> close,
Frame frame,
FieldAbsoluteDate<T> date)
Get the intersection point of a line with the surface of the body.
|
GeodeticPoint |
BodyShape.getIntersectionPoint(Line line,
Vector3D close,
Frame frame,
AbsoluteDate date)
Get the intersection point of a line with the surface of the body.
|
GeodeticPoint |
OneAxisEllipsoid.getIntersectionPoint(Line line,
Vector3D close,
Frame frame,
AbsoluteDate date)
Get the intersection point of a line with the surface of the body.
|
static CelestialBody |
CelestialBodyFactory.getJupiter()
Get the Jupiter singleton body.
|
double |
JPLEphemeridesLoader.getLoadedAstronomicalUnit()
Get astronomical unit.
|
double |
JPLEphemeridesLoader.getLoadedConstant(String... names)
Get a constant defined in the ephemerides headers.
|
double |
JPLEphemeridesLoader.getLoadedEarthMoonMassRatio()
Get Earth/Moon mass ratio.
|
double |
JPLEphemeridesLoader.getLoadedGravitationalCoefficient(JPLEphemeridesLoader.EphemerisType body)
Get the gravitational coefficient of a body.
|
static CelestialBody |
CelestialBodyFactory.getMars()
Get the Mars singleton body.
|
static CelestialBody |
CelestialBodyFactory.getMercury()
Get the Mercury singleton body.
|
static CelestialBody |
CelestialBodyFactory.getMoon()
Get the Moon singleton body.
|
static CelestialBody |
CelestialBodyFactory.getNeptune()
Get the Neptune singleton body.
|
static CelestialBody |
CelestialBodyFactory.getPluto()
Get the Pluto singleton body.
|
<T extends RealFieldElement<T>> |
CelestialBody.getPVCoordinates(FieldAbsoluteDate<T> date,
Frame frame)
Get the
TimeStampedFieldPVCoordinates of the body in the selected frame. |
PVCoordinates |
JPLEphemeridesLoader.RawPVProvider.getRawPV(AbsoluteDate date)
Get the position-velocity at date.
|
<T extends RealFieldElement<T>> |
JPLEphemeridesLoader.RawPVProvider.getRawPV(FieldAbsoluteDate<T> date)
Get the position-velocity at date.
|
static CelestialBody |
CelestialBodyFactory.getSaturn()
Get the Saturn singleton body.
|
static CelestialBody |
CelestialBodyFactory.getSolarSystemBarycenter()
Get the solar system barycenter aggregated body.
|
static CelestialBody |
CelestialBodyFactory.getSun()
Get the Sun singleton body.
|
static CelestialBody |
CelestialBodyFactory.getUranus()
Get the Uranus singleton body.
|
static CelestialBody |
CelestialBodyFactory.getVenus()
Get the Venus singleton body.
|
CelestialBody |
CelestialBodyLoader.loadCelestialBody(String name)
Load celestial body.
|
CelestialBody |
JPLEphemeridesLoader.loadCelestialBody(String name)
Load celestial body.
|
Vector3D |
Ellipsoid.pointOnLimb(Vector3D observer,
Vector3D outside)
Find a point on ellipsoid limb, as seen by an external observer.
|
TimeStampedPVCoordinates |
BodyShape.projectToGround(TimeStampedPVCoordinates pv,
Frame frame)
Project a moving point to the ground.
|
TimeStampedPVCoordinates |
OneAxisEllipsoid.projectToGround(TimeStampedPVCoordinates pv,
Frame frame)
Project a moving point to the ground.
|
Vector3D |
BodyShape.projectToGround(Vector3D point,
AbsoluteDate date,
Frame frame)
Project a point to the ground.
|
Vector3D |
OneAxisEllipsoid.projectToGround(Vector3D point,
AbsoluteDate date,
Frame frame)
Project a point to the ground.
|
<T extends RealFieldElement<T>> |
BodyShape.transform(FieldVector3D<T> point,
Frame frame,
FieldAbsoluteDate<T> date)
Transform a Cartesian point to a surface-relative point.
|
<T extends RealFieldElement<T>> |
OneAxisEllipsoid.transform(FieldVector3D<T> point,
Frame frame,
FieldAbsoluteDate<T> date)
Transform a Cartesian point to a surface-relative point.
|
FieldGeodeticPoint<DerivativeStructure> |
OneAxisEllipsoid.transform(PVCoordinates point,
Frame frame,
AbsoluteDate date)
Transform a Cartesian point to a surface-relative point.
|
GeodeticPoint |
BodyShape.transform(Vector3D point,
Frame frame,
AbsoluteDate date)
Transform a Cartesian point to a surface-relative point.
|
GeodeticPoint |
OneAxisEllipsoid.transform(Vector3D point,
Frame frame,
AbsoluteDate date)
Transform a Cartesian point to a surface-relative point.
|
| Constructor and Description |
|---|
JPLEphemeridesLoader(String supportedNames,
JPLEphemeridesLoader.EphemerisType generateType)
Create a loader for JPL ephemerides binary files.
|
| Modifier and Type | Method and Description |
|---|---|
void |
DataProvidersManager.addDefaultProviders()
Add the default providers configuration.
|
S |
SimpleTimeStampedTableParser.RowConverter.convert(double[] rawFields)
Convert a row.
|
boolean |
NetworkCrawler.feed(Pattern supported,
DataLoader visitor)
Feed a data file loader by browsing the data collection.
|
boolean |
ZipJarCrawler.feed(Pattern supported,
DataLoader visitor)
Feed a data file loader by browsing the data collection.
|
boolean |
DataProvider.feed(Pattern supported,
DataLoader visitor)
Feed a data file loader by browsing the data collection.
|
boolean |
DirectoryCrawler.feed(Pattern supported,
DataLoader visitor)
Feed a data file loader by browsing the data collection.
|
boolean |
ClasspathCrawler.feed(Pattern supported,
DataLoader visitor)
Feed a data file loader by browsing the data collection.
|
boolean |
DataProvidersManager.feed(String supportedNames,
DataLoader loader)
Feed a data file loader by browsing all data providers.
|
void |
DataLoader.loadData(InputStream input,
String name)
Load data from a stream.
|
List<T> |
SimpleTimeStampedTableParser.parse(InputStream stream,
String name)
Parse a stream.
|
PoissonSeries |
PoissonSeriesParser.parse(InputStream stream,
String name)
Parse a stream.
|
PoissonSeriesParser |
PoissonSeriesParser.withDoodson(int firstMultiplierColumn,
int numberColumn)
Set up columns for Doodson multiplers and Doodson number.
|
PoissonSeriesParser |
PoissonSeriesParser.withGamma(int column)
Set up column of GMST tide multiplier.
|
| Constructor and Description |
|---|
ClasspathCrawler(ClassLoader classLoader,
String... list)
Build a data classpath crawler.
|
ClasspathCrawler(String... list)
Build a data classpath crawler.
|
DirectoryCrawler(File root)
Build a data files crawler.
|
FundamentalNutationArguments(IERSConventions conventions,
TimeScale timeScale,
InputStream stream,
String name)
Build a model of fundamental arguments from an IERS table file.
|
FundamentalNutationArguments(IERSConventions conventions,
TimeScale timeScale,
List<double[]> coefficients)
Build a model of fundamental arguments from an IERS table file.
|
ZipJarCrawler(ClassLoader classLoader,
String resource)
Build a zip crawler for an archive file in classpath.
|
ZipJarCrawler(String resource)
Build a zip crawler for an archive file in classpath.
|
ZipJarCrawler(URL url)
Build a zip crawler for an archive file on network.
|
| Modifier and Type | Class and Description |
|---|---|
class |
FrameAncestorException
This class is the base class for exception thrown by
the
UpdatableFrame.updateTransform method. |
class |
TimeStampedCacheException
This class is the base class for all specific exceptions thrown by
during the
GenericTimeStampedCache. |
| Modifier and Type | Method and Description |
|---|---|
OrekitException |
OrekitExceptionWrapper.getException()
Get the wrapped exception.
|
static OrekitException |
OrekitException.unwrap(MathRuntimeException exception)
Recover a OrekitException, possibly embedded in a
MathRuntimeException. |
| Modifier and Type | Method and Description |
|---|---|
static TimeStampedCacheException |
TimeStampedCacheException.unwrap(OrekitException oe)
Recover a TimeStampedCacheException, possibly embedded in a
OrekitException. |
| Constructor and Description |
|---|
OrekitException(OrekitException exception)
Copy constructor.
|
OrekitExceptionWrapper(OrekitException wrappedException)
Simple constructor.
|
TimeStampedCacheException(OrekitException exception)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
void |
BatchLSEstimator.addMeasurement(ObservedMeasurement<?> measurement)
Add a measurement.
|
NumericalPropagator[] |
BatchLSEstimator.estimate()
Estimate the orbital, propagation and measurements parameters.
|
void |
BatchLSObserver.evaluationPerformed(int iterationsCount,
int evaluationsCount,
Orbit[] orbits,
ParameterDriversList estimatedOrbitalParameters,
ParameterDriversList estimatedPropagatorParameters,
ParameterDriversList estimatedMeasurementsParameters,
EstimationsProvider evaluationsProvider,
LeastSquaresProblem.Evaluation lspEvaluation)
Notification callback for the end of each evaluation.
|
ParameterDriversList |
BatchLSEstimator.getMeasurementsParametersDrivers(boolean estimatedOnly)
Get the measurements parameters supported by this estimator (including measurements and modifiers).
|
ParameterDriversList |
BatchLSEstimator.getOrbitalParametersDrivers(boolean estimatedOnly)
Get the orbital parameters supported by this estimator.
|
ParameterDriversList |
BatchLSEstimator.getPropagatorParametersDrivers(boolean estimatedOnly)
Get the propagator parameters supported by this estimator.
|
| Constructor and Description |
|---|
BatchLSEstimator(LeastSquaresOptimizer optimizer,
NumericalPropagatorBuilder... propagatorBuilder)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
void |
AbstractMeasurement.addModifier(EstimationModifier<T> modifier)
Add a modifier.
|
void |
ObservedMeasurement.addModifier(EstimationModifier<T> modifier)
Add a modifier.
|
EstimatedMeasurement<T> |
AbstractMeasurement.estimate(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value of the measurement.
|
EstimatedMeasurement<T> |
ObservedMeasurement.estimate(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value of the measurement.
|
EstimatedMeasurement<?> |
EstimationsProvider.getEstimatedMeasurement(int index)
Get one estimated measurement.
|
GeodeticPoint |
GroundStation.getOffsetGeodeticPoint()
Get the geodetic point at the center of the offset frame.
|
Transform |
GroundStation.getOffsetToInertial(Frame inertial,
AbsoluteDate date)
Get the transform between offset frame and inertial frame.
|
FieldTransform<DerivativeStructure> |
GroundStation.getOffsetToInertial(Frame inertial,
FieldAbsoluteDate<DerivativeStructure> date,
DSFactory factory,
Map<String,Integer> indices)
Get the transform between offset frame and inertial frame with derivatives.
|
void |
EstimationModifier.modify(EstimatedMeasurement<T> estimated)
Apply a modifier to an estimated measurement.
|
protected EstimatedMeasurement<AngularRaDec> |
AngularRaDec.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected EstimatedMeasurement<Range> |
Range.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected EstimatedMeasurement<PV> |
PV.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected EstimatedMeasurement<RangeRate> |
RangeRate.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected EstimatedMeasurement<AngularAzEl> |
AngularAzEl.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected EstimatedMeasurement<TurnAroundRange> |
TurnAroundRange.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected abstract EstimatedMeasurement<T> |
AbstractMeasurement.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
protected EstimatedMeasurement<InterSatellitesRange> |
InterSatellitesRange.theoreticalEvaluation(int iteration,
int evaluation,
SpacecraftState[] states)
Estimate the theoretical value.
|
| Constructor and Description |
|---|
AngularAzEl(GroundStation station,
AbsoluteDate date,
double[] angular,
double[] sigma,
double[] baseWeight)
Simple constructor.
|
AngularAzEl(GroundStation station,
AbsoluteDate date,
double[] angular,
double[] sigma,
double[] baseWeight,
int propagatorIndex)
Simple constructor.
|
AngularRaDec(GroundStation station,
Frame referenceFrame,
AbsoluteDate date,
double[] angular,
double[] sigma,
double[] baseWeight)
Simple constructor.
|
AngularRaDec(GroundStation station,
Frame referenceFrame,
AbsoluteDate date,
double[] angular,
double[] sigma,
double[] baseWeight,
int propagatorIndex)
Simple constructor.
|
GroundStation(TopocentricFrame baseFrame)
Simple constructor.
|
InterSatellitesRange(int satellite1Index,
int satellite2Index,
boolean twoWay,
AbsoluteDate date,
double range,
double sigma,
double baseWeight)
Simple constructor.
|
Range(GroundStation station,
AbsoluteDate date,
double range,
double sigma,
double baseWeight)
Simple constructor.
|
Range(GroundStation station,
AbsoluteDate date,
double range,
double sigma,
double baseWeight,
int propagatorIndex)
Simple constructor.
|
RangeRate(GroundStation station,
AbsoluteDate date,
double rangeRate,
double sigma,
double baseWeight,
boolean twoway)
Simple constructor.
|
RangeRate(GroundStation station,
AbsoluteDate date,
double rangeRate,
double sigma,
double baseWeight,
boolean twoway,
int propagatorIndex)
Simple constructor.
|
TurnAroundRange(GroundStation masterStation,
GroundStation slaveStation,
AbsoluteDate date,
double turnAroundRange,
double sigma,
double baseWeight)
Simple constructor.
|
TurnAroundRange(GroundStation masterStation,
GroundStation slaveStation,
AbsoluteDate date,
double turnAroundRange,
double sigma,
double baseWeight,
int propagatorIndex)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
void |
AngularIonosphericDelayModifier.modify(EstimatedMeasurement<AngularAzEl> estimated) |
void |
AngularRadioRefractionModifier.modify(EstimatedMeasurement<AngularAzEl> estimated) |
void |
AngularTroposphericDelayModifier.modify(EstimatedMeasurement<AngularAzEl> estimated) |
void |
RangeIonosphericDelayModifier.modify(EstimatedMeasurement<Range> estimated) |
void |
RangeTroposphericDelayModifier.modify(EstimatedMeasurement<Range> estimated)
Apply a modifier to an estimated measurement.
|
void |
RangeRateIonosphericDelayModifier.modify(EstimatedMeasurement<RangeRate> estimated)
Apply a modifier to an estimated measurement.
|
void |
RangeRateTroposphericDelayModifier.modify(EstimatedMeasurement<RangeRate> estimated)
Apply a modifier to an estimated measurement.
|
void |
TurnAroundRangeTroposphericDelayModifier.modify(EstimatedMeasurement<TurnAroundRange> estimated)
Apply a modifier to an estimated measurement.
|
void |
TurnAroundRangeIonosphericDelayModifier.modify(EstimatedMeasurement<TurnAroundRange> estimated) |
double |
RangeRateTroposphericDelayModifier.rangeRateErrorTroposphericModel(GroundStation station,
SpacecraftState state)
Compute the measurement error due to Troposphere.
|
| Constructor and Description |
|---|
Bias(String[] name,
double[] bias,
double[] scale,
double[] min,
double[] max)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
void |
TDMFile.checkTimeSystems()
Check that, according to the CCSDS standard, every ObservationsBlock has the same time system.
|
CartesianOrbit |
OPMFile.generateCartesianOrbit()
Generate a
CartesianOrbit from the OPM state vector data. |
CartesianOrbit |
OMMFile.generateCartesianOrbit()
Generate a
CartesianOrbit from the KeplerianOrbit. |
KeplerianOrbit |
OPMFile.generateKeplerianOrbit()
Generate a
KeplerianOrbit from the OPM Keplerian elements if hasKeplerianElements is true,
or from the state vector data otherwise. |
KeplerianOrbit |
OMMFile.generateKeplerianOrbit()
Generate a
KeplerianOrbit based on the OMM mean Keplerian elements. |
SpacecraftState |
OPMFile.generateSpacecraftState()
Generate spacecraft state from the
CartesianOrbit generated by generateCartesianOrbit. |
SpacecraftState |
OMMFile.generateSpacecraftState()
Generate spacecraft state from the
KeplerianOrbit generated by generateKeplerianOrbit. |
abstract CelestialBody |
CenterName.getCelestialBody()
Get the celestial body corresponding to the CCSDS constant.
|
IERSConventions |
ODMFile.getConventions()
Get IERS conventions.
|
Frame |
ODMMetaData.getFrame()
Get the reference frame in which data are given: used for state vector and
Keplerian elements data (and for the covariance reference frame if none is given).
|
Frame |
OEMFile.EphemeridesBlock.getFrame() |
Frame |
CCSDSFrame.getFrame(IERSConventions conventions,
boolean simpleEOP)
Get the frame corresponding to the CCSDS constant.
|
double |
OGMFile.getMass()
Get the spacecraft mass.
|
TimeScale |
OEMFile.EphemeridesBlock.getTimeScale() |
abstract TimeScale |
CcsdsTimeScale.getTimeScale(IERSConventions conventions)
Get the corresponding
TimeScale. |
void |
StreamingOemWriter.Segment.handleStep(SpacecraftState s,
boolean isLast) |
void |
StreamingOemWriter.Segment.init(SpacecraftState s0,
AbsoluteDate t,
double step)
Initialize step handler at the start of a propagation.
|
OEMFile |
OEMParser.parse(BufferedReader reader,
String fileName) |
OMMFile |
OMMParser.parse(InputStream stream)
Parse a CCSDS Orbit Data Message.
|
OPMFile |
OPMParser.parse(InputStream stream)
Parse a CCSDS Orbit Data Message.
|
TDMFile |
TDMParser.parse(InputStream stream)
Parse a CCSDS Tracking Data Message.
|
ODMFile |
ODMParser.parse(InputStream stream)
Parse a CCSDS Orbit Data Message.
|
OEMFile |
OEMParser.parse(InputStream stream)
Parse a CCSDS Orbit Data Message.
|
OMMFile |
OMMParser.parse(InputStream stream,
String fileName)
Parse a CCSDS Orbit Data Message.
|
OPMFile |
OPMParser.parse(InputStream stream,
String fileName)
Parse a CCSDS Orbit Data Message.
|
TDMFile |
TDMParser.parse(InputStream stream,
String fileName)
Parse a CCSDS Tracking Data Message.
|
abstract ODMFile |
ODMParser.parse(InputStream stream,
String fileName)
Parse a CCSDS Orbit Data Message.
|
OEMFile |
OEMParser.parse(InputStream stream,
String fileName)
Parse a CCSDS Orbit Data Message.
|
OMMFile |
OMMParser.parse(String fileName)
Parse a CCSDS Orbit Data Message.
|
OPMFile |
OPMParser.parse(String fileName)
Parse a CCSDS Orbit Data Message.
|
TDMFile |
TDMParser.parse(String fileName)
Parse a CCSDS Tracking Data Message.
|
ODMFile |
ODMParser.parse(String fileName)
Parse a CCSDS Orbit Data Message.
|
OEMFile |
OEMParser.parse(String fileName)
Parse a CCSDS Orbit Data Message.
|
protected AbsoluteDate |
ODMParser.parseDate(String date,
CcsdsTimeScale timeSystem)
Parse a date.
|
AbsoluteDate |
CcsdsTimeScale.parseDate(String date,
IERSConventions conventions,
AbsoluteDate missionReferenceDate)
Parse a date in this time scale.
|
protected boolean |
ODMParser.parseGeneralStateDataEntry(org.orekit.files.ccsds.KeyValue keyValue,
OGMFile general,
List<String> comment)
Parse a general state data key = value entry.
|
protected boolean |
ODMParser.parseHeaderEntry(org.orekit.files.ccsds.KeyValue keyValue,
ODMFile odmFile,
List<String> comment)
Parse an entry from the header.
|
TDMFile |
TDMParser.parseKeyValue(InputStream stream,
String fileName)
Parse a CCSDS Tracking Data Message with KEYVALUE format.
|
protected boolean |
ODMParser.parseMetaDataEntry(org.orekit.files.ccsds.KeyValue keyValue,
ODMMetaData metaData,
List<String> comment)
Parse a meta-data key = value entry.
|
TDMFile |
TDMParser.parseXml(InputStream stream,
String fileName)
Parse a CCSDS Tracking Data Message with XML format.
|
protected void |
ODMFile.setMuUsed()
Set the gravitational coefficient created from the knowledge of the central body.
|
void |
OEMWriter.write(Appendable writer,
EphemerisFile ephemerisFile)
Write the passed in
EphemerisFile using the passed in
Appendable. |
| Constructor and Description |
|---|
StreamingOemWriter(Appendable writer,
TimeScale timeScale,
Map<Keyword,String> metadata)
Create an OEM writer than streams data to the given output stream.
|
| Modifier and Type | Method and Description |
|---|---|
OrekitEphemerisFile.OrekitEphemerisSegment |
OrekitEphemerisFile.OrekitSatelliteEphemeris.addNewSegment(List<SpacecraftState> states)
Injects pre-computed satellite states into this ephemeris file
object, returning the generated
OrekitEphemerisFile.OrekitEphemerisSegment that
has been stored internally. |
OrekitEphemerisFile.OrekitEphemerisSegment |
OrekitEphemerisFile.OrekitSatelliteEphemeris.addNewSegment(List<SpacecraftState> states,
CelestialBody body,
int interpolationSampleSize)
Injects pre-computed satellite states into this ephemeris file
object, returning the generated
OrekitEphemerisFile.OrekitEphemerisSegment that
has been stored internally. |
OrekitEphemerisFile.OrekitEphemerisSegment |
OrekitEphemerisFile.OrekitSatelliteEphemeris.addNewSegment(List<SpacecraftState> states,
int interpolationSampleSize)
Injects pre-computed satellite states into this ephemeris file
object, returning the generated
OrekitEphemerisFile.OrekitEphemerisSegment that
has been stored internally. |
Frame |
OrekitEphemerisFile.OrekitEphemerisSegment.getFrame() |
Frame |
EphemerisFile.EphemerisSegment.getFrame()
Get the reference frame for this ephemeris segment.
|
default BoundedPropagator |
EphemerisFile.SatelliteEphemeris.getPropagator()
View this ephemeris as a propagator, combining data from all
segments. |
default BoundedPropagator |
EphemerisFile.EphemerisSegment.getPropagator()
View this ephemeris segment as a propagator.
|
TimeScale |
OrekitEphemerisFile.OrekitEphemerisSegment.getTimeScale() |
TimeScale |
EphemerisFile.EphemerisSegment.getTimeScale()
Get the time scale for this ephemeris segment.
|
EphemerisFile |
EphemerisFileParser.parse(BufferedReader reader,
String fileName)
Parse an ephemeris file from a stream.
|
EphemerisFile |
EphemerisFileParser.parse(String fileName)
Parse an ephemeris file from a file on the local file system.
|
void |
EphemerisFileWriter.write(Appendable writer,
EphemerisFile ephemerisFile)
Write the passed in
EphemerisFile using the passed in
Appendable. |
default void |
EphemerisFileWriter.write(String outputFilePath,
EphemerisFile ephemerisFile)
Write the passed in
EphemerisFile to a file at the output path
specified. |
| Modifier and Type | Method and Description |
|---|---|
Frame |
SP3File.SP3Ephemeris.getFrame() |
BoundedPropagator |
SP3File.SP3Ephemeris.getPropagator() |
TimeScale |
SP3File.SP3Ephemeris.getTimeScale() |
SP3File |
SP3Parser.parse(BufferedReader reader,
String fileName) |
SP3File |
SP3Parser.parse(InputStream stream)
Parse a SP3 file from an input stream using the UTF-8 charset.
|
SP3File |
SP3Parser.parse(String fileName) |
| Modifier and Type | Method and Description |
|---|---|
<T extends RealFieldElement<T>> |
AbstractParametricAcceleration.acceleration(FieldSpacecraftState<T> state,
T[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
ForceModel.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
Vector3D |
AbstractParametricAcceleration.acceleration(SpacecraftState state,
double[] parameters)
Compute acceleration.
|
Vector3D |
ForceModel.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
default <T extends RealFieldElement<T>> |
ForceModel.addContribution(FieldSpacecraftState<T> s,
FieldTimeDerivativesEquations<T> adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
default void |
ForceModel.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
protected void |
AbstractForceModel.complainIfNotSupported(String name)
Complain if a parameter is not supported.
|
Vector3D |
BoxAndSolarArraySpacecraft.dragAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
double density,
Vector3D relativeVelocity,
double[] parameters)
Compute the acceleration due to drag.
|
FieldVector3D<DerivativeStructure> |
BoxAndSolarArraySpacecraft.dragAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
double density,
Vector3D relativeVelocity,
double[] parameters,
String paramName)
Compute acceleration due to drag, with parameters derivatives.
|
<T extends RealFieldElement<T>> |
BoxAndSolarArraySpacecraft.dragAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
T density,
FieldVector3D<T> relativeVelocity,
T[] parameters)
Compute the acceleration due to drag.
|
FieldVector3D<DerivativeStructure> |
BoxAndSolarArraySpacecraft.getNormal(AbsoluteDate date,
Frame frame,
FieldVector3D<DerivativeStructure> position,
FieldRotation<DerivativeStructure> rotation)
Get solar array normal in spacecraft frame.
|
Vector3D |
BoxAndSolarArraySpacecraft.getNormal(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation)
Get solar array normal in spacecraft frame.
|
<T extends RealFieldElement<T>> |
BoxAndSolarArraySpacecraft.getNormal(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation)
Get solar array normal in spacecraft frame.
|
ParameterDriver |
ForceModel.getParameterDriver(String name)
Get parameter value from its name.
|
ParameterDriver |
AbstractForceModel.getParameterDriver(String name)
Get parameter value from its name.
|
void |
HarmonicParametricAcceleration.init(SpacecraftState initialState,
AbsoluteDate target)
Initialize the force model at the start of propagation.
|
default void |
ForceModel.init(SpacecraftState initialState,
AbsoluteDate target)
Initialize the force model at the start of propagation.
|
void |
PolynomialParametricAcceleration.init(SpacecraftState initialState,
AbsoluteDate target)
Initialize the force model at the start of propagation.
|
Vector3D |
BoxAndSolarArraySpacecraft.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters)
Compute the acceleration due to radiation pressure.
|
FieldVector3D<DerivativeStructure> |
BoxAndSolarArraySpacecraft.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters,
String paramName)
Compute the acceleration due to radiation pressure, with parameters derivatives.
|
<T extends RealFieldElement<T>> |
BoxAndSolarArraySpacecraft.radiationPressureAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
FieldVector3D<T> flux,
T[] parameters)
Compute the acceleration due to radiation pressure.
|
| Modifier and Type | Method and Description |
|---|---|
<T extends RealFieldElement<T>> |
DragForce.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
Vector3D |
DragForce.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
Vector3D |
DragSensitive.dragAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
double density,
Vector3D relativeVelocity,
double[] parameters)
Compute the acceleration due to drag.
|
FieldVector3D<DerivativeStructure> |
IsotropicDrag.dragAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
double density,
Vector3D relativeVelocity,
double[] parameters,
String paramName)
Compute acceleration due to drag, with parameters derivatives.
|
FieldVector3D<DerivativeStructure> |
DragSensitive.dragAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
double density,
Vector3D relativeVelocity,
double[] parameters,
String paramName)
Compute acceleration due to drag, with parameters derivatives.
|
<T extends RealFieldElement<T>> |
IsotropicDrag.dragAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
T density,
FieldVector3D<T> relativeVelocity,
T[] parameters)
Compute the acceleration due to drag.
|
<T extends RealFieldElement<T>> |
DragSensitive.dragAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
T density,
FieldVector3D<T> relativeVelocity,
T[] parameters)
Compute the acceleration due to drag.
|
| Modifier and Type | Method and Description |
|---|---|
double |
DTM2000InputParameters.get24HoursKp(AbsoluteDate date)
Get the last 24H mean geomagnetic index.
|
double[] |
NRLMSISE00InputParameters.getAp(AbsoluteDate date)
Get the Ap geomagnetic indices.
|
double |
NRLMSISE00InputParameters.getAverageFlux(AbsoluteDate date)
Get the value of the 81 day average of F10.7 solar flux centered on current day.
|
double |
NRLMSISE00InputParameters.getDailyFlux(AbsoluteDate date)
Get the value of the daily F10.7 solar flux for previous day.
|
double |
JB2008.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
Atmosphere.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
DTM2000.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
NRLMSISE00.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
SimpleExponentialAtmosphere.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
HarrisPriester.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density at some position.
|
double |
JB2008.getDensity(double dateMJD,
double sunRA,
double sunDecli,
double satLon,
double satLat,
double satAlt,
double f10,
double f10B,
double s10,
double s10B,
double xm10,
double xm10B,
double y10,
double y10B,
double dstdtc)
Get the local density with initial entries.
|
<T extends RealFieldElement<T>> |
JB2008.getDensity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame)
Get the local density.
|
<T extends RealFieldElement<T>> |
Atmosphere.getDensity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame)
Get the local density.
|
<T extends RealFieldElement<T>> |
DTM2000.getDensity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame)
Get the local density.
|
<T extends RealFieldElement<T>> |
NRLMSISE00.getDensity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame)
Get the local density.
|
<T extends RealFieldElement<T>> |
SimpleExponentialAtmosphere.getDensity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame) |
<T extends RealFieldElement<T>> |
HarrisPriester.getDensity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame)
Get the local density at some position.
|
double |
DTM2000.getDensity(int day,
double alti,
double lon,
double lat,
double hl,
double f,
double fbar,
double akp3,
double akp24)
Get the local density with initial entries.
|
<T extends RealFieldElement<T>> |
DTM2000.getDensity(int day,
T alti,
T lon,
T lat,
T hl,
double f,
double fbar,
double akp3,
double akp24)
Get the local density with initial entries.
|
<T extends RealFieldElement<T>> |
JB2008.getDensity(T dateMJD,
T sunRA,
T sunDecli,
T satLon,
T satLat,
T satAlt,
double f10,
double f10B,
double s10,
double s10B,
double xm10,
double xm10B,
double y10,
double y10B,
double dstdtc)
Get the local density with initial entries.
|
<T extends RealFieldElement<T>> |
HarrisPriester.getDensity(Vector3D sunInEarth,
FieldVector3D<T> posInEarth)
Get the local density.
|
double |
HarrisPriester.getDensity(Vector3D sunInEarth,
Vector3D posInEarth)
Get the local density.
|
double |
JB2008InputParameters.getDSTDTC(AbsoluteDate date)
Get the temperature change computed from Dst index.
|
double |
JB2008InputParameters.getF10(AbsoluteDate date)
Get the value of the instantaneous solar flux index
(1e-22*Watt/(m²*Hertz)).
|
double |
JB2008InputParameters.getF10B(AbsoluteDate date)
Get the value of the mean solar flux.
|
double |
DTM2000InputParameters.getInstantFlux(AbsoluteDate date)
Get the value of the instantaneous solar flux.
|
AbsoluteDate |
NRLMSISE00InputParameters.getMaxDate()
Gets the available data range maximum date.
|
AbsoluteDate |
DTM2000InputParameters.getMaxDate()
Gets the available data range maximum date.
|
double |
DTM2000InputParameters.getMeanFlux(AbsoluteDate date)
Get the value of the mean solar flux.
|
AbsoluteDate |
NRLMSISE00InputParameters.getMinDate()
Gets the available data range minimum date.
|
AbsoluteDate |
DTM2000InputParameters.getMinDate()
Gets the available data range minimum date.
|
double |
JB2008InputParameters.getS10(AbsoluteDate date)
Get the EUV index (26-34 nm) scaled to F10.
|
double |
JB2008InputParameters.getS10B(AbsoluteDate date)
Get the EUV 81-day averaged centered index.
|
double |
DTM2000InputParameters.getThreeHourlyKP(AbsoluteDate date)
Get the value of the 3 hours geomagnetic index.
|
default Vector3D |
Atmosphere.getVelocity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the inertial velocity of atmosphere molecules.
|
default <T extends RealFieldElement<T>> |
Atmosphere.getVelocity(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
Frame frame)
Get the inertial velocity of atmosphere molecules.
|
double |
JB2008InputParameters.getXM10(AbsoluteDate date)
Get the MG2 index scaled to F10.
|
double |
JB2008InputParameters.getXM10B(AbsoluteDate date)
Get the MG2 81-day average centered index.
|
double |
JB2008InputParameters.getY10(AbsoluteDate date)
Get the Solar X-Ray & Lya index scaled to F10.
|
double |
JB2008InputParameters.getY10B(AbsoluteDate date)
Get the Solar X-Ray & Lya 81-day ave.
|
NRLMSISE00 |
NRLMSISE00.withSwitch(int number,
int value)
Change a switch.
|
| Constructor and Description |
|---|
DTM2000(DTM2000InputParameters parameters,
PVCoordinatesProvider sun,
BodyShape earth)
Simple constructor for independent computation.
|
| Modifier and Type | Method and Description |
|---|---|
double |
MarshallSolarActivityFutureEstimation.get24HoursKp(AbsoluteDate date)
The Kp index is derived from the Ap index.
|
DateComponents |
MarshallSolarActivityFutureEstimation.getFileDate(AbsoluteDate date)
Get the date of the file from which data at the specified date comes from.
|
double |
MarshallSolarActivityFutureEstimation.getInstantFlux(AbsoluteDate date)
Get the value of the instantaneous solar flux.
|
AbsoluteDate |
MarshallSolarActivityFutureEstimation.getMaxDate()
Gets the available data range maximum date.
|
double |
MarshallSolarActivityFutureEstimation.getMeanFlux(AbsoluteDate date)
Get the value of the mean solar flux.
|
AbsoluteDate |
MarshallSolarActivityFutureEstimation.getMinDate()
Gets the available data range minimum date.
|
double |
MarshallSolarActivityFutureEstimation.getThreeHourlyKP(AbsoluteDate date)
Get the value of the 3 hours geomagnetic index.
|
void |
MarshallSolarActivityFutureEstimation.loadData(InputStream input,
String name)
Load data from a stream.
|
| Modifier and Type | Method and Description |
|---|---|
<T extends RealFieldElement<T>> |
HolmesFeatherstoneAttractionModel.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
OceanTides.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
NewtonianAttraction.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
Relativity.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
ThirdBodyAttraction.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
SolidTides.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
Vector3D |
HolmesFeatherstoneAttractionModel.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
Vector3D |
OceanTides.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
Vector3D |
NewtonianAttraction.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
Vector3D |
Relativity.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
Vector3D |
ThirdBodyAttraction.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
Vector3D |
SolidTides.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
NewtonianAttraction.addContribution(FieldSpacecraftState<T> s,
FieldTimeDerivativesEquations<T> adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
NewtonianAttraction.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
double[] |
HolmesFeatherstoneAttractionModel.gradient(AbsoluteDate date,
Vector3D position,
double mu)
Compute the gradient of the non-central part of the gravity field.
|
<T extends RealFieldElement<T>> |
HolmesFeatherstoneAttractionModel.gradient(FieldAbsoluteDate<T> date,
FieldVector3D<T> position,
T mu)
Compute the gradient of the non-central part of the gravity field.
|
double |
HolmesFeatherstoneAttractionModel.nonCentralPart(AbsoluteDate date,
Vector3D position,
double mu)
Compute the non-central part of the gravity field.
|
double |
HolmesFeatherstoneAttractionModel.value(AbsoluteDate date,
Vector3D position,
double mu)
Compute the value of the gravity field.
|
| Constructor and Description |
|---|
OceanTides(Frame centralBodyFrame,
double ae,
double mu,
boolean poleTide,
double step,
int nbPoints,
int degree,
int order,
IERSConventions conventions,
UT1Scale ut1)
Simple constructor.
|
OceanTides(Frame centralBodyFrame,
double ae,
double mu,
int degree,
int order,
IERSConventions conventions,
UT1Scale ut1)
Simple constructor.
|
SolidTides(Frame centralBodyFrame,
double ae,
double mu,
TideSystem centralTideSystem,
boolean poleTide,
double step,
int nbPoints,
IERSConventions conventions,
UT1Scale ut1,
CelestialBody... bodies)
Simple constructor.
|
SolidTides(Frame centralBodyFrame,
double ae,
double mu,
TideSystem centralTideSystem,
IERSConventions conventions,
UT1Scale ut1,
CelestialBody... bodies)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
static void |
GravityFieldFactory.addDefaultOceanTidesReaders()
Add the default READERS for ocean tides.
|
protected void |
OceanTidesReader.addWaveCoefficients(int doodson,
int n,
int m,
double cPlus,
double sPlus,
double cMinus,
double sMinus,
int lineNumber,
String line)
Add parsed coefficients.
|
protected void |
OceanTidesReader.endParse()
End parsing.
|
abstract double[] |
OceanLoadDeformationCoefficients.getCoefficients()
Get the load deformation coefficients for ocean tides.
|
static NormalizedSphericalHarmonicsProvider |
GravityFieldFactory.getConstantNormalizedProvider(int degree,
int order)
Get the constant gravity field coefficients provider from the first supported file.
|
protected org.orekit.forces.gravity.potential.ConstantSphericalHarmonics |
PotentialCoefficientsReader.getConstantProvider(boolean wantNormalized,
int degree,
int order)
Get a time-independent provider for read spherical harmonics coefficients.
|
static UnnormalizedSphericalHarmonicsProvider |
GravityFieldFactory.getConstantUnnormalizedProvider(int degree,
int order)
Get the constant gravity field coefficients provider from the first supported file.
|
double |
NormalizedSphericalHarmonicsProvider.NormalizedSphericalHarmonics.getNormalizedCnm(int n,
int m)
Get a spherical harmonic cosine coefficient.
|
static NormalizedSphericalHarmonicsProvider |
GravityFieldFactory.getNormalizedProvider(int degree,
int order)
Get the gravity field coefficients provider from the first supported file.
|
static NormalizedSphericalHarmonicsProvider |
GravityFieldFactory.getNormalizedProvider(UnnormalizedSphericalHarmonicsProvider unnormalized)
Create a
NormalizedSphericalHarmonicsProvider from an UnnormalizedSphericalHarmonicsProvider. |
double |
NormalizedSphericalHarmonicsProvider.NormalizedSphericalHarmonics.getNormalizedSnm(int n,
int m)
Get a spherical harmonic sine coefficient.
|
static List<OceanTidesWave> |
GravityFieldFactory.getOceanTidesWaves(int degree,
int order)
Get the ocean tides waves from the first supported file.
|
RawSphericalHarmonicsProvider |
EGMFormatReader.getProvider(boolean wantNormalized,
int degree,
int order)
Get a provider for read spherical harmonics coefficients.
|
RawSphericalHarmonicsProvider |
ICGEMFormatReader.getProvider(boolean wantNormalized,
int degree,
int order)
Get a provider for read spherical harmonics coefficients.
|
abstract RawSphericalHarmonicsProvider |
PotentialCoefficientsReader.getProvider(boolean wantNormalized,
int degree,
int order)
Get a provider for read spherical harmonics coefficients.
|
RawSphericalHarmonicsProvider |
SHMFormatReader.getProvider(boolean wantNormalized,
int degree,
int order)
Get a provider for read spherical harmonics coefficients.
|
RawSphericalHarmonicsProvider |
GRGSFormatReader.getProvider(boolean wantNormalized,
int degree,
int order)
Get a provider for read spherical harmonics coefficients.
|
double |
RawSphericalHarmonicsProvider.RawSphericalHarmonics.getRawCnm(int n,
int m)
Get a spherical harmonic cosine coefficient.
|
double |
RawSphericalHarmonicsProvider.RawSphericalHarmonics.getRawSnm(int n,
int m)
Get a spherical harmonic sine coefficient.
|
static double[][] |
GravityFieldFactory.getUnnormalizationFactors(int degree,
int order)
Get a un-normalization factors array.
|
double |
UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics.getUnnormalizedCnm(int n,
int m)
Get a spherical harmonic cosine coefficient.
|
static UnnormalizedSphericalHarmonicsProvider |
GravityFieldFactory.getUnnormalizedProvider(int degree,
int order)
Get the gravity field coefficients provider from the first supported file.
|
static UnnormalizedSphericalHarmonicsProvider |
GravityFieldFactory.getUnnormalizedProvider(NormalizedSphericalHarmonicsProvider normalized)
Create an
UnnormalizedSphericalHarmonicsProvider from a NormalizedSphericalHarmonicsProvider. |
double |
UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics.getUnnormalizedSnm(int n,
int m)
Get a spherical harmonic sine coefficient.
|
void |
EGMFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
ICGEMFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
abstract void |
PotentialCoefficientsReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
AstronomicalAmplitudeReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
FESCnmSnmReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
FESCHatEpsilonReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
SHMFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
GRGSFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
NormalizedSphericalHarmonicsProvider.NormalizedSphericalHarmonics |
NormalizedSphericalHarmonicsProvider.onDate(AbsoluteDate date)
Get the normalized spherical harmonic coefficients at a specific instance in time.
|
RawSphericalHarmonicsProvider.RawSphericalHarmonics |
RawSphericalHarmonicsProvider.onDate(AbsoluteDate date)
Get the raw spherical harmonic coefficients on a specific date.
|
UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics |
UnnormalizedSphericalHarmonicsProvider.onDate(AbsoluteDate date)
Get the un-normalized spherical harmonic coefficients at a specific instance in time.
|
protected void |
PotentialCoefficientsReader.parseCoefficient(String field,
double[][] array,
int i,
int j,
String cName,
String name)
Parse a coefficient.
|
protected void |
PotentialCoefficientsReader.parseCoefficient(String field,
List<List<Double>> list,
int i,
int j,
String cName,
String name)
Parse a coefficient.
|
static PotentialCoefficientsReader |
GravityFieldFactory.readGravityField(int maxParseDegree,
int maxParseOrder)
Read a gravity field coefficients provider from the first supported file.
|
protected static void |
PotentialCoefficientsReader.rescale(double scale,
boolean normalizedOrigin,
double[][] originC,
double[][] originS,
boolean wantNormalized,
double[][] rescaledC,
double[][] rescaledS)
Rescale coefficients arrays.
|
protected void |
PotentialCoefficientsReader.setRawCoefficients(boolean rawNormalized,
double[][] c,
double[][] s,
String name)
Set the tesseral-sectorial coefficients matrix.
|
| Modifier and Type | Method and Description |
|---|---|
<T extends RealFieldElement<T>> |
ConstantThrustManeuver.addContribution(FieldSpacecraftState<T> s,
FieldTimeDerivativesEquations<T> adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
ConstantThrustManeuver.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
double |
ImpulseManeuver.g(SpacecraftState s)
Compute the value of the switching function.
|
void |
SmallManeuverAnalyticalModel.getJacobian(Orbit orbit1,
PositionAngle positionAngle,
double[][] jacobian)
Compute the Jacobian of the orbit with respect to maneuver parameters.
|
| Constructor and Description |
|---|
SmallManeuverAnalyticalModel(SpacecraftState state0,
Frame frame,
Vector3D dV,
double isp)
Build a maneuver defined in user-specified frame.
|
SmallManeuverAnalyticalModel(SpacecraftState state0,
Vector3D dV,
double isp)
Build a maneuver defined in spacecraft frame.
|
| Modifier and Type | Method and Description |
|---|---|
<T extends RealFieldElement<T>> |
SolarRadiationPressure.acceleration(FieldSpacecraftState<T> s,
T[] parameters)
Compute acceleration.
|
Vector3D |
SolarRadiationPressure.acceleration(SpacecraftState s,
double[] parameters)
Compute acceleration.
|
<T extends RealFieldElement<T>> |
SolarRadiationPressure.getLightingRatio(FieldVector3D<T> position,
Frame frame,
FieldAbsoluteDate<T> date)
Get the lighting ratio ([0-1]).
|
double |
SolarRadiationPressure.getLightingRatio(Vector3D position,
Frame frame,
AbsoluteDate date)
Get the lighting ratio ([0-1]).
|
Vector3D |
RadiationSensitive.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters)
Compute the acceleration due to radiation pressure.
|
FieldVector3D<DerivativeStructure> |
IsotropicRadiationSingleCoefficient.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters,
String paramName)
Compute the acceleration due to radiation pressure, with parameters derivatives.
|
FieldVector3D<DerivativeStructure> |
RadiationSensitive.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters,
String paramName)
Compute the acceleration due to radiation pressure, with parameters derivatives.
|
FieldVector3D<DerivativeStructure> |
IsotropicRadiationClassicalConvention.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters,
String paramName)
Compute the acceleration due to radiation pressure, with parameters derivatives.
|
FieldVector3D<DerivativeStructure> |
IsotropicRadiationCNES95Convention.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters,
String paramName)
Compute the acceleration due to radiation pressure, with parameters derivatives.
|
<T extends RealFieldElement<T>> |
IsotropicRadiationSingleCoefficient.radiationPressureAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
FieldVector3D<T> flux,
T[] parameters)
Compute the acceleration due to radiation pressure.
|
<T extends RealFieldElement<T>> |
RadiationSensitive.radiationPressureAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
FieldVector3D<T> flux,
T[] parameters)
Compute the acceleration due to radiation pressure.
|
<T extends RealFieldElement<T>> |
IsotropicRadiationClassicalConvention.radiationPressureAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
FieldVector3D<T> flux,
T[] parameters)
Compute the acceleration due to radiation pressure.
|
<T extends RealFieldElement<T>> |
IsotropicRadiationCNES95Convention.radiationPressureAcceleration(FieldAbsoluteDate<T> date,
Frame frame,
FieldVector3D<T> position,
FieldRotation<T> rotation,
T mass,
FieldVector3D<T> flux,
T[] parameters)
Compute the acceleration due to radiation pressure.
|
| Modifier and Type | Method and Description |
|---|---|
void |
OrphanFrame.addChild(OrphanFrame child,
Transform transform,
boolean isPseudoInertial)
Add a child.
|
void |
OrphanFrame.addChild(OrphanFrame child,
TransformProvider transformProvider,
boolean isPseudoInertial)
Add a child.
|
void |
OrphanFrame.attachTo(Frame parent,
Transform transform,
boolean isPseudoInertial)
Attach the instance (and all its children down to leafs) to the main tree.
|
void |
OrphanFrame.attachTo(Frame parent,
TransformProvider transformProvider,
boolean isPseudoInertial)
Attach the instance (and all its children down to leafs) to the main tree.
|
void |
EOPHistory.checkEOPContinuity(double maxGap)
Check Earth orientation parameters continuity.
|
GeodeticPoint |
TopocentricFrame.computeLimitVisibilityPoint(double radius,
double azimuth,
double elevation)
Compute the limit visibility point for a satellite in a given direction.
|
void |
EOPHistoryLoader.fillHistory(IERSConventions.NutationCorrectionConverter converter,
SortedSet<EOPEntry> history)
Load celestial body.
|
double |
TopocentricFrame.getAzimuth(Vector3D extPoint,
Frame frame,
AbsoluteDate date)
Get the azimuth of a point with regards to the topocentric frame center point.
|
static FactoryManagedFrame |
FramesFactory.getCIRF(IERSConventions conventions,
boolean simpleEOP)
Get the CIRF2000 reference frame.
|
static Frame |
FramesFactory.getEcliptic(IERSConventions conventions)
Get the ecliptic frame.
|
double |
TopocentricFrame.getElevation(Vector3D extPoint,
Frame frame,
AbsoluteDate date)
Get the elevation of a point with regards to the local point.
|
static EOPHistory |
FramesFactory.getEOPHistory(IERSConventions conventions,
boolean simpleEOP)
Get Earth Orientation Parameters history.
|
Frame |
OrphanFrame.getFrame()
Get the associated
frame. |
static Frame |
FramesFactory.getFrame(Predefined factoryKey)
Get one of the predefined frames.
|
Frame |
Frame.getFrozenFrame(Frame reference,
AbsoluteDate freezingDate,
String frozenName)
Get a new version of the instance, frozen with respect to a reference frame.
|
static FactoryManagedFrame |
FramesFactory.getGTOD(boolean applyEOPCorr)
Get the GTOD reference frame.
|
static FactoryManagedFrame |
FramesFactory.getGTOD(IERSConventions conventions,
boolean simpleEOP)
Get the GTOD reference frame.
|
static Frame |
FramesFactory.getICRF()
Get the unique ICRF frame.
|
static FactoryManagedFrame |
FramesFactory.getITRF(IERSConventions conventions,
boolean simpleEOP)
Get the ITRF2008 reference frame, using IERS 2010 conventions.
|
static FactoryManagedFrame |
FramesFactory.getITRFEquinox(IERSConventions conventions,
boolean simpleEOP)
Get the equinox-based ITRF reference frame.
|
static FactoryManagedFrame |
FramesFactory.getMOD(boolean applyEOPCorr)
Get the MOD reference frame.
|
static FactoryManagedFrame |
FramesFactory.getMOD(IERSConventions conventions)
Get the MOD reference frame.
|
EOPHistory |
EOPHistory.getNonInterpolatingEOPHistory()
Get non-interpolating version of the instance.
|
GTODProvider |
GTODProvider.getNonInterpolatingProvider()
Get a version of the provider that does not cache tidal corrections.
|
EOPBasedTransformProvider |
EOPBasedTransformProvider.getNonInterpolatingProvider()
Get a version of the provider that does not cache tidal corrections.
|
static Transform |
FramesFactory.getNonInterpolatingTransform(Frame from,
Frame to,
AbsoluteDate date)
Get the transform between two frames, suppressing all interpolation.
|
static <T extends RealFieldElement<T>> |
FramesFactory.getNonInterpolatingTransform(Frame from,
Frame to,
FieldAbsoluteDate<T> date)
Get the transform between two frames, suppressing all interpolation.
|
TimeStampedPVCoordinates |
TopocentricFrame.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the topocentric frame origin in the selected frame. |
double |
TopocentricFrame.getRange(Vector3D extPoint,
Frame frame,
AbsoluteDate date)
Get the range of a point with regards to the topocentric frame center point.
|
double |
TopocentricFrame.getRangeRate(PVCoordinates extPV,
Frame frame,
AbsoluteDate date)
Get the range rate of a point with regards to the topocentric frame center point.
|
static FactoryManagedFrame |
FramesFactory.getTEME()
Get the TEME reference frame.
|
static FactoryManagedFrame |
FramesFactory.getTIRF(IERSConventions conventions)
Get the TIRF reference frame, ignoring tidal effects.
|
static FactoryManagedFrame |
FramesFactory.getTIRF(IERSConventions conventions,
boolean simpleEOP)
Get the TIRF reference frame.
|
static FactoryManagedFrame |
FramesFactory.getTOD(boolean applyEOPCorr)
Get the TOD reference frame.
|
static FactoryManagedFrame |
FramesFactory.getTOD(IERSConventions conventions,
boolean simpleEOP)
Get the TOD reference frame.
|
Transform |
ShiftingTransformProvider.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
Transform |
TransformProvider.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
Transform |
InterpolatingTransformProvider.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
Transform |
GTODProvider.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
Transform |
EclipticProvider.getTransform(AbsoluteDate date) |
<T extends RealFieldElement<T>> |
ShiftingTransformProvider.getTransform(FieldAbsoluteDate<T> date)
Get the
FieldTransform corresponding to specified date. |
<T extends RealFieldElement<T>> |
TransformProvider.getTransform(FieldAbsoluteDate<T> date)
Get the
FieldTransform corresponding to specified date. |
<T extends RealFieldElement<T>> |
InterpolatingTransformProvider.getTransform(FieldAbsoluteDate<T> date)
Get the
FieldTransform corresponding to specified date. |
<T extends RealFieldElement<T>> |
EclipticProvider.getTransform(FieldAbsoluteDate<T> date) |
Transform |
Frame.getTransformTo(Frame destination,
AbsoluteDate date)
Get the transform from the instance to another frame.
|
<T extends RealFieldElement<T>> |
Frame.getTransformTo(Frame destination,
FieldAbsoluteDate<T> date)
Get the transform from the instance to another frame.
|
static FactoryManagedFrame |
FramesFactory.getVeis1950()
Get the VEIS 1950 reference frame.
|
static Transform |
Transform.interpolate(AbsoluteDate date,
CartesianDerivativesFilter cFilter,
AngularDerivativesFilter aFilter,
Collection<Transform> sample)
Interpolate a transform from a sample set of existing transforms.
|
Transform |
Transform.interpolate(AbsoluteDate interpolationDate,
Stream<Transform> sample)
Get an interpolated instance.
|
static <T extends RealFieldElement<T>> |
FieldTransform.interpolate(FieldAbsoluteDate<T> date,
CartesianDerivativesFilter cFilter,
AngularDerivativesFilter aFilter,
Collection<FieldTransform<T>> sample)
Interpolate a transform from a sample set of existing transforms.
|
static <T extends RealFieldElement<T>> |
FieldTransform.interpolate(FieldAbsoluteDate<T> date,
CartesianDerivativesFilter cFilter,
AngularDerivativesFilter aFilter,
Stream<FieldTransform<T>> sample)
Interpolate a transform from a sample set of existing transforms.
|
static <T extends RealFieldElement<T>> |
FieldTransform.interpolate(FieldAbsoluteDate<T> interpolationDate,
Collection<FieldTransform<T>> sample)
Interpolate a transform from a sample set of existing transforms.
|
GeodeticPoint |
TopocentricFrame.pointAtDistance(double azimuth,
double elevation,
double distance)
Compute the point observed from the station at some specified distance.
|
void |
UpdatableFrame.updateTransform(Frame f1,
Frame f2,
Transform f1Tof2,
AbsoluteDate date)
Update the transform from parent frame implicitly according to two other
frames.
|
| Constructor and Description |
|---|
EclipticProvider(IERSConventions conventions)
Create a transform provider from MOD to an ecliptically aligned frame.
|
EOPEntry(int mjd,
double dt,
double lod,
double x,
double y,
double ddPsi,
double ddEps,
double dx,
double dy)
Simple constructor.
|
EOPHistory(IERSConventions conventions,
Collection<EOPEntry> data,
boolean simpleEOP)
Simple constructor.
|
GTODProvider(IERSConventions conventions,
EOPHistory eopHistory)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
DOP |
DOPComputer.compute(AbsoluteDate date,
List<Propagator> gnss)
Compute the
DOP at a given date for a set of GNSS spacecrafts. |
void |
YUMAParser.loadData()
Loads almanacs.
|
void |
SEMParser.loadData()
Loads almanacs.
|
void |
YUMAParser.loadData(InputStream input,
String name) |
void |
SEMParser.loadData(InputStream input,
String name) |
| Modifier and Type | Method and Description |
|---|---|
static FixedTroposphericDelay |
FixedTroposphericDelay.getDefaultModel()
Returns the default model, loading delay values from the file
"tropospheric-delay.txt".
|
static GeoMagneticField |
GeoMagneticFieldFactory.getField(GeoMagneticFieldFactory.FieldModel type,
double year)
Get the
GeoMagneticField for the given model type and year. |
static GeoMagneticField |
GeoMagneticFieldFactory.getIGRF(double year)
Get the IGRF model for the given year.
|
<T extends RealFieldElement<T>> |
Geoid.getIntersectionPoint(FieldLine<T> lineInFrame,
FieldVector3D<T> closeInFrame,
Frame frame,
FieldAbsoluteDate<T> date)
Get the intersection point of a line with the surface of the body.
|
GeodeticPoint |
Geoid.getIntersectionPoint(Line lineInFrame,
Vector3D closeInFrame,
Frame frame,
AbsoluteDate date)
Get the intersection point of a line with the surface of the body.
|
static SaastamoinenModel |
SaastamoinenModel.getStandardModel()
Create a new Saastamoinen model using a standard atmosphere model.
|
double |
Geoid.getUndulation(double geodeticLatitude,
double longitude,
AbsoluteDate date)
Gets the Undulation of the Geoid, N at the given position.
|
static GeoMagneticField |
GeoMagneticFieldFactory.getWMM(double year)
Get the WMM model for the given year.
|
void |
KlobucharIonoCoefficientsLoader.loadData(InputStream input,
String name)
Load Klobuchar-Style ionospheric coefficients read from some file.
|
void |
KlobucharIonoCoefficientsLoader.loadKlobucharIonosphericCoefficients()
Load the data using supported names .
|
void |
KlobucharIonoCoefficientsLoader.loadKlobucharIonosphericCoefficients(DateComponents dateComponents)
Load the data for a given day.
|
TimeStampedPVCoordinates |
Geoid.projectToGround(TimeStampedPVCoordinates pv,
Frame frame) |
Vector3D |
Geoid.projectToGround(Vector3D point,
AbsoluteDate date,
Frame frame) |
<T extends RealFieldElement<T>> |
Geoid.transform(FieldVector3D<T> point,
Frame frame,
FieldAbsoluteDate<T> date)
Transform a Cartesian point to a surface-relative point.
|
GeodeticPoint |
Geoid.transform(Vector3D point,
Frame frame,
AbsoluteDate date)
Transform a Cartesian point to a surface-relative point.
|
GeoMagneticField |
GeoMagneticField.transformModel(double year)
Time transform the model coefficients from the base year of the model
using secular variation coefficients.
|
GeoMagneticField |
GeoMagneticField.transformModel(GeoMagneticField otherModel,
double year)
Time transform the model coefficients from the base year of the model
using a linear interpolation with a second model.
|
| Constructor and Description |
|---|
FixedTroposphericDelay(String supportedName)
Creates a new
FixedTroposphericDelay instance, and loads the
delay values from the given resource via the DataProvidersManager. |
SaastamoinenModel(double t0,
double p0,
double r0,
String deltaRFileName)
Create a new Saastamoinen model for the troposphere using the given
environmental conditions.
|
| Modifier and Type | Method and Description |
|---|---|
Vector3D |
AlongTrackAiming.alongTileDirection(Vector3D point,
GeodeticPoint gp)
Find the along tile direction for tessellation at specified point.
|
Vector3D |
TileAiming.alongTileDirection(Vector3D point,
GeodeticPoint gp)
Find the along tile direction for tessellation at specified point.
|
List<List<GeodeticPoint>> |
EllipsoidTessellator.sample(SphericalPolygonsSet zone,
double width,
double length)
Sample a zone of interest into a grid sample of
geodetic points. |
List<List<Tile>> |
EllipsoidTessellator.tessellate(SphericalPolygonsSet zone,
double fullWidth,
double fullLength,
double widthOverlap,
double lengthOverlap,
boolean truncateLastWidth,
boolean truncateLastLength)
Tessellate a zone of interest into tiles.
|
| Constructor and Description |
|---|
AlongTrackAiming(OneAxisEllipsoid ellipsoid,
Orbit orbit,
boolean isAscending)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
abstract ParameterDriversList |
OrbitType.getDrivers(double dP,
Orbit orbit,
PositionAngle type)
Get parameters drivers initialized from a reference orbit.
|
TimeStampedPVCoordinates |
Orbit.getPVCoordinates(AbsoluteDate otherDate,
Frame otherFrame)
Get the
PVCoordinates of the body in the selected frame. |
TimeStampedFieldPVCoordinates<T> |
FieldOrbit.getPVCoordinates(FieldAbsoluteDate<T> otherDate,
Frame otherFrame)
Get the
FieldPVCoordinates of the body in the selected frame. |
TimeStampedPVCoordinates |
Orbit.getPVCoordinates(Frame outputFrame)
Get the
TimeStampedPVCoordinates in a specified frame. |
TimeStampedFieldPVCoordinates<T> |
FieldOrbit.getPVCoordinates(Frame outputFrame)
Get the
TimeStampedPVCoordinates in a specified frame. |
protected double[] |
OrbitType.scale(double dP,
Orbit orbit)
Compute scaling factor for parameters drivers.
|
| Modifier and Type | Method and Description |
|---|---|
void |
Propagator.addAdditionalStateProvider(AdditionalStateProvider additionalStateProvider)
Add a set of user-specified state parameters to be computed along with the orbit propagation.
|
void |
AbstractPropagator.addAdditionalStateProvider(AdditionalStateProvider additionalStateProvider)
Add a set of user-specified state parameters to be computed along with the orbit propagation.
|
void |
FieldAbstractPropagator.addAdditionalStateProvider(FieldAdditionalStateProvider<T> additionalStateProvider)
Add a set of user-specified state parameters to be computed along with the orbit propagation.
|
void |
FieldPropagator.addAdditionalStateProvider(FieldAdditionalStateProvider<T> additionalStateProvider)
Add a set of user-specified state parameters to be computed along with the orbit propagation.
|
void |
FieldSpacecraftState.ensureCompatibleAdditionalStates(FieldSpacecraftState<T> state)
Check if two instances have the same set of additional states available.
|
void |
SpacecraftState.ensureCompatibleAdditionalStates(SpacecraftState state)
Check if two instances have the same set of additional states available.
|
T[] |
FieldAdditionalStateProvider.getAdditionalState(FieldSpacecraftState<T> state)
Get the additional state.
|
double[] |
AdditionalStateProvider.getAdditionalState(SpacecraftState state)
Get the additional state.
|
double[] |
SpacecraftState.getAdditionalState(String name)
Get an additional state.
|
T[] |
FieldSpacecraftState.getAdditionalState(String name)
Get an additional state.
|
SpacecraftState |
Propagator.getInitialState()
Get the propagator initial state.
|
FieldSpacecraftState<T> |
FieldAbstractPropagator.getInitialState()
Get the propagator initial state.
|
SpacecraftState |
AbstractPropagator.getInitialState()
Get the propagator initial state.
|
FieldSpacecraftState<T> |
FieldPropagator.getInitialState()
Get the propagator initial state.
|
TimeStampedPVCoordinates |
AbstractPropagator.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
TimeStampedFieldPVCoordinates<T> |
FieldAbstractPropagator.getPVCoordinates(FieldAbsoluteDate<T> date,
Frame frame)
Get the
FieldPVCoordinates of the body in the selected frame. |
TimeStampedPVCoordinates |
SpacecraftState.getPVCoordinates(Frame outputFrame)
Get the
TimeStampedPVCoordinates in given output frame. |
TimeStampedFieldPVCoordinates<T> |
FieldSpacecraftState.getPVCoordinates(Frame outputFrame)
Get the
TimeStampedFieldPVCoordinates in given output frame. |
SpacecraftState |
SpacecraftState.interpolate(AbsoluteDate date,
Stream<SpacecraftState> sample)
Get an interpolated instance.
|
FieldSpacecraftState<T> |
FieldSpacecraftState.interpolate(FieldAbsoluteDate<T> date,
Stream<FieldSpacecraftState<T>> sample)
Get an interpolated instance.
|
SpacecraftState |
Propagator.propagate(AbsoluteDate target)
Propagate towards a target date.
|
SpacecraftState |
AbstractPropagator.propagate(AbsoluteDate target)
Propagate towards a target date.
|
SpacecraftState |
Propagator.propagate(AbsoluteDate start,
AbsoluteDate target)
Propagate from a start date towards a target date.
|
List<SpacecraftState> |
PropagatorsParallelizer.propagate(AbsoluteDate start,
AbsoluteDate target)
Propagate from a start date towards a target date.
|
FieldSpacecraftState<T> |
FieldAbstractPropagator.propagate(FieldAbsoluteDate<T> target)
Propagate towards a target date.
|
FieldSpacecraftState<T> |
FieldPropagator.propagate(FieldAbsoluteDate<T> target)
Propagate towards a target date.
|
FieldSpacecraftState<T> |
FieldPropagator.propagate(FieldAbsoluteDate<T> start,
FieldAbsoluteDate<T> target)
Propagate from a start date towards a target date.
|
void |
FieldAbstractPropagator.resetInitialState(FieldSpacecraftState<T> state)
Reset the propagator initial state.
|
void |
FieldPropagator.resetInitialState(FieldSpacecraftState<T> state)
Reset the propagator initial state.
|
void |
Propagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
void |
AbstractPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
protected FieldSpacecraftState<T> |
FieldAbstractPropagator.updateAdditionalStates(FieldSpacecraftState<T> original)
Update state by adding all additional states.
|
protected SpacecraftState |
AbstractPropagator.updateAdditionalStates(SpacecraftState original)
Update state by adding all additional states.
|
| Constructor and Description |
|---|
FieldSpacecraftState(FieldOrbit<T> orbit)
Build a spacecraft state from orbit only.
|
FieldSpacecraftState(FieldOrbit<T> orbit,
Map<String,T[]> additional)
Build a spacecraft state from orbit only.
|
FieldSpacecraftState(FieldOrbit<T> orbit,
T mass)
Create a new instance from orbit and mass.
|
FieldSpacecraftState(FieldOrbit<T> orbit,
T mass,
Map<String,T[]> additional)
Create a new instance from orbit and mass.
|
SpacecraftState(Orbit orbit)
Build a spacecraft state from orbit only.
|
SpacecraftState(Orbit orbit,
double mass)
Create a new instance from orbit and mass.
|
SpacecraftState(Orbit orbit,
double mass,
Map<String,double[]> additional)
Create a new instance from orbit and mass.
|
SpacecraftState(Orbit orbit,
Map<String,double[]> additional)
Build a spacecraft state from orbit only.
|
| Modifier and Type | Method and Description |
|---|---|
protected FieldSpacecraftState<T> |
FieldAbstractAnalyticalPropagator.acceptStep(org.orekit.propagation.analytical.FieldAbstractAnalyticalPropagator.FieldBasicStepInterpolator interpolator,
FieldAbsoluteDate<T> target,
double epsilon)
Accept a step, triggering events and step handlers.
|
protected SpacecraftState |
AbstractAnalyticalPropagator.acceptStep(OrekitStepInterpolator interpolator,
AbsoluteDate target,
double epsilon)
Accept a step, triggering events and step handlers.
|
SpacecraftState |
AdapterPropagator.DifferentialEffect.apply(SpacecraftState original)
Apply the effect to a
spacecraft state. |
protected SpacecraftState |
AbstractAnalyticalPropagator.basicPropagate(AbsoluteDate date)
Propagate an orbit without any fancy features.
|
SpacecraftState |
Ephemeris.basicPropagate(AbsoluteDate date) |
protected SpacecraftState |
AdapterPropagator.basicPropagate(AbsoluteDate date)
Propagate an orbit without any fancy features.
|
protected FieldSpacecraftState<T> |
FieldAbstractAnalyticalPropagator.basicPropagate(FieldAbsoluteDate<T> date)
Propagate an orbit without any fancy features.
|
SpacecraftState |
Ephemeris.getInitialState()
Get the propagator initial state.
|
SpacecraftState |
AdapterPropagator.getInitialState()
Get the propagator initial state.
|
SpacecraftState |
AggregateBoundedPropagator.getInitialState() |
protected abstract double |
AbstractAnalyticalPropagator.getMass(AbsoluteDate date)
Get the mass.
|
protected double |
Ephemeris.getMass(AbsoluteDate date)
Get the mass.
|
protected double |
AdapterPropagator.getMass(AbsoluteDate date)
Get the mass.
|
protected double |
AggregateBoundedPropagator.getMass(AbsoluteDate date) |
protected abstract T |
FieldAbstractAnalyticalPropagator.getMass(FieldAbsoluteDate<T> date)
Get the mass.
|
TimeStampedPVCoordinates |
Ephemeris.getPVCoordinates(AbsoluteDate date,
Frame f)
Get the
PVCoordinates of the body in the selected frame. |
TimeStampedPVCoordinates |
AggregateBoundedPropagator.getPVCoordinates(AbsoluteDate date,
Frame frame) |
SpacecraftState |
AbstractAnalyticalPropagator.propagate(AbsoluteDate start,
AbsoluteDate target)
Propagate from a start date towards a target date.
|
FieldSpacecraftState<T> |
FieldAbstractAnalyticalPropagator.propagate(FieldAbsoluteDate<T> start,
FieldAbsoluteDate<T> target)
Propagate from a start date towards a target date.
|
CartesianOrbit |
EcksteinHechlerPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
protected abstract Orbit |
AbstractAnalyticalPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
protected Orbit |
Ephemeris.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
protected Orbit |
AdapterPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
protected Orbit |
AggregateBoundedPropagator.propagateOrbit(AbsoluteDate date) |
protected Orbit |
KeplerianPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
protected FieldOrbit<T> |
FieldKeplerianPropagator.propagateOrbit(FieldAbsoluteDate<T> date)
Extrapolate an orbit up to a specific target date.
|
FieldCartesianOrbit<T> |
FieldEcksteinHechlerPropagator.propagateOrbit(FieldAbsoluteDate<T> date)
Extrapolate an orbit up to a specific target date.
|
protected abstract FieldOrbit<T> |
FieldAbstractAnalyticalPropagator.propagateOrbit(FieldAbsoluteDate<T> date)
Extrapolate an orbit up to a specific target date.
|
void |
FieldKeplerianPropagator.resetInitialState(FieldSpacecraftState<T> state)
Reset the propagator initial state.
|
void |
FieldEcksteinHechlerPropagator.resetInitialState(FieldSpacecraftState<T> state)
Reset the propagator initial state.
|
void |
EcksteinHechlerPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
void |
Ephemeris.resetInitialState(SpacecraftState state)
Try (and fail) to reset the initial state.
|
void |
AdapterPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
void |
AggregateBoundedPropagator.resetInitialState(SpacecraftState state) |
void |
KeplerianPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
protected void |
FieldKeplerianPropagator.resetIntermediateState(FieldSpacecraftState<T> state,
boolean forward)
Reset an intermediate state.
|
protected void |
FieldEcksteinHechlerPropagator.resetIntermediateState(FieldSpacecraftState<T> state,
boolean forward)
Reset an intermediate state.
|
protected abstract void |
FieldAbstractAnalyticalPropagator.resetIntermediateState(FieldSpacecraftState<T> state,
boolean forward)
Reset an intermediate state.
|
protected void |
EcksteinHechlerPropagator.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
protected abstract void |
AbstractAnalyticalPropagator.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
protected void |
Ephemeris.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
protected void |
AdapterPropagator.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
protected void |
AggregateBoundedPropagator.resetIntermediateState(SpacecraftState state,
boolean forward) |
protected void |
KeplerianPropagator.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
| Constructor and Description |
|---|
AggregateBoundedPropagator(Collection<? extends BoundedPropagator> propagators)
Create a propagator by concatenating several
BoundedPropagators. |
EcksteinHechlerPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from orbit, attitude provider and potential.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv,
double mass,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from orbit, attitude provider, mass and potential.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv,
double mass,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from orbit, attitude provider, mass and potential provider.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
AttitudeProvider attitude,
double mass,
UnnormalizedSphericalHarmonicsProvider provider,
UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics harmonics)
Private helper constructor.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from orbit, attitude provider and potential provider.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from orbit and potential.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
double mass,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from orbit, mass and potential.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
double mass,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from orbit, mass and potential provider.
|
EcksteinHechlerPropagator(Orbit initialOrbit,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from orbit and potential provider.
|
Ephemeris(List<SpacecraftState> states,
int interpolationPoints)
Constructor with tabulated states.
|
Ephemeris(List<SpacecraftState> states,
int interpolationPoints,
double extrapolationThreshold)
Constructor with tabulated states.
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
AttitudeProvider attitudeProv,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
AttitudeProvider attitudeProv,
T mass,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
AttitudeProvider attitudeProv,
T mass,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
AttitudeProvider attitude,
T mass,
UnnormalizedSphericalHarmonicsProvider provider,
UnnormalizedSphericalHarmonicsProvider.UnnormalizedSphericalHarmonics harmonics)
Private helper constructor.
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
AttitudeProvider attitudeProv,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
T mass,
double referenceRadius,
double mu,
double c20,
double c30,
double c40,
double c50,
double c60)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
T mass,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from FieldOrbit
|
FieldEcksteinHechlerPropagator(FieldOrbit<T> initialOrbit,
UnnormalizedSphericalHarmonicsProvider provider)
Build a propagator from FieldOrbit
|
FieldKeplerianPropagator(FieldOrbit<T> initialFieldOrbit)
Build a propagator from orbit only.
|
FieldKeplerianPropagator(FieldOrbit<T> initialFieldOrbit,
AttitudeProvider attitudeProv)
Build a propagator from orbit and attitude provider.
|
FieldKeplerianPropagator(FieldOrbit<T> initialFieldOrbit,
AttitudeProvider attitudeProv,
double mu)
Build a propagator from orbit, attitude provider and central attraction
coefficient μ.
|
FieldKeplerianPropagator(FieldOrbit<T> initialOrbit,
AttitudeProvider attitudeProv,
double mu,
T mass)
Build propagator from orbit, attitude provider, central attraction
coefficient μ and mass.
|
FieldKeplerianPropagator(FieldOrbit<T> initialFieldOrbit,
double mu)
Build a propagator from orbit and central attraction coefficient μ.
|
J2DifferentialEffect(Orbit orbit0,
Orbit orbit1,
boolean applyBefore,
UnnormalizedSphericalHarmonicsProvider gravityField)
Simple constructor.
|
J2DifferentialEffect(SpacecraftState original,
AdapterPropagator.DifferentialEffect directEffect,
boolean applyBefore,
double referenceRadius,
double mu,
double j2)
Simple constructor.
|
J2DifferentialEffect(SpacecraftState original,
AdapterPropagator.DifferentialEffect directEffect,
boolean applyBefore,
UnnormalizedSphericalHarmonicsProvider gravityField)
Simple constructor.
|
KeplerianPropagator(Orbit initialOrbit)
Build a propagator from orbit only.
|
KeplerianPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv)
Build a propagator from orbit and attitude provider.
|
KeplerianPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv,
double mu)
Build a propagator from orbit, attitude provider and central attraction
coefficient μ.
|
KeplerianPropagator(Orbit initialOrbit,
AttitudeProvider attitudeProv,
double mu,
double mass)
Build propagator from orbit, attitude provider, central attraction
coefficient μ and mass.
|
KeplerianPropagator(Orbit initialOrbit,
double mu)
Build a propagator from orbit and central attraction coefficient μ.
|
| Modifier and Type | Method and Description |
|---|---|
protected Orbit |
GPSPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
void |
GPSPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
protected void |
GPSPropagator.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
| Constructor and Description |
|---|
Builder(GPSOrbitalElements gpsOrbElt)
Initializes the builder.
|
| Modifier and Type | Method and Description |
|---|---|
Set<Integer> |
TLESeries.getAvailableSatelliteNumbers()
Deprecated.
Get the available satellite numbers.
|
String |
TLE.getLine1()
Get the first line.
|
String |
TLE.getLine2()
Get the second line.
|
PVCoordinates |
TLEPropagator.getPVCoordinates(AbsoluteDate date)
Get the extrapolated position and velocity from an initial TLE.
|
PVCoordinates |
TLESeries.getPVCoordinates(AbsoluteDate date)
Deprecated.
Get the extrapolated position and velocity from an initial date.
|
static boolean |
TLE.isFormatOK(String line1,
String line2)
Check the lines format validity.
|
void |
TLESeries.loadData(InputStream input,
String name)
Deprecated.
Load data from a stream.
|
void |
TLESeries.loadTLEData()
Deprecated.
Load TLE data for a specified object.
|
void |
TLESeries.loadTLEData(int satelliteNumber)
Deprecated.
Load TLE data for a specified object.
|
void |
TLESeries.loadTLEData(int launchYear,
int launchNumber,
String launchPiece)
Deprecated.
Load TLE data for a specified object.
|
protected void |
DeepSDP4.luniSolarTermsComputation()
Computes luni - solar terms from initial coordinates and epoch.
|
protected Orbit |
TLEPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
void |
TLEPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
protected void |
TLEPropagator.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
static TLEPropagator |
TLEPropagator.selectExtrapolator(TLE tle)
Selects the extrapolator to use with the selected TLE.
|
static TLEPropagator |
TLEPropagator.selectExtrapolator(TLE tle,
AttitudeProvider attitudeProvider,
double mass)
Selects the extrapolator to use with the selected TLE.
|
protected abstract void |
TLEPropagator.sxpInitialize()
Initialization proper to each propagator (SGP or SDP).
|
protected abstract void |
TLEPropagator.sxpPropagate(double t)
Propagation proper to each propagator (SGP or SDP).
|
| Constructor and Description |
|---|
DeepSDP4(TLE initialTLE,
AttitudeProvider attitudeProvider,
double mass)
Constructor for a unique initial TLE.
|
SGP4(TLE initialTLE,
AttitudeProvider attitudeProvider,
double mass)
Constructor for a unique initial TLE.
|
TLE(String line1,
String line2)
Simple constructor from unparsed two lines.
|
TLEPropagator(TLE initialTLE,
AttitudeProvider attitudeProvider,
double mass)
Protected constructor for derived classes.
|
| Modifier and Type | Method and Description |
|---|---|
void |
NumericalPropagatorBuilder.addForceModel(ForceModel model)
Add a force model to the global perturbation model.
|
protected void |
AbstractPropagatorBuilder.addSupportedParameter(ParameterDriver driver)
Add a supported parameter.
|
AbstractIntegrator |
ODEIntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
AbstractIntegrator |
AdamsBashforthIntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
AbstractIntegrator |
AdamsMoultonIntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
AbstractIntegrator |
GraggBulirschStoerIntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
AbstractIntegrator |
HighamHall54IntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
AbstractIntegrator |
DormandPrince853IntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
AbstractIntegrator |
DormandPrince54IntegratorBuilder.buildIntegrator(Orbit orbit,
OrbitType orbitType)
Build a first order integrator.
|
Propagator |
TLEPropagatorBuilder.buildPropagator(double[] normalizedParameters)
Build a propagator.
|
Propagator |
EcksteinHechlerPropagatorBuilder.buildPropagator(double[] normalizedParameters)
Build a propagator.
|
Propagator |
PropagatorBuilder.buildPropagator(double[] normalizedParameters)
Build a propagator.
|
Propagator |
KeplerianPropagatorBuilder.buildPropagator(double[] normalizedParameters)
Build a propagator.
|
NumericalPropagator |
NumericalPropagatorBuilder.buildPropagator(double[] normalizedParameters)
Build a propagator.
|
SpacecraftState |
OsculatingToMeanElementsConverter.convert()
Convert an osculating orbit into a mean orbit, in DSST sense.
|
Propagator |
AbstractPropagatorConverter.convert(List<SpacecraftState> states,
boolean positionOnly,
List<String> freeParameters)
Find the propagator that minimize the mean square error for a sample of
states. |
Propagator |
PropagatorConverter.convert(List<SpacecraftState> states,
boolean positionOnly,
List<String> freeParameters)
Find the propagator that minimize the mean square error for a sample of
states. |
Propagator |
AbstractPropagatorConverter.convert(List<SpacecraftState> states,
boolean positionOnly,
String... freeParameters)
Find the propagator that minimize the mean square error for a sample of
states. |
Propagator |
PropagatorConverter.convert(List<SpacecraftState> states,
boolean positionOnly,
String... freeParameters)
Find the propagator that minimize the mean square error for a sample of
states. |
Propagator |
AbstractPropagatorConverter.convert(Propagator source,
double timeSpan,
int nbPoints,
List<String> freeParameters)
Convert a propagator to another.
|
Propagator |
PropagatorConverter.convert(Propagator source,
double timeSpan,
int nbPoints,
List<String> freeParameters)
Convert a propagator into another one.
|
Propagator |
AbstractPropagatorConverter.convert(Propagator source,
double timeSpan,
int nbPoints,
String... freeParameters)
Convert a propagator to another.
|
Propagator |
PropagatorConverter.convert(Propagator source,
double timeSpan,
int nbPoints,
String... freeParameters)
Convert a propagator into another one.
|
protected void |
AbstractPropagatorBuilder.setParameters(double[] normalizedParameters)
Set the selected parameters.
|
| Constructor and Description |
|---|
AbstractPropagatorBuilder(Orbit templateOrbit,
PositionAngle positionAngle,
double positionScale,
boolean addDriverForCentralAttraction)
Build a new instance.
|
EcksteinHechlerPropagatorBuilder(Orbit templateOrbit,
double referenceRadius,
double mu,
TideSystem tideSystem,
double c20,
double c30,
double c40,
double c50,
double c60,
OrbitType orbitType,
PositionAngle positionAngle,
double positionScale)
Build a new instance.
|
EcksteinHechlerPropagatorBuilder(Orbit templateOrbit,
UnnormalizedSphericalHarmonicsProvider provider,
PositionAngle positionAngle,
double positionScale)
Build a new instance.
|
JacobianPropagatorConverter(NumericalPropagatorBuilder builder,
double threshold,
int maxIterations)
Simple constructor.
|
KeplerianPropagatorBuilder(Orbit templateOrbit,
PositionAngle positionAngle,
double positionScale)
Build a new instance.
|
NumericalPropagatorBuilder(Orbit referenceOrbit,
ODEIntegratorBuilder builder,
PositionAngle positionAngle,
double positionScale)
Build a new instance.
|
TLEPropagatorBuilder(TLE templateTLE,
PositionAngle positionAngle,
double positionScale)
Build a new instance.
|
| Modifier and Type | Method and Description |
|---|---|
FieldEventState.EventOccurrence<T> |
FieldEventState.doEvent(FieldSpacecraftState<T> state)
Notify the user's listener of the event.
|
EventState.EventOccurrence |
EventState.doEvent(SpacecraftState state)
Notify the user's listener of the event.
|
boolean |
FieldEventState.evaluateStep(FieldOrekitStepInterpolator<T> interpolator)
Evaluate the impact of the proposed step on the event detector.
|
boolean |
EventState.evaluateStep(OrekitStepInterpolator interpolator)
Evaluate the impact of the proposed step on the event detector.
|
boolean |
EnablingPredicate.eventIsEnabled(SpacecraftState state,
S eventDetector,
double g)
Compute an event enabling function of state.
|
FieldEventHandler.Action |
FieldEventDetector.eventOccurred(FieldSpacecraftState<T> s,
boolean increasing)
Handle the event.
|
FieldEventHandler.Action |
FieldAbstractDetector.eventOccurred(FieldSpacecraftState<T> s,
boolean increasing)
Handle the event.
|
EventHandler.Action |
EventDetector.eventOccurred(SpacecraftState s,
boolean increasing)
Handle the event.
|
EventHandler.Action |
AbstractDetector.eventOccurred(SpacecraftState s,
boolean increasing)
Handle the event.
|
T |
FieldEventDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
T |
FieldDateDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
abstract T |
FieldAbstractDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
T |
FieldElevationDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
T |
FieldAltitudeDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
T |
FieldApsideDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
T |
FieldNodeDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
T |
FieldEclipseDetector.g(FieldSpacecraftState<T> s)
Compute the value of the switching function.
|
double |
NodeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LatitudeExtremumDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
double |
NegateDetector.g(SpacecraftState s) |
double |
PositionAngleDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
double |
ElevationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AlignmentDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
GeographicZoneDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
double |
CircularFieldOfViewDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
EventDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
FieldOfViewDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
EventShifter.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ApsideDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AngularSeparationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
EventSlopeFilter.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LongitudeExtremumDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
double |
EventEnablingPredicateFilter.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LongitudeCrossingDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
abstract double |
AbstractDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LatitudeCrossingDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
double |
FootprintOverlapDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
GroundFieldOfViewDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ElevationExtremumDetector.g(SpacecraftState s)
Compute the value of the detection function.
|
double |
EclipseDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AltitudeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
DateDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
BooleanDetector.g(SpacecraftState s) |
double |
ElevationExtremumDetector.getElevation(SpacecraftState s)
Get the elevation value.
|
void |
FieldEventState.reinitializeBegin(FieldOrekitStepInterpolator<T> interpolator)
Reinitialize the beginning of the step.
|
void |
EventState.reinitializeBegin(OrekitStepInterpolator interpolator)
Reinitialize the beginning of the step.
|
default FieldSpacecraftState<T> |
FieldEventDetector.resetState(FieldSpacecraftState<T> oldState)
Reset the state prior to continue propagation.
|
FieldSpacecraftState<T> |
FieldAbstractDetector.resetState(FieldSpacecraftState<T> oldState)
Reset the state prior to continue propagation.
|
default SpacecraftState |
EventDetector.resetState(SpacecraftState oldState)
Reset the state prior to continue propagation.
|
SpacecraftState |
AbstractDetector.resetState(SpacecraftState oldState)
Reset the state prior to continue propagation.
|
boolean |
FieldEventState.tryAdvance(FieldSpacecraftState<T> state,
FieldOrekitStepInterpolator<T> interpolator)
Try to accept the current history up to the given time.
|
boolean |
EventState.tryAdvance(SpacecraftState state,
OrekitStepInterpolator interpolator)
Try to accept the current history up to the given time.
|
| Constructor and Description |
|---|
FieldOfView(Vector3D center,
Vector3D axis1,
double halfAperture1,
Vector3D axis2,
double halfAperture2,
double margin)
Build a Field Of View with dihedral shape (i.e.
|
FootprintOverlapDetector(FieldOfView fov,
OneAxisEllipsoid body,
SphericalPolygonsSet zone,
double samplingStep)
Build a new instance.
|
| Modifier and Type | Method and Description |
|---|---|
FieldEventHandler.Action |
FieldStopOnDecreasing.eventOccurred(FieldSpacecraftState<T> s,
KK detector,
boolean increasing)
Handle a detection event and choose what to do next.
|
FieldEventHandler.Action |
FieldEventHandler.eventOccurred(FieldSpacecraftState<T> s,
KK detector,
boolean increasing)
eventOccurred method mirrors the same interface method as in
EventDetector
and its subclasses, but with an additional parameter that allows the calling
method to pass in an object from the detector which would have potential
additional data to allow the implementing class to determine the correct
return state. |
FieldEventHandler.Action |
FieldStopOnEvent.eventOccurred(FieldSpacecraftState<T> s,
KK detector,
boolean increasing)
Specific implementation of the eventOccurred interface.
|
FieldEventHandler.Action |
FieldStopOnIncreasing.eventOccurred(FieldSpacecraftState<T> s,
KK detector,
boolean increasing)
Handle a detection event and choose what to do next.
|
FieldEventHandler.Action |
FieldContinueOnEvent.eventOccurred(FieldSpacecraftState<T> s,
KK detector,
boolean increasing)
Specific implementation of the eventOccurred interface.
|
EventHandler.Action |
StopOnEvent.eventOccurred(SpacecraftState s,
T detector,
boolean increasing)
Specific implementation of the eventOccurred interface.
|
EventHandler.Action |
ContinueOnEvent.eventOccurred(SpacecraftState s,
T detector,
boolean increasing)
Specific implementation of the eventOccurred interface.
|
EventHandler.Action |
StopOnDecreasing.eventOccurred(SpacecraftState s,
T detector,
boolean increasing)
Handle a detection event and choose what to do next.
|
EventHandler.Action |
StopOnIncreasing.eventOccurred(SpacecraftState s,
T detector,
boolean increasing)
Handle a detection event and choose what to do next.
|
EventHandler.Action |
EventHandler.eventOccurred(SpacecraftState s,
T detector,
boolean increasing)
eventOccurred method mirrors the same interface method as in
EventDetector
and its subclasses, but with an additional parameter that allows the calling
method to pass in an object from the detector which would have potential
additional data to allow the implementing class to determine the correct
return state. |
default FieldSpacecraftState<T> |
FieldEventHandler.resetState(KK detector,
FieldSpacecraftState<T> oldState)
Reset the state prior to continue propagation.
|
default SpacecraftState |
EventHandler.resetState(T detector,
SpacecraftState oldState)
Reset the state prior to continue propagation.
|
| Modifier and Type | Method and Description |
|---|---|
void |
AbstractIntegratedPropagator.addAdditionalEquations(AdditionalEquations additional)
Add a set of user-specified equations to be integrated along with the orbit propagation.
|
void |
FieldAbstractIntegratedPropagator.addAdditionalEquations(FieldAdditionalEquations<T> additional)
Add a set of user-specified equations to be integrated along with the orbit propagation.
|
protected void |
AbstractIntegratedPropagator.afterIntegration()
Method called just after integration.
|
protected void |
FieldAbstractIntegratedPropagator.afterIntegration()
Method called just after integration.
|
protected SpacecraftState |
IntegratedEphemeris.basicPropagate(AbsoluteDate date)
Propagate an orbit without any fancy features.
|
protected FieldSpacecraftState<T> |
FieldIntegratedEphemeris.basicPropagate(FieldAbsoluteDate<T> date)
Propagate an orbit without any fancy features.
|
protected void |
FieldAbstractIntegratedPropagator.beforeIntegration(FieldSpacecraftState<T> initialState,
FieldAbsoluteDate<T> tEnd)
Method called just before integration.
|
protected void |
AbstractIntegratedPropagator.beforeIntegration(SpacecraftState initialState,
AbsoluteDate tEnd)
Method called just before integration.
|
T[] |
FieldAbstractIntegratedPropagator.MainStateEquations.computeDerivatives(FieldSpacecraftState<T> state)
Compute differential equations for main state.
|
T[] |
FieldAdditionalEquations.computeDerivatives(FieldSpacecraftState<T> s,
T[] pDot)
Compute the derivatives related to the additional state parameters.
|
double[] |
AbstractIntegratedPropagator.MainStateEquations.computeDerivatives(SpacecraftState state)
Compute differential equations for main state.
|
double[] |
AdditionalEquations.computeDerivatives(SpacecraftState s,
double[] pDot)
Compute the derivatives related to the additional state parameters.
|
protected SpacecraftState |
AbstractIntegratedPropagator.getInitialIntegrationState()
Get the initial state for integration.
|
protected FieldSpacecraftState<T> |
FieldAbstractIntegratedPropagator.getInitialIntegrationState()
Get the initial state for integration.
|
FieldSpacecraftState<T> |
FieldIntegratedEphemeris.getInitialState()
Get the propagator initial state.
|
SpacecraftState |
IntegratedEphemeris.getInitialState()
Get the propagator initial state.
|
protected double |
IntegratedEphemeris.getMass(AbsoluteDate date)
Get the mass.
|
protected T |
FieldIntegratedEphemeris.getMass(FieldAbsoluteDate<T> date)
Get the mass.
|
TimeStampedPVCoordinates |
IntegratedEphemeris.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
TimeStampedFieldPVCoordinates<T> |
FieldIntegratedEphemeris.getPVCoordinates(FieldAbsoluteDate<T> date,
Frame frame)
Get the
FieldPVCoordinates of the body in the selected frame. |
void |
FieldAbstractIntegratedPropagator.MainStateEquations.init(FieldSpacecraftState<T> initialState,
FieldAbsoluteDate<T> target)
Initialize the equations at the start of propagation.
|
default void |
AbstractIntegratedPropagator.MainStateEquations.init(SpacecraftState initialState,
AbsoluteDate target)
Initialize the equations at the start of propagation.
|
abstract SpacecraftState |
StateMapper.mapArrayToState(AbsoluteDate date,
double[] y,
double[] yDot,
boolean meanOnly)
Map the raw double components to a spacecraft state.
|
SpacecraftState |
StateMapper.mapArrayToState(double t,
double[] y,
double[] yDot,
boolean meanOnly)
Map the raw double components to a spacecraft state.
|
abstract FieldSpacecraftState<T> |
FieldStateMapper.mapArrayToState(FieldAbsoluteDate<T> date,
T[] y,
T[] yDot,
boolean meanOnly)
Map the raw double components to a spacecraft state.
|
FieldSpacecraftState<T> |
FieldStateMapper.mapArrayToState(T t,
T[] y,
T[] yDot,
boolean meanOnly)
Map the raw double components to a spacecraft state.
|
abstract void |
FieldStateMapper.mapStateToArray(FieldSpacecraftState<T> state,
T[] y,
T[] yDot)
Map a spacecraft state to raw double components.
|
abstract void |
StateMapper.mapStateToArray(SpacecraftState state,
double[] y,
double[] yDot)
Map a spacecraft state to raw double components.
|
SpacecraftState |
AbstractIntegratedPropagator.propagate(AbsoluteDate target)
Propagate towards a target date.
|
SpacecraftState |
AbstractIntegratedPropagator.propagate(AbsoluteDate tStart,
AbsoluteDate tEnd)
Propagate from a start date towards a target date.
|
protected SpacecraftState |
AbstractIntegratedPropagator.propagate(AbsoluteDate tEnd,
boolean activateHandlers)
Propagation with or without event detection.
|
FieldSpacecraftState<T> |
FieldAbstractIntegratedPropagator.propagate(FieldAbsoluteDate<T> target)
Propagate towards a target date.
|
protected FieldSpacecraftState<T> |
FieldAbstractIntegratedPropagator.propagate(FieldAbsoluteDate<T> tEnd,
boolean activateHandlers)
Propagation with or without event detection.
|
FieldSpacecraftState<T> |
FieldAbstractIntegratedPropagator.propagate(FieldAbsoluteDate<T> tStart,
FieldAbsoluteDate<T> tEnd)
Propagate from a start date towards a target date.
|
protected Orbit |
IntegratedEphemeris.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
protected FieldOrbit<T> |
FieldIntegratedEphemeris.propagateOrbit(FieldAbsoluteDate<T> date)
Extrapolate an orbit up to a specific target date.
|
void |
FieldIntegratedEphemeris.resetInitialState(FieldSpacecraftState<T> state)
Reset the propagator initial state.
|
void |
IntegratedEphemeris.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
protected void |
FieldIntegratedEphemeris.resetIntermediateState(FieldSpacecraftState<T> state,
boolean forward)
Reset an intermediate state.
|
protected void |
IntegratedEphemeris.resetIntermediateState(SpacecraftState state,
boolean forward)
Reset an intermediate state.
|
| Constructor and Description |
|---|
FieldIntegratedEphemeris(FieldAbsoluteDate<T> startDate,
FieldAbsoluteDate<T> minDate,
FieldAbsoluteDate<T> maxDate,
FieldStateMapper<T> mapper,
boolean meanFieldOrbit,
FieldDenseOutputModel<T> model,
Map<String,T[]> unmanaged,
List<FieldAdditionalStateProvider<T>> providers,
String[] equations)
Creates a new instance of IntegratedEphemeris.
|
IntegratedEphemeris(AbsoluteDate startDate,
AbsoluteDate minDate,
AbsoluteDate maxDate,
StateMapper mapper,
boolean meanOrbit,
DenseOutputModel model,
Map<String,double[]> unmanaged,
List<AdditionalStateProvider> providers,
String[] equations)
Creates a new instance of IntegratedEphemeris.
|
| Modifier and Type | Method and Description |
|---|---|
void |
FieldTimeDerivativesEquations.addNonKeplerianAcceleration(FieldVector3D<T> gamma)
Add the contribution of an acceleration expressed in some inertial frame.
|
void |
TimeDerivativesEquations.addNonKeplerianAcceleration(Vector3D gamma)
Add the contribution of a non-Keplerian acceleration.
|
double[] |
PartialDerivativesEquations.computeDerivatives(SpacecraftState s,
double[] pDot)
Compute the derivatives related to the additional state parameters.
|
JacobiansMapper |
PartialDerivativesEquations.getMapper()
Get a mapper between two-dimensional Jacobians and one-dimensional additional state.
|
void |
JacobiansMapper.getParametersJacobian(SpacecraftState state,
double[][] dYdP)
Get theJacobian with respect to parameters from a one-dimensional additional state array.
|
TimeStampedPVCoordinates |
NumericalPropagator.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
TimeStampedFieldPVCoordinates<T> |
FieldNumericalPropagator.getPVCoordinates(FieldAbsoluteDate<T> date,
Frame frame)
Get the
FieldPVCoordinates of the body in the selected frame. |
ParameterDriversList |
PartialDerivativesEquations.getSelectedParameters()
Get the selected parameters, in Jacobian matrix column order.
|
void |
JacobiansMapper.getStateJacobian(SpacecraftState state,
double[][] dYdY0)
Get the Jacobian with respect to state from a one-dimensional additional state array.
|
void |
FieldNumericalPropagator.resetInitialState(FieldSpacecraftState<T> state)
Reset the propagator initial state.
|
void |
NumericalPropagator.resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
SpacecraftState |
PartialDerivativesEquations.setInitialJacobians(SpacecraftState s0)
Set the initial value of the Jacobian with respect to state and parameter.
|
SpacecraftState |
PartialDerivativesEquations.setInitialJacobians(SpacecraftState s1,
double[][] dY1dY0,
double[][] dY1dP)
Set the initial value of the Jacobian with respect to state and parameter.
|
SpacecraftState |
PartialDerivativesEquations.setInitialJacobians(SpacecraftState s0,
int stateDimension)
Deprecated.
as of 9.0, replaced by
PartialDerivativesEquations.setInitialJacobians(SpacecraftState) |
void |
FieldNumericalPropagator.setInitialState(FieldSpacecraftState<T> initialState)
Set the initial state.
|
void |
NumericalPropagator.setInitialState(SpacecraftState initialState)
Set the initial state.
|
static double[][] |
NumericalPropagator.tolerances(double dP,
Orbit orbit,
OrbitType type)
Estimate tolerance vectors for integrators.
|
static <T extends RealFieldElement<T>> |
FieldNumericalPropagator.tolerances(T dP,
FieldOrbit<T> orbit,
OrbitType type)
Estimate tolerance vectors for integrators.
|
| Constructor and Description |
|---|
PartialDerivativesEquations(String name,
NumericalPropagator propagator)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
FieldSpacecraftState<T> |
FieldOrekitStepInterpolator.getCurrentState()
Get the state at previous grid point date.
|
SpacecraftState |
OrekitStepInterpolator.getCurrentState()
Get the state at current grid point date.
|
SpacecraftState |
OrekitStepInterpolator.getInterpolatedState(AbsoluteDate date)
Get the state at interpolated date.
|
FieldSpacecraftState<T> |
FieldOrekitStepInterpolator.getInterpolatedState(FieldAbsoluteDate<T> date)
Get the state at interpolated date.
|
FieldSpacecraftState<T> |
FieldOrekitStepInterpolator.getPreviousState()
Get the state at previous grid point date.
|
SpacecraftState |
OrekitStepInterpolator.getPreviousState()
Get the state at previous grid point date.
|
void |
FieldOrekitStepHandler.handleStep(FieldOrekitStepInterpolator<T> interpolator,
boolean isLast)
Handle the current step.
|
void |
FieldOrekitStepHandlerMultiplexer.handleStep(FieldOrekitStepInterpolator<T> interpolator,
boolean isLast)
Handle the current step.
|
void |
FieldOrekitStepNormalizer.handleStep(FieldOrekitStepInterpolator<T> interpolator,
boolean isLast)
Handle the last accepted step.
|
void |
FieldOrekitFixedStepHandler.handleStep(FieldSpacecraftState<T> currentState,
boolean isLast)
Handle the current step.
|
void |
MultiSatStepHandler.handleStep(List<OrekitStepInterpolator> interpolators,
boolean isLast)
Handle the current step.
|
void |
OrekitStepHandler.handleStep(OrekitStepInterpolator interpolator,
boolean isLast)
Handle the current step.
|
void |
OrekitStepNormalizer.handleStep(OrekitStepInterpolator interpolator,
boolean isLast)
Handle the last accepted step.
|
void |
OrekitStepHandlerMultiplexer.handleStep(OrekitStepInterpolator interpolator,
boolean isLast)
Handle the current step.
|
void |
OrekitFixedStepHandler.handleStep(SpacecraftState currentState,
boolean isLast)
Handle the current step.
|
void |
FieldOrekitStepHandler.init(FieldSpacecraftState<T> s0,
FieldAbsoluteDate<T> t)
Initialize step handler at the start of a propagation.
|
void |
FieldOrekitStepHandlerMultiplexer.init(FieldSpacecraftState<T> s0,
FieldAbsoluteDate<T> t)
Initialize step handler at the start of a propagation.
|
void |
FieldOrekitStepNormalizer.init(FieldSpacecraftState<T> s0,
FieldAbsoluteDate<T> t)
Initialize step handler at the start of a propagation.
|
default void |
FieldOrekitFixedStepHandler.init(FieldSpacecraftState<T> s0,
FieldAbsoluteDate<T> t,
T step)
Initialize step handler at the start of a propagation.
|
default void |
MultiSatStepHandler.init(List<SpacecraftState> states0,
AbsoluteDate t)
Initialize step handler at the start of a propagation.
|
default void |
OrekitStepHandler.init(SpacecraftState s0,
AbsoluteDate t)
Initialize step handler at the start of a propagation.
|
default void |
OrekitFixedStepHandler.init(SpacecraftState s0,
AbsoluteDate t)
Deprecated.
as of 9.0, replaced by
OrekitFixedStepHandler.init(SpacecraftState, AbsoluteDate, double) |
void |
OrekitStepNormalizer.init(SpacecraftState s0,
AbsoluteDate t)
Initialize step handler at the start of a propagation.
|
void |
OrekitStepHandlerMultiplexer.init(SpacecraftState s0,
AbsoluteDate t)
Initialize step handler at the start of a propagation.
|
default void |
OrekitFixedStepHandler.init(SpacecraftState s0,
AbsoluteDate t,
double step)
Initialize step handler at the start of a propagation.
|
OrekitStepInterpolator |
OrekitStepInterpolator.restrictStep(SpacecraftState newPreviousState,
SpacecraftState newCurrentState)
Create a new restricted version of the instance.
|
| Modifier and Type | Method and Description |
|---|---|
protected void |
DSSTPropagator.afterIntegration()
Method called just after integration.
|
protected void |
DSSTPropagator.beforeIntegration(SpacecraftState initialState,
AbsoluteDate tEnd)
Method called just before integration.
|
static SpacecraftState |
DSSTPropagator.computeMeanState(SpacecraftState osculating,
AttitudeProvider attitudeProvider,
Collection<DSSTForceModel> forceModels)
Conversion from osculating to mean orbit.
|
static SpacecraftState |
DSSTPropagator.computeOsculatingState(SpacecraftState mean,
AttitudeProvider attitudeProvider,
Collection<DSSTForceModel> forces)
Conversion from mean to osculating orbit.
|
protected SpacecraftState |
DSSTPropagator.getInitialIntegrationState()
Get the initial state for integration.
|
void |
DSSTPropagator.resetInitialState(SpacecraftState state)
Reset the initial state.
|
void |
DSSTPropagator.setInitialState(SpacecraftState initialState)
Set the initial state with osculating orbital elements.
|
void |
DSSTPropagator.setInitialState(SpacecraftState initialState,
boolean isOsculating)
Set the initial state.
|
static double[][] |
DSSTPropagator.tolerances(double dP,
Orbit orbit)
Estimate tolerance vectors for an AdaptativeStepsizeIntegrator.
|
| Modifier and Type | Method and Description |
|---|---|
Map<String,double[]> |
ShortPeriodTerms.getCoefficients(AbsoluteDate date,
Set<String> selected)
Computes the coefficients involved in the contributions.
|
protected double[] |
DSSTSolarRadiationPressure.getLLimits(SpacecraftState state)
Compute the limits in L, the true longitude, for integration.
|
protected abstract double[] |
AbstractGaussianContribution.getLLimits(SpacecraftState state)
Compute the limits in L, the true longitude, for integration.
|
protected double[] |
DSSTAtmosphericDrag.getLLimits(SpacecraftState state)
Compute the limits in L, the true longitude, for integration.
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double[] |
DSSTZonal.getMeanElementRate(SpacecraftState spacecraftState)
Computes the mean equinoctial elements rates dai / dt.
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double[] |
DSSTForceModel.getMeanElementRate(SpacecraftState state)
Computes the mean equinoctial elements rates dai / dt.
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double[] |
DSSTTesseral.getMeanElementRate(SpacecraftState spacecraftState)
Computes the mean equinoctial elements rates dai / dt.
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double[] |
AbstractGaussianContribution.getMeanElementRate(SpacecraftState state)
Computes the mean equinoctial elements rates dai / dt.
|
List<ShortPeriodTerms> |
DSSTThirdBody.initialize(AuxiliaryElements aux,
boolean meanOnly)
Computes the highest power of the eccentricity and the highest power
of a/R3 to appear in the truncated analytical power series expansion.
|
List<ShortPeriodTerms> |
DSSTZonal.initialize(AuxiliaryElements aux,
boolean meanOnly)
Performs initialization prior to propagation for the current force model.
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List<ShortPeriodTerms> |
DSSTForceModel.initialize(AuxiliaryElements aux,
boolean meanOnly)
Performs initialization prior to propagation for the current force model.
|
List<ShortPeriodTerms> |
DSSTTesseral.initialize(AuxiliaryElements aux,
boolean meanOnly)
Performs initialization prior to propagation for the current force model.
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void |
DSSTThirdBody.initializeStep(AuxiliaryElements aux)
Performs initialization at each integration step for the current force model.
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void |
DSSTZonal.initializeStep(AuxiliaryElements aux)
Performs initialization at each integration step for the current force model.
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void |
DSSTForceModel.initializeStep(AuxiliaryElements aux)
Performs initialization at each integration step for the current force model.
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void |
DSSTTesseral.initializeStep(AuxiliaryElements aux)
Performs initialization at each integration step for the current force model.
|
void |
AbstractGaussianContribution.initializeStep(AuxiliaryElements aux)
Performs initialization at each integration step for the current force model.
|
void |
DSSTThirdBody.updateShortPeriodTerms(SpacecraftState... meanStates)
Update the short period terms.
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void |
DSSTZonal.updateShortPeriodTerms(SpacecraftState... meanStates)
Update the short period terms.
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void |
DSSTForceModel.updateShortPeriodTerms(SpacecraftState... meanStates)
Update the short period terms.
|
void |
DSSTTesseral.updateShortPeriodTerms(SpacecraftState... meanStates)
Update the short period terms.
|
void |
AbstractGaussianContribution.updateShortPeriodTerms(SpacecraftState... meanStates)
Update the short period terms.
|
double[] |
ShortPeriodTerms.value(Orbit meanOrbit)
Evaluate the contributions of the short period terms.
|
| Constructor and Description |
|---|
DSSTTesseral(Frame centralBodyFrame,
double centralBodyRotationRate,
UnnormalizedSphericalHarmonicsProvider provider,
int maxDegreeTesseralSP,
int maxOrderTesseralSP,
int maxEccPowTesseralSP,
int maxFrequencyShortPeriodics,
int maxDegreeMdailyTesseralSP,
int maxOrderMdailyTesseralSP,
int maxEccPowMdailyTesseralSP)
Simple constructor.
|
DSSTZonal(UnnormalizedSphericalHarmonicsProvider provider,
int maxDegreeShortPeriodics,
int maxEccPowShortPeriodics,
int maxFrequencyShortPeriodics)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
static double |
CoefficientsFactory.getVmns(int m,
int n,
int s)
Get the Vn,sm coefficient from Vn,s.
|
| Modifier and Type | Method and Description |
|---|---|
DateTimeComponents |
AbsoluteDate.getComponents(int minutesFromUTC)
Split the instance into date/time components for a local time.
|
DateTimeComponents |
FieldAbsoluteDate.getComponents(int minutesFromUTC)
Split the instance into date/time components for a local time.
|
DateTimeComponents |
AbsoluteDate.getComponents(TimeZone timeZone)
Split the instance into date/time components for a time zone.
|
DateTimeComponents |
FieldAbsoluteDate.getComponents(TimeZone timeZone)
Split the instance into date/time components for a time zone.
|
static GLONASSScale |
TimeScalesFactory.getGLONASS()
Get the GLObal NAvigation Satellite System time scale.
|
static GMSTScale |
TimeScalesFactory.getGMST(IERSConventions conventions,
boolean simpleEOP)
Get the Greenwich Mean Sidereal Time scale.
|
static UT1Scale |
TimeScalesFactory.getUT1(EOPHistory history)
Get the Universal Time 1 scale.
|
static UT1Scale |
TimeScalesFactory.getUT1(IERSConventions conventions,
boolean simpleEOP)
Get the Universal Time 1 scale.
|
static UTCScale |
TimeScalesFactory.getUTC()
Get the Universal Time Coordinate scale.
|
default T |
TimeInterpolable.interpolate(AbsoluteDate date,
Collection<T> sample)
Get an interpolated instance.
|
T |
TimeInterpolable.interpolate(AbsoluteDate date,
Stream<T> sample)
Get an interpolated instance.
|
default T |
FieldTimeInterpolable.interpolate(FieldAbsoluteDate<KK> date,
Collection<T> sample)
Get an interpolated instance.
|
T |
FieldTimeInterpolable.interpolate(FieldAbsoluteDate<KK> date,
Stream<T> sample)
Get an interpolated instance.
|
List<OffsetModel> |
UTCTAIHistoryFilesLoader.loadOffsets()
Load UTC-TAI offsets entries.
|
List<OffsetModel> |
UTCTAIBulletinAFilesLoader.loadOffsets()
Load UTC-TAI offsets entries.
|
List<OffsetModel> |
TAIUTCDatFilesLoader.loadOffsets()
Load UTC-TAI offsets entries.
|
List<OffsetModel> |
UTCTAIOffsetsLoader.loadOffsets()
Load UTC-TAI offsets entries.
|
static AbsoluteDate |
AbsoluteDate.parseCCSDSCalendarSegmentedTimeCode(byte preambleField,
byte[] timeField)
Build an instance from a CCSDS Calendar Segmented Time Code (CCS).
|
FieldAbsoluteDate<T> |
FieldAbsoluteDate.parseCCSDSCalendarSegmentedTimeCode(byte preambleField,
byte[] timeField)
Build an instance from a CCSDS Calendar Segmented Time Code (CCS).
|
static AbsoluteDate |
AbsoluteDate.parseCCSDSDaySegmentedTimeCode(byte preambleField,
byte[] timeField,
DateComponents agencyDefinedEpoch)
Build an instance from a CCSDS Day Segmented Time Code (CDS).
|
static <T extends RealFieldElement<T>> |
FieldAbsoluteDate.parseCCSDSDaySegmentedTimeCode(Field<T> field,
byte preambleField,
byte[] timeField,
DateComponents agencyDefinedEpoch)
Build an instance from a CCSDS Day Segmented Time Code (CDS).
|
static AbsoluteDate |
AbsoluteDate.parseCCSDSUnsegmentedTimeCode(byte preambleField1,
byte preambleField2,
byte[] timeField,
AbsoluteDate agencyDefinedEpoch)
Build an instance from a CCSDS Unsegmented Time Code (CUC).
|
static <T extends RealFieldElement<T>> |
FieldAbsoluteDate.parseCCSDSUnsegmentedTimeCode(Field<T> field,
byte preambleField1,
byte preambleField2,
byte[] timeField,
FieldAbsoluteDate<T> agencyDefinedEpoch)
Build an instance from a CCSDS Unsegmented Time Code (CUC).
|
String |
AbsoluteDate.toString(int minutesFromUTC)
Get a String representation of the instant location for a local time.
|
String |
FieldAbsoluteDate.toString(int minutesFromUTC)
Get a String representation of the instant location for a local time.
|
String |
AbsoluteDate.toString(TimeZone timeZone)
Get a String representation of the instant location for a time zone.
|
String |
FieldAbsoluteDate.toString(TimeZone timeZone)
Get a String representation of the instant location for a time zone.
|
| Modifier and Type | Method and Description |
|---|---|
void |
ParameterDriversList.add(ParameterDriver driver)
Add a driver.
|
void |
ParameterDriver.addObserver(ParameterObserver observer)
Add an observer for this driver.
|
abstract TimeVectorFunction |
IERSConventions.getEOPTidalCorrection()
Get the function computing tidal corrections for Earth Orientation Parameters.
|
abstract TimeScalarFunction |
IERSConventions.getGASTFunction(TimeScale ut1,
EOPHistory eopHistory)
Get the function computing Greenwich apparent sidereal time, in radians.
|
abstract TimeScalarFunction |
IERSConventions.getGMSTFunction(TimeScale ut1)
Get the function computing Greenwich mean sidereal time, in radians.
|
abstract TimeScalarFunction |
IERSConventions.getGMSTRateFunction(TimeScale ut1)
Get the function computing Greenwich mean sidereal time rate, in radians per second.
|
abstract LoveNumbers |
IERSConventions.getLoveNumbers()
Get the Love numbers.
|
abstract TimeScalarFunction |
IERSConventions.getMeanObliquityFunction()
Get the function computing mean obliquity of the ecliptic.
|
abstract FundamentalNutationArguments |
IERSConventions.getNutationArguments(TimeScale timeScale)
Get the fundamental nutation arguments.
|
IERSConventions.NutationCorrectionConverter |
IERSConventions.getNutationCorrectionConverter()
Create a function converting nutation corrections between
δX/δY and δΔψ/δΔε.
|
abstract TimeVectorFunction |
IERSConventions.getNutationFunction()
Get the function computing the nutation angles.
|
abstract TimeVectorFunction |
IERSConventions.getOceanPoleTide(EOPHistory eopHistory)
Get the function computing ocean pole tide (ΔC₂₁, ΔS₂₁).
|
abstract double |
IERSConventions.getPermanentTide()
Get the permanent tide to be removed from ΔC₂₀ when zero-tide potentials are used.
|
abstract TimeVectorFunction |
IERSConventions.getPrecessionFunction()
Get the function computing the precession angles.
|
TimeStampedPVCoordinates |
PVCoordinatesProvider.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
TimeStampedFieldPVCoordinates<T> |
FieldPVCoordinatesProvider.getPVCoordinates(FieldAbsoluteDate<T> date,
Frame frame)
Get the
FieldPVCoordinates of the body in the selected frame. |
abstract TimeVectorFunction |
IERSConventions.getSolidPoleTide(EOPHistory eopHistory)
Get the function computing solid pole tide (ΔC₂₁, ΔS₂₁).
|
abstract TimeVectorFunction |
IERSConventions.getTideFrequencyDependenceFunction(TimeScale ut1)
Get the function computing frequency dependent terms (ΔC₂₀, ΔC₂₁, ΔS₂₁, ΔC₂₂, ΔS₂₂).
|
abstract TimeVectorFunction |
IERSConventions.getXYSpXY2Function()
Get the function computing the Celestial Intermediate Pole and Celestial Intermediate Origin components.
|
static TimeStampedAngularCoordinates |
TimeStampedAngularCoordinates.interpolate(AbsoluteDate date,
AngularDerivativesFilter filter,
Collection<TimeStampedAngularCoordinates> sample)
Interpolate angular coordinates.
|
static <T extends RealFieldElement<T>> |
TimeStampedFieldAngularCoordinates.interpolate(AbsoluteDate date,
AngularDerivativesFilter filter,
Collection<TimeStampedFieldAngularCoordinates<T>> sample)
Interpolate angular coordinates.
|
static <T extends RealFieldElement<T>> |
TimeStampedFieldAngularCoordinates.interpolate(FieldAbsoluteDate<T> date,
AngularDerivativesFilter filter,
Collection<TimeStampedFieldAngularCoordinates<T>> sample)
Interpolate angular coordinates.
|
void |
InterpolationTableLoader.loadData(InputStream input,
String name)
Loads an bi-variate interpolation table from the given
InputStream. |
protected LoveNumbers |
IERSConventions.loadLoveNumbers(String nameLove)
Load the Love numbers.
|
void |
ParameterDriver.setNormalizedValue(double normalized)
Set normalized value.
|
void |
ParameterDriver.setValue(double newValue)
Set parameter value.
|
FieldRotation<DerivativeStructure> |
AngularCoordinates.toDerivativeStructureRotation(int order)
Transform the instance to a
FieldRotation<DerivativeStructure>. |
FieldVector3D<DerivativeStructure> |
PVCoordinates.toDerivativeStructureVector(int order)
Transform the instance to a
FieldVector3D<DerivativeStructure>. |
double[] |
IERSConventions.NutationCorrectionConverter.toEquinox(AbsoluteDate date,
double dX,
double dY)
Convert nutation corrections.
|
double[] |
IERSConventions.NutationCorrectionConverter.toNonRotating(AbsoluteDate date,
double ddPsi,
double ddEpsilon)
Convert nutation corrections.
|
double |
ParameterFunction.value(ParameterDriver parameterDriver)
Evaluate the function.
|
double[] |
StateFunction.value(SpacecraftState state)
Evaluate the function.
|
double[][] |
StateJacobian.value(SpacecraftState state)
Evaluate the Jacobian of the function.
|
void |
ParameterObserver.valueChanged(double previousValue,
ParameterDriver driver)
Notify that a parameter value has been changed.
|
| Constructor and Description |
|---|
AngularCoordinates(PVCoordinates u,
PVCoordinates v)
Build one of the rotations that transform one pv coordinates into another one.
|
AngularCoordinates(PVCoordinates u1,
PVCoordinates u2,
PVCoordinates v1,
PVCoordinates v2,
double tolerance)
Build the rotation that transforms a pair of pv coordinates into another one.
|
FieldAngularCoordinates(FieldPVCoordinates<T> u1,
FieldPVCoordinates<T> u2,
FieldPVCoordinates<T> v1,
FieldPVCoordinates<T> v2,
double tolerance)
Build the rotation that transforms a pair of pv coordinates into another one.
|
ParameterDriver(String name,
double referenceValue,
double scale,
double minValue,
double maxValue)
Simple constructor.
|
TimeStampedAngularCoordinates(AbsoluteDate date,
PVCoordinates u,
PVCoordinates v)
Build one of the rotations that transform one pv coordinates into another one.
|
TimeStampedAngularCoordinates(AbsoluteDate date,
PVCoordinates u1,
PVCoordinates u2,
PVCoordinates v1,
PVCoordinates v2,
double tolerance)
Build the rotation that transforms a pair of pv coordinates into another pair.
|
TimeStampedFieldAngularCoordinates(AbsoluteDate date,
FieldPVCoordinates<T> u1,
FieldPVCoordinates<T> u2,
FieldPVCoordinates<T> v1,
FieldPVCoordinates<T> v2,
double tolerance)
Build the rotation that transforms a pair of pv coordinates into another pair.
|
TimeStampedFieldAngularCoordinates(FieldAbsoluteDate<T> date,
FieldPVCoordinates<T> u1,
FieldPVCoordinates<T> u2,
FieldPVCoordinates<T> v1,
FieldPVCoordinates<T> v2,
double tolerance)
Build the rotation that transforms a pair of pv coordinates into another pair.
|
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