| 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.estimation.iod | |
| org.orekit.estimation.measurements |
The measurements package defines everything that is related to orbit
determination measurements.
|
| 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.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 |
|---|---|
AbsoluteDate |
Attitude.getDate()
Get the date of attitude parameters.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
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.
|
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.
|
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.
|
| Constructor and Description |
|---|
Attitude(AbsoluteDate date,
Frame referenceFrame,
AngularCoordinates orientation)
Creates a new instance.
|
Attitude(AbsoluteDate date,
Frame referenceFrame,
Rotation attitude,
Vector3D spin,
Vector3D acceleration)
Creates a new instance.
|
SpinStabilized(AttitudeProvider nonRotatingLaw,
AbsoluteDate start,
Vector3D axis,
double rate)
Creates a new instance.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
Vector3D |
IAUPole.getPole(AbsoluteDate date)
Get the body North pole direction in ICRF frame.
|
double |
IAUPole.getPrimeMeridianAngle(AbsoluteDate date)
Get the prime meridian angle.
|
PVCoordinates |
JPLEphemeridesLoader.RawPVProvider.getRawPV(AbsoluteDate date)
Get the position-velocity at date.
|
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.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
DelaunayArguments.getDate()
Get the date.
|
| Modifier and Type | Method and Description |
|---|---|
BodiesElements |
FundamentalNutationArguments.evaluateAll(AbsoluteDate date)
Evaluate all fundamental arguments for the current date (Delaunay plus planetary).
|
| Constructor and Description |
|---|
BodiesElements(AbsoluteDate date,
double tc,
double gamma,
double gammaDot,
double l,
double lDot,
double lPrime,
double lPrimeDot,
double f,
double fDot,
double d,
double dDot,
double omega,
double omegaDot,
double lMe,
double lMeDot,
double lVe,
double lVeDot,
double lE,
double lEDot,
double lMa,
double lMaDot,
double lJu,
double lJuDot,
double lSa,
double lSaDot,
double lUr,
double lUrDot,
double lNe,
double lNeDot,
double pa,
double paDot)
Simple constructor.
|
DelaunayArguments(AbsoluteDate date,
double tc,
double gamma,
double gammaDot,
double l,
double lDot,
double lPrime,
double lPrimeDot,
double f,
double fDot,
double d,
double dDot,
double omega,
double omegaDot)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
KeplerianOrbit |
IodLambert.estimate(Frame frame,
boolean posigrade,
int nRev,
Vector3D p1,
AbsoluteDate t1,
Vector3D p2,
AbsoluteDate t2)
Estimate a Keplerian orbit given two position vectors and a duration.
|
KeplerianOrbit |
IodGibbs.estimate(Frame frame,
Vector3D r1,
AbsoluteDate date1,
Vector3D r2,
AbsoluteDate date2,
Vector3D r3,
AbsoluteDate date3)
Give an initial orbit estimation, assuming Keplerian motion.
|
KeplerianOrbit |
IodGooding.estimate(Vector3D O1,
Vector3D O2,
Vector3D O3,
Vector3D lineOfSight1,
AbsoluteDate dateObs1,
Vector3D lineOfSight2,
AbsoluteDate dateObs2,
Vector3D lineOfSight3,
AbsoluteDate dateObs3,
double rho1init,
double rho3init)
Orbit got from Observed Three Lines of Sight (angles only).
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
EstimatedMeasurement.getDate()
Get the date.
|
AbsoluteDate |
AbstractMeasurement.getDate()
Get the date.
|
| Modifier and Type | Method and Description |
|---|---|
Transform |
GroundStation.getOffsetToInertial(Frame inertial,
AbsoluteDate date)
Get the transform between offset frame and inertial frame.
|
static double |
AbstractMeasurement.signalTimeOfFlight(TimeStampedPVCoordinates adjustableEmitterPV,
Vector3D receiverPosition,
AbsoluteDate signalArrivalDate)
Compute propagation delay on a link leg (typically downlink or uplink).
|
| Constructor and Description |
|---|
AbstractMeasurement(AbsoluteDate date,
double[] observed,
double[] sigma,
double[] baseWeight,
List<Integer> propagatorsIndices,
ParameterDriver... supportedParameters)
Simple constructor, for multi-dimensional measurements.
|
AbstractMeasurement(AbsoluteDate date,
double observed,
double sigma,
double baseWeight,
List<Integer> propagatorsIndices,
ParameterDriver... supportedParameters)
Simple constructor for mono-dimensional measurements.
|
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.
|
InterSatellitesRange(int satellite1Index,
int satellite2Index,
boolean twoWay,
AbsoluteDate date,
double range,
double sigma,
double baseWeight)
Simple constructor.
|
PV(AbsoluteDate date,
Vector3D position,
Vector3D velocity,
double sigmaPosition,
double sigmaVelocity,
double baseWeight)
Simple constructor.
|
PV(AbsoluteDate date,
Vector3D position,
Vector3D velocity,
double sigmaPosition,
double sigmaVelocity,
double baseWeight,
int propagatorIndex)
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 |
|---|---|
AbsoluteDate |
ODMFile.getCreationDate()
Get the file creation date and time in UTC.
|
AbsoluteDate |
TDMFile.getCreationDate()
Get the file creation date and time in UTC.
|
AbsoluteDate |
OGMFile.getEpoch()
Get epoch of state vector, Keplerian elements and covariance matrix data.
|
AbsoluteDate |
TDMFile.Observation.getEpoch()
Getter for the epoch.
|
AbsoluteDate |
OEMFile.CovarianceMatrix.getEpoch()
Get the epoch relative to the covariance matrix.
|
AbsoluteDate |
OPMFile.Maneuver.getEpochIgnition()
Get epoch ignition.
|
AbsoluteDate |
ODMMetaData.getFrameEpoch()
Get epoch of reference frame, if not intrinsic to the definition of the
reference frame.
|
AbsoluteDate |
TDMParser.getMissionReferenceDate()
Get initial date.
|
AbsoluteDate |
ODMFile.getMissionReferenceDate()
Get reference date for Mission Elapsed Time and Mission Relative Time time systems.
|
AbsoluteDate |
ODMParser.getMissionReferenceDate()
Get initial date.
|
AbsoluteDate |
OEMFile.EphemeridesBlock.getStart() |
AbsoluteDate |
OEMFile.OemSatelliteEphemeris.getStart() |
AbsoluteDate |
TDMFile.TDMMetaData.getStartTime()
Getter for the startTime.
|
AbsoluteDate |
OEMFile.EphemeridesBlock.getStartTime()
Get start of total time span covered by ephemerides data and
covariance data.
|
AbsoluteDate |
OEMFile.EphemeridesBlock.getStop() |
AbsoluteDate |
OEMFile.OemSatelliteEphemeris.getStop() |
AbsoluteDate |
TDMFile.TDMMetaData.getStopTime()
Getter for the stopTime.
|
AbsoluteDate |
OEMFile.EphemeridesBlock.getStopTime()
Get end of total time span covered by ephemerides data and covariance
data.
|
AbsoluteDate |
OEMFile.EphemeridesBlock.getUseableStartTime()
Get start of useable time span covered by ephemerides data, it may be
necessary to allow for proper interpolation.
|
AbsoluteDate |
OEMFile.EphemeridesBlock.getUseableStopTime()
Get end of useable time span covered by ephemerides data, it may be
necessary to allow for proper interpolation.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
void |
TDMFile.ObservationsBlock.addObservation(String keyword,
AbsoluteDate epoch,
double measurement)
Adds an observation data line.
|
void |
StreamingOemWriter.Segment.init(SpacecraftState s0,
AbsoluteDate t,
double step)
Initialize step handler at the start of a propagation.
|
AbsoluteDate |
CcsdsTimeScale.parseDate(String date,
IERSConventions conventions,
AbsoluteDate missionReferenceDate)
Parse a date in this time scale.
|
void |
TDMFile.Observation.setEpoch(AbsoluteDate epoch)
Setter for the epoch.
|
void |
TDMFile.TDMMetaData.setStartTime(AbsoluteDate startTime)
Setter for the startTime.
|
void |
TDMFile.TDMMetaData.setStopTime(AbsoluteDate stopTime)
Setter for the stopTime.
|
OMMParser |
OMMParser.withMissionReferenceDate(AbsoluteDate newMissionReferenceDate)
Set initial date.
|
OPMParser |
OPMParser.withMissionReferenceDate(AbsoluteDate newMissionReferenceDate)
Set initial date.
|
TDMParser |
TDMParser.withMissionReferenceDate(AbsoluteDate newMissionReferenceDate)
Set initial date.
|
abstract ODMParser |
ODMParser.withMissionReferenceDate(AbsoluteDate newMissionReferenceDate)
Set initial date.
|
OEMParser |
OEMParser.withMissionReferenceDate(AbsoluteDate newMissionReferenceDate)
Set initial date.
|
| Constructor and Description |
|---|
ODMParser(AbsoluteDate missionReferenceDate,
double mu,
IERSConventions conventions,
boolean simpleEOP,
int launchYear,
int launchNumber,
String launchPiece)
Complete constructor.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
OrekitEphemerisFile.OrekitSatelliteEphemeris.getStart() |
AbsoluteDate |
OrekitEphemerisFile.OrekitEphemerisSegment.getStart() |
AbsoluteDate |
EphemerisFile.SatelliteEphemeris.getStart()
Get the start date of the ephemeris.
|
AbsoluteDate |
EphemerisFile.EphemerisSegment.getStart()
Get the start date of this ephemeris segment.
|
AbsoluteDate |
OrekitEphemerisFile.OrekitSatelliteEphemeris.getStop() |
AbsoluteDate |
OrekitEphemerisFile.OrekitEphemerisSegment.getStop() |
AbsoluteDate |
EphemerisFile.SatelliteEphemeris.getStop()
Get the end date of the ephemeris.
|
AbsoluteDate |
EphemerisFile.EphemerisSegment.getStop()
Get the end date of this ephemeris segment.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
SP3File.getEpoch()
Returns the start epoch of the orbit file.
|
AbsoluteDate |
SP3File.SP3Ephemeris.getStart() |
AbsoluteDate |
SP3File.SP3Ephemeris.getStop() |
| Modifier and Type | Method and Description |
|---|---|
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.
|
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.
|
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.
|
| Constructor and Description |
|---|
BoxAndSolarArraySpacecraft(BoxAndSolarArraySpacecraft.Facet[] facets,
PVCoordinatesProvider sun,
double solarArrayArea,
Vector3D solarArrayAxis,
AbsoluteDate referenceDate,
Vector3D referenceNormal,
double rotationRate,
double dragCoeff,
double absorptionCoeff,
double reflectionCoeff)
Build a spacecraft model with linear rotation of solar array.
|
BoxAndSolarArraySpacecraft(BoxAndSolarArraySpacecraft.Facet[] facets,
PVCoordinatesProvider sun,
double solarArrayArea,
Vector3D solarArrayAxis,
AbsoluteDate referenceDate,
Vector3D referenceNormal,
double rotationRate,
double dragCoeff,
double liftRatio,
double absorptionCoeff,
double reflectionCoeff)
Build a spacecraft model with linear rotation of solar array.
|
BoxAndSolarArraySpacecraft(double xLength,
double yLength,
double zLength,
PVCoordinatesProvider sun,
double solarArrayArea,
Vector3D solarArrayAxis,
AbsoluteDate referenceDate,
Vector3D referenceNormal,
double rotationRate,
double dragCoeff,
double absorptionCoeff,
double reflectionCoeff)
Build a spacecraft model with linear rotation of solar array.
|
BoxAndSolarArraySpacecraft(double xLength,
double yLength,
double zLength,
PVCoordinatesProvider sun,
double solarArrayArea,
Vector3D solarArrayAxis,
AbsoluteDate referenceDate,
Vector3D referenceNormal,
double rotationRate,
double dragCoeff,
double liftRatio,
double absorptionCoeff,
double reflectionCoeff)
Build a spacecraft model with linear rotation of solar array.
|
HarmonicParametricAcceleration(Vector3D direction,
AttitudeProvider attitudeOverride,
String prefix,
AbsoluteDate referenceDate,
double fundamentalPeriod,
int harmonicMultiplier)
Simple constructor.
|
HarmonicParametricAcceleration(Vector3D direction,
boolean isInertial,
String prefix,
AbsoluteDate referenceDate,
double fundamentalPeriod,
int harmonicMultiplier)
Simple constructor.
|
PolynomialParametricAcceleration(Vector3D direction,
AttitudeProvider attitudeOverride,
String prefix,
AbsoluteDate referenceDate,
int degree)
Simple constructor.
|
PolynomialParametricAcceleration(Vector3D direction,
boolean isInertial,
String prefix,
AbsoluteDate referenceDate,
int degree)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
Vector3D |
IsotropicDrag.dragAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
double density,
Vector3D relativeVelocity,
double[] parameters)
Compute the acceleration due to drag.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
JB2008InputParameters.getMaxDate()
Gets the available data range maximum date.
|
AbsoluteDate |
NRLMSISE00InputParameters.getMaxDate()
Gets the available data range maximum date.
|
AbsoluteDate |
DTM2000InputParameters.getMaxDate()
Gets the available data range maximum date.
|
AbsoluteDate |
JB2008InputParameters.getMinDate()
Gets the available data range minimum date.
|
AbsoluteDate |
NRLMSISE00InputParameters.getMinDate()
Gets the available data range minimum date.
|
AbsoluteDate |
DTM2000InputParameters.getMinDate()
Gets the available data range minimum date.
|
| 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 |
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.
|
double |
DTM2000InputParameters.getMeanFlux(AbsoluteDate date)
Get the value of the mean solar flux.
|
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.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
MarshallSolarActivityFutureEstimation.getMaxDate()
Gets the available data range maximum date.
|
AbsoluteDate |
MarshallSolarActivityFutureEstimation.getMinDate()
Gets the available data range minimum date.
|
| 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.
|
double |
MarshallSolarActivityFutureEstimation.getMeanFlux(AbsoluteDate date)
Get the value of the mean solar flux.
|
double |
MarshallSolarActivityFutureEstimation.getThreeHourlyKP(AbsoluteDate date)
Get the value of the 3 hours geomagnetic index.
|
| Modifier and Type | Method and Description |
|---|---|
double[] |
HolmesFeatherstoneAttractionModel.gradient(AbsoluteDate date,
Vector3D position,
double 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.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
CachedNormalizedSphericalHarmonicsProvider.getReferenceDate()
Get the reference date for the harmonics.
|
AbsoluteDate |
SphericalHarmonicsProvider.getReferenceDate()
Get the reference date for the harmonics.
|
| Modifier and Type | Method and Description |
|---|---|
double |
CachedNormalizedSphericalHarmonicsProvider.getOffset(AbsoluteDate date)
Get the offset from
reference date for the harmonics. |
double |
SphericalHarmonicsProvider.getOffset(AbsoluteDate date)
Get the offset from
reference date for the harmonics. |
NormalizedSphericalHarmonicsProvider.NormalizedSphericalHarmonics |
CachedNormalizedSphericalHarmonicsProvider.onDate(AbsoluteDate date)
Get the normalized spherical harmonic coefficients at a specific instance in time.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
SmallManeuverAnalyticalModel.getDate()
Get the date of the maneuver.
|
| Modifier and Type | Method and Description |
|---|---|
void |
ImpulseManeuver.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
ConstantThrustManeuver.init(SpacecraftState s0,
AbsoluteDate t)
Initialize the force model at the start of propagation.
|
| Constructor and Description |
|---|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
double thrust,
double isp,
Vector3D direction)
Simple constructor for a constant direction and constant thrust.
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
double thrust,
double isp,
Vector3D direction,
String driversNamePrefix)
Simple constructor for a constant direction and constant thrust.
|
| Modifier and Type | Method and Description |
|---|---|
double |
SolarRadiationPressure.getLightingRatio(Vector3D position,
Frame frame,
AbsoluteDate date)
Get the lighting ratio ([0-1]).
|
Vector3D |
IsotropicRadiationSingleCoefficient.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters)
Compute the acceleration due to radiation pressure.
|
Vector3D |
RadiationSensitive.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters)
Compute the acceleration due to radiation pressure.
|
Vector3D |
IsotropicRadiationClassicalConvention.radiationPressureAcceleration(AbsoluteDate date,
Frame frame,
Vector3D position,
Rotation rotation,
double mass,
Vector3D flux,
double[] parameters)
Compute the acceleration due to radiation pressure.
|
Vector3D |
IsotropicRadiationCNES95Convention.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.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
Transform.getDate()
Get the date.
|
AbsoluteDate |
EOPEntry.getDate()
Get the date.
|
AbsoluteDate |
FieldTransform.getDate()
Get the date.
|
AbsoluteDate |
EOPHistory.getEndDate()
Get the date of the last available Earth Orientation Parameters.
|
AbsoluteDate |
HelmertTransformation.getEpoch()
Get the reference epoch of the transform.
|
AbsoluteDate |
EOPHistory.getStartDate()
Get the date of the first available Earth Orientation Parameters.
|
| Modifier and Type | Method and Description |
|---|---|
List<FieldTransform<T>> |
FieldTransformGenerator.generate(AbsoluteDate existingDate,
AbsoluteDate date)
Generate a chronologically sorted list of entries to be cached.
|
List<Transform> |
TransformGenerator.generate(AbsoluteDate existingDate,
AbsoluteDate date)
Generate a chronologically sorted list of entries to be cached.
|
double |
TopocentricFrame.getAzimuth(Vector3D extPoint,
Frame frame,
AbsoluteDate date)
Get the azimuth of a point with regards to the topocentric frame center point.
|
double |
TopocentricFrame.getElevation(Vector3D extPoint,
Frame frame,
AbsoluteDate date)
Get the elevation of a point with regards to the local point.
|
double[] |
EOPHistory.getEquinoxNutationCorrection(AbsoluteDate date)
Get the correction to the nutation parameters for equinox-based paradigm.
|
Frame |
Frame.getFrozenFrame(Frame reference,
AbsoluteDate freezingDate,
String frozenName)
Get a new version of the instance, frozen with respect to a reference frame.
|
double |
EOPHistory.getLOD(AbsoluteDate date)
Get the LoD (Length of Day) value.
|
protected Stream<EOPEntry> |
EOPHistory.getNeighbors(AbsoluteDate central)
Get the entries surrounding a central date.
|
static Transform |
FramesFactory.getNonInterpolatingTransform(Frame from,
Frame to,
AbsoluteDate date)
Get the transform between two frames, suppressing all interpolation.
|
double[] |
EOPHistory.getNonRotatinOriginNutationCorrection(AbsoluteDate date)
Get the correction to the nutation parameters for Non-Rotating Origin paradigm.
|
PoleCorrection |
EOPHistory.getPoleCorrection(AbsoluteDate date)
Get the pole IERS Reference Pole correction.
|
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.
|
Transform |
HelmertTransformation.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
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 |
FixedTransformProvider.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) |
Transform |
Frame.getTransformTo(Frame destination,
AbsoluteDate date)
Get the transform from the instance to another frame.
|
double |
EOPHistory.getUT1MinusUTC(AbsoluteDate date)
Get the UT1-UTC value.
|
protected boolean |
EOPHistory.hasDataFor(AbsoluteDate date)
Check if the cache has data for the given date using
EOPHistory.getStartDate() and EOPHistory.getEndDate(). |
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.
|
Transform |
LOFType.transformFromInertial(AbsoluteDate date,
PVCoordinates pv)
Get the transform from an inertial frame defining position-velocity and the local orbital frame.
|
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 |
|---|
HelmertTransformation(AbsoluteDate epoch,
double t1,
double t2,
double t3,
double r1,
double r2,
double r3,
double t1Dot,
double t2Dot,
double t3Dot,
double r1Dot,
double r2Dot,
double r3Dot)
Build a transform from its primitive operations.
|
InterpolatingTransformProvider(TransformProvider rawProvider,
CartesianDerivativesFilter cFilter,
AngularDerivativesFilter aFilter,
AbsoluteDate earliest,
AbsoluteDate latest,
int gridPoints,
double step,
int maxSlots,
double maxSpan,
double newSlotInterval)
Simple constructor.
|
ShiftingTransformProvider(TransformProvider rawProvider,
CartesianDerivativesFilter cFilter,
AngularDerivativesFilter aFilter,
AbsoluteDate earliest,
AbsoluteDate latest,
int gridPoints,
double step,
int maxSlots,
double maxSpan,
double newSlotInterval)
Simple constructor.
|
Transform(AbsoluteDate date,
AngularCoordinates angular)
Build a rotation transform.
|
Transform(AbsoluteDate date,
PVCoordinates cartesian)
Build a translation transform, with its first time derivative.
|
Transform(AbsoluteDate date,
Rotation rotation)
Build a rotation transform.
|
Transform(AbsoluteDate date,
Rotation rotation,
Vector3D rotationRate)
Build a rotation transform.
|
Transform(AbsoluteDate date,
Rotation rotation,
Vector3D rotationRate,
Vector3D rotationAcceleration)
Build a rotation transform.
|
Transform(AbsoluteDate date,
Transform first,
Transform second)
Build a transform by combining two existing ones.
|
Transform(AbsoluteDate date,
Vector3D translation)
Build a translation transform.
|
Transform(AbsoluteDate date,
Vector3D translation,
Vector3D velocity)
Build a translation transform, with its first time derivative.
|
Transform(AbsoluteDate date,
Vector3D translation,
Vector3D velocity,
Vector3D acceleration)
Build a translation transform, with its first and second time derivatives.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
GPSAlmanac.getDate() |
AbsoluteDate |
DOP.getDate()
Gets the calculation date of the DOP.
|
| 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. |
| Constructor and Description |
|---|
DOP(GeodeticPoint location,
AbsoluteDate date,
int gnssNb,
double gdop,
double pdop,
double hdop,
double vdop,
double tdop)
Constructor.
|
| Modifier and Type | Method and Description |
|---|---|
GeodeticPoint |
Geoid.getIntersectionPoint(Line lineInFrame,
Vector3D closeInFrame,
Frame frame,
AbsoluteDate date)
Get the intersection point of a line with the surface of the body.
|
double |
Geoid.getUndulation(double geodeticLatitude,
double longitude,
AbsoluteDate date)
Gets the Undulation of the Geoid, N at the given position.
|
double |
KlobucharIonoModel.pathDelay(AbsoluteDate date,
GeodeticPoint geo,
double elevation,
double azimuth)
Calculates the ionospheric path delay for the signal path from a ground
station to a satellite.
|
double |
IonosphericModel.pathDelay(AbsoluteDate date,
GeodeticPoint geo,
double elevation,
double azimuth)
Calculates the ionospheric path delay for the signal path from a ground
station to a satellite.
|
Vector3D |
Geoid.projectToGround(Vector3D point,
AbsoluteDate date,
Frame frame) |
GeodeticPoint |
Geoid.transform(Vector3D point,
Frame frame,
AbsoluteDate date)
Transform a Cartesian point to a surface-relative point.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
Orbit.getDate()
Get the date of orbital parameters.
|
| Modifier and Type | Method and Description |
|---|---|
TimeStampedPVCoordinates |
Orbit.getPVCoordinates(AbsoluteDate otherDate,
Frame otherFrame)
Get the
PVCoordinates of the body in the selected frame. |
CircularOrbit |
CircularOrbit.interpolate(AbsoluteDate date,
Stream<Orbit> sample)
Get an interpolated instance.
|
KeplerianOrbit |
KeplerianOrbit.interpolate(AbsoluteDate date,
Stream<Orbit> sample)
Get an interpolated instance.
|
CartesianOrbit |
CartesianOrbit.interpolate(AbsoluteDate date,
Stream<Orbit> sample)
Get an interpolated instance.
|
EquinoctialOrbit |
EquinoctialOrbit.interpolate(AbsoluteDate date,
Stream<Orbit> sample)
Get an interpolated instance.
|
abstract Orbit |
OrbitType.mapArrayToOrbit(double[] array,
double[] arrayDot,
PositionAngle type,
AbsoluteDate date,
double mu,
Frame frame)
Convert state array to orbital parameters.
|
| Constructor and Description |
|---|
CartesianOrbit(PVCoordinates pvaCoordinates,
Frame frame,
AbsoluteDate date,
double mu)
Constructor from Cartesian parameters.
|
CircularOrbit(double a,
double ex,
double ey,
double i,
double raan,
double alpha,
double aDot,
double exDot,
double eyDot,
double iDot,
double raanDot,
double alphaDot,
PositionAngle type,
Frame frame,
AbsoluteDate date,
double mu)
Creates a new instance.
|
CircularOrbit(double a,
double ex,
double ey,
double i,
double raan,
double alpha,
PositionAngle type,
Frame frame,
AbsoluteDate date,
double mu)
Creates a new instance.
|
CircularOrbit(PVCoordinates pvCoordinates,
Frame frame,
AbsoluteDate date,
double mu)
Constructor from Cartesian parameters.
|
EquinoctialOrbit(double a,
double ex,
double ey,
double hx,
double hy,
double l,
double aDot,
double exDot,
double eyDot,
double hxDot,
double hyDot,
double lDot,
PositionAngle type,
Frame frame,
AbsoluteDate date,
double mu)
Creates a new instance.
|
EquinoctialOrbit(double a,
double ex,
double ey,
double hx,
double hy,
double l,
PositionAngle type,
Frame frame,
AbsoluteDate date,
double mu)
Creates a new instance.
|
EquinoctialOrbit(PVCoordinates pvCoordinates,
Frame frame,
AbsoluteDate date,
double mu)
Constructor from Cartesian parameters.
|
KeplerianOrbit(double a,
double e,
double i,
double pa,
double raan,
double anomaly,
double aDot,
double eDot,
double iDot,
double paDot,
double raanDot,
double anomalyDot,
PositionAngle type,
Frame frame,
AbsoluteDate date,
double mu)
Creates a new instance.
|
KeplerianOrbit(double a,
double e,
double i,
double pa,
double raan,
double anomaly,
PositionAngle type,
Frame frame,
AbsoluteDate date,
double mu)
Creates a new instance.
|
KeplerianOrbit(PVCoordinates pvCoordinates,
Frame frame,
AbsoluteDate date,
double mu)
Constructor from Cartesian parameters.
|
Orbit(Frame frame,
AbsoluteDate date,
double mu)
Default constructor.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
SpacecraftState.getDate()
Get the date.
|
AbsoluteDate |
BoundedPropagator.getMaxDate()
Get the last date of the range.
|
AbsoluteDate |
BoundedPropagator.getMinDate()
Get the first date of the range.
|
protected AbsoluteDate |
AbstractPropagator.getStartDate()
Get the start date.
|
| Modifier and Type | Method and Description |
|---|---|
TimeStampedPVCoordinates |
AbstractPropagator.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
SpacecraftState |
SpacecraftState.interpolate(AbsoluteDate date,
Stream<SpacecraftState> 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.
|
protected void |
AbstractPropagator.setStartDate(AbsoluteDate startDate)
Set a start date.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
Ephemeris.getMaxDate()
Get the last date of the range.
|
AbsoluteDate |
AggregateBoundedPropagator.getMaxDate() |
AbsoluteDate |
Ephemeris.getMinDate()
Get the first date of the range.
|
AbsoluteDate |
AggregateBoundedPropagator.getMinDate() |
| Modifier and Type | Method and Description |
|---|---|
protected SpacecraftState |
AbstractAnalyticalPropagator.acceptStep(OrekitStepInterpolator interpolator,
AbsoluteDate target,
double epsilon)
Accept a step, triggering events and step handlers.
|
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 double |
EcksteinHechlerPropagator.getMass(AbsoluteDate date)
Get the mass.
|
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 double |
KeplerianPropagator.getMass(AbsoluteDate 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.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
protected double |
GPSPropagator.getMass(AbsoluteDate date)
Get the mass.
|
PVCoordinates |
GPSPropagator.propagateInEcef(AbsoluteDate date)
Gets the PVCoordinates of the GPS SV in
ECEF frame. |
protected Orbit |
GPSPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
TLE.getDate()
Get the TLE current date.
|
AbsoluteDate |
TLESeries.getFirstDate()
Deprecated.
Get the start date of the series.
|
AbsoluteDate |
TLESeries.getLastDate()
Deprecated.
Get the last date of the series.
|
| Modifier and Type | Method and Description |
|---|---|
TLE |
TLESeries.getClosestTLE(AbsoluteDate date)
Deprecated.
Get the closest TLE to the selected date.
|
protected double |
TLEPropagator.getMass(AbsoluteDate date)
Get the mass.
|
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.
|
protected Orbit |
TLEPropagator.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
| Constructor and Description |
|---|
TLE(int satelliteNumber,
char classification,
int launchYear,
int launchNumber,
String launchPiece,
int ephemerisType,
int elementNumber,
AbsoluteDate epoch,
double meanMotion,
double meanMotionFirstDerivative,
double meanMotionSecondDerivative,
double e,
double i,
double pa,
double raan,
double meanAnomaly,
int revolutionNumberAtEpoch,
double bStar)
Simple constructor from already parsed elements.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
AbstractPropagatorBuilder.getInitialOrbitDate()
Get the date of the initial orbit.
|
AbsoluteDate |
PropagatorBuilder.getInitialOrbitDate()
Get the date of the initial orbit.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
DateDetector.getDate()
Get the current event date according to the propagator.
|
AbsoluteDate |
EventState.getEventDate()
Get the occurrence time of the event triggered in the current
step.
|
AbsoluteDate |
EventState.EventOccurrence.getStopDate()
Get the new time for a stop action.
|
| Modifier and Type | Method and Description |
|---|---|
void |
DateDetector.addEventDate(AbsoluteDate target)
Add an event date.
|
void |
NegateDetector.init(SpacecraftState s0,
AbsoluteDate t) |
default void |
EventDetector.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
EventShifter.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
EventSlopeFilter.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
EventEnablingPredicateFilter.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
AbstractDetector.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
EventState.init(SpacecraftState s0,
AbsoluteDate t)
Initialize event handler at the start of a propagation.
|
void |
BooleanDetector.init(SpacecraftState s0,
AbsoluteDate t) |
| Constructor and Description |
|---|
DateDetector(AbsoluteDate target)
Build a new instance.
|
| Modifier and Type | Method and Description |
|---|---|
default void |
EventHandler.init(SpacecraftState initialState,
AbsoluteDate target)
Initialize event handler at the start of a propagation.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
IntegratedEphemeris.getMaxDate()
Get the last date of the range.
|
AbsoluteDate |
IntegratedEphemeris.getMinDate()
Get the first date of the range.
|
AbsoluteDate |
StateMapper.getReferenceDate()
Get reference date.
|
AbsoluteDate |
StateMapper.mapDoubleToDate(double t)
Map the raw double time offset to a date.
|
AbsoluteDate |
StateMapper.mapDoubleToDate(double t,
AbsoluteDate date)
Map the raw double time offset to a date.
|
| Modifier and Type | Method and Description |
|---|---|
protected SpacecraftState |
IntegratedEphemeris.basicPropagate(AbsoluteDate date)
Propagate an orbit without any fancy features.
|
protected void |
AbstractIntegratedPropagator.beforeIntegration(SpacecraftState initialState,
AbsoluteDate tEnd)
Method called just before integration.
|
protected abstract StateMapper |
AbstractIntegratedPropagator.createMapper(AbsoluteDate referenceDate,
double mu,
OrbitType orbitType,
PositionAngle positionAngleType,
AttitudeProvider attitudeProvider,
Frame frame)
Create a mapper between raw double components and spacecraft state.
|
protected double |
IntegratedEphemeris.getMass(AbsoluteDate date)
Get the mass.
|
TimeStampedPVCoordinates |
IntegratedEphemeris.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
default void |
AbstractIntegratedPropagator.MainStateEquations.init(SpacecraftState initialState,
AbsoluteDate target)
Initialize the equations at the start of propagation.
|
void |
ModeHandler.initialize(boolean activateHandlers,
AbsoluteDate targetDate)
Initialize the mode handler.
|
abstract SpacecraftState |
StateMapper.mapArrayToState(AbsoluteDate date,
double[] y,
double[] yDot,
boolean meanOnly)
Map the raw double components to a spacecraft state.
|
double |
StateMapper.mapDateToDouble(AbsoluteDate date)
Map a date to a raw double time offset.
|
AbsoluteDate |
StateMapper.mapDoubleToDate(double t,
AbsoluteDate date)
Map the raw double time offset to a date.
|
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.
|
protected Orbit |
IntegratedEphemeris.propagateOrbit(AbsoluteDate date)
Extrapolate an orbit up to a specific target date.
|
| Constructor and Description |
|---|
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.
|
StateMapper(AbsoluteDate referenceDate,
double mu,
OrbitType orbitType,
PositionAngle positionAngleType,
AttitudeProvider attitudeProvider,
Frame frame)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
protected StateMapper |
NumericalPropagator.createMapper(AbsoluteDate referenceDate,
double mu,
OrbitType orbitType,
PositionAngle positionAngleType,
AttitudeProvider attitudeProvider,
Frame frame)
Create a mapper between raw double components and spacecraft state.
|
TimeStampedPVCoordinates |
NumericalPropagator.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
| Modifier and Type | Method and Description |
|---|---|
SpacecraftState |
OrekitStepInterpolator.getInterpolatedState(AbsoluteDate date)
Get the state at interpolated date.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
protected void |
DSSTPropagator.beforeIntegration(SpacecraftState initialState,
AbsoluteDate tEnd)
Method called just before integration.
|
protected StateMapper |
DSSTPropagator.createMapper(AbsoluteDate referenceDate,
double mu,
OrbitType ignoredOrbitType,
PositionAngle ignoredPositionAngleType,
AttitudeProvider attitudeProvider,
Frame frame)
Create a mapper between raw double components and spacecraft state.
|
| Modifier and Type | Method and Description |
|---|---|
Map<String,double[]> |
ShortPeriodTerms.getCoefficients(AbsoluteDate date,
Set<String> selected)
Computes the coefficients involved in the contributions.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
AuxiliaryElements.getDate()
Get the date of the orbit.
|
| Modifier and Type | Method and Description |
|---|---|
void |
ShortPeriodicsInterpolatedCoefficient.addGridPoint(AbsoluteDate date,
double[] value)
Add a point to the interpolation grid.
|
double[] |
ShortPeriodicsInterpolatedCoefficient.value(AbsoluteDate date)
Compute the value of the coefficient.
|
| Modifier and Type | Field and Description |
|---|---|
static AbsoluteDate |
AbsoluteDate.CCSDS_EPOCH
Reference epoch for CCSDS Time Code Format (CCSDS 301.0-B-4):
1958-01-01T00:00:00 International Atomic Time (not UTC).
|
static AbsoluteDate |
AbsoluteDate.FIFTIES_EPOCH
Reference epoch for 1950 dates: 1950-01-01T00:00:00 Terrestrial Time.
|
static AbsoluteDate |
AbsoluteDate.FUTURE_INFINITY
Dummy date at infinity in the future direction.
|
static AbsoluteDate |
AbsoluteDate.GALILEO_EPOCH
Reference epoch for Galileo System Time: 1999-08-22T00:00:00 UTC.
|
static AbsoluteDate |
AbsoluteDate.GPS_EPOCH
Reference epoch for GPS weeks: 1980-01-06T00:00:00 GPS time.
|
static AbsoluteDate |
AbsoluteDate.J2000_EPOCH
J2000.0 Reference epoch: 2000-01-01T12:00:00 Terrestrial Time (not UTC).
|
static AbsoluteDate |
AbsoluteDate.JAVA_EPOCH
Java Reference epoch: 1970-01-01T00:00:00 Universal Time Coordinate.
|
static AbsoluteDate |
AbsoluteDate.JULIAN_EPOCH
Reference epoch for julian dates: -4712-01-01T12:00:00 Terrestrial Time.
|
static AbsoluteDate |
AbsoluteDate.MODIFIED_JULIAN_EPOCH
Reference epoch for modified julian dates: 1858-11-17T00:00:00 Terrestrial Time.
|
static AbsoluteDate |
AbsoluteDate.PAST_INFINITY
Dummy date at infinity in the past direction.
|
| Modifier and Type | Method and Description |
|---|---|
static AbsoluteDate |
AbsoluteDate.createBesselianEpoch(double besselianEpoch)
Build an instance corresponding to a Besselian Epoch (BE).
|
static AbsoluteDate |
AbsoluteDate.createGPSDate(int weekNumber,
double milliInWeek)
Build an instance corresponding to a GPS date.
|
static AbsoluteDate |
AbsoluteDate.createJDDate(int jd,
double secondsSinceNoon,
TimeScale timeScale)
Build an instance corresponding to a Julian Day date.
|
static AbsoluteDate |
AbsoluteDate.createJulianEpoch(double julianEpoch)
Build an instance corresponding to a Julian Epoch (JE).
|
static AbsoluteDate |
AbsoluteDate.createMJDDate(int mjd,
double secondsInDay,
TimeScale timeScale)
Build an instance corresponding to a Modified Julian Day date.
|
AbsoluteDate |
TimeStamped.getDate()
Get the date.
|
AbsoluteDate |
AbsoluteDate.getDate()
Get the date.
|
AbsoluteDate |
UTCScale.getFirstKnownLeapSecond()
Get the date of the first known leap second.
|
AbsoluteDate |
UTCScale.getLastKnownLeapSecond()
Get the date of the last known leap second.
|
static AbsoluteDate |
AbsoluteDate.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 AbsoluteDate |
AbsoluteDate.parseCCSDSUnsegmentedTimeCode(byte preambleField1,
byte preambleField2,
byte[] timeField,
AbsoluteDate agencyDefinedEpoch)
Build an instance from a CCSDS Unsegmented Time Code (CUC).
|
AbsoluteDate |
AbsoluteDate.shiftedBy(double dt)
Get a time-shifted date.
|
AbsoluteDate |
FieldAbsoluteDate.toAbsoluteDate()
Transform the FieldAbsoluteDate in an AbsoluteDate.
|
| Modifier and Type | Method and Description |
|---|---|
int |
AbsoluteDate.compareTo(AbsoluteDate date)
Compare the instance with another date.
|
double |
AbsoluteDate.durationFrom(AbsoluteDate instant)
Compute the physically elapsed duration between two instants.
|
T |
FieldAbsoluteDate.durationFrom(AbsoluteDate instant)
Compute the physically elapsed duration between two instants.
|
default double |
TimeScale.getLeap(AbsoluteDate date)
Get the value of the previous leap.
|
double |
UTCScale.getLeap(AbsoluteDate date)
Get the value of the previous leap.
|
double |
GLONASSScale.getLeap(AbsoluteDate date)
Get the value of the previous leap.
|
default boolean |
TimeScale.insideLeap(AbsoluteDate date)
Check if date is within a leap second introduction in this time scale.
|
boolean |
UTCScale.insideLeap(AbsoluteDate date)
Check if date is within a leap second introduction in this time scale.
|
boolean |
GLONASSScale.insideLeap(AbsoluteDate date)
Check if date is within a leap second introduction in this time 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 int |
TimeScale.minuteDuration(AbsoluteDate date)
Check length of the current minute in this time scale.
|
int |
UTCScale.minuteDuration(AbsoluteDate date)
Check length of the current minute in this time scale.
|
int |
GLONASSScale.minuteDuration(AbsoluteDate date)
Check length of the current minute in this time scale.
|
double |
AbsoluteDate.offsetFrom(AbsoluteDate instant,
TimeScale timeScale)
Compute the apparent clock offset between two instant in the
perspective of a specific
time scale. |
double |
TimeScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
GMSTScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
UT1Scale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
TTScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
GalileoScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
TAIScale.offsetFromTAI(AbsoluteDate taiTime)
Get the offset to convert locations from
TAIScale to instance. |
double |
QZSSScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
TCGScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
GPSScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
UTCScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
GLONASSScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
TDBScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
double |
TCBScale.offsetFromTAI(AbsoluteDate date)
Get the offset to convert locations from
TAIScale to instance. |
static AbsoluteDate |
AbsoluteDate.parseCCSDSUnsegmentedTimeCode(byte preambleField1,
byte preambleField2,
byte[] timeField,
AbsoluteDate agencyDefinedEpoch)
Build an instance from a CCSDS Unsegmented Time Code (CUC).
|
double[] |
TimeVectorFunction.value(AbsoluteDate date)
Compute a function of time.
|
double |
TimeScalarFunction.value(AbsoluteDate date)
Compute a function of time.
|
| Constructor and Description |
|---|
AbsoluteDate(AbsoluteDate since,
double elapsedDuration)
Build an instance from an elapsed duration since to another instant.
|
AbsoluteDate(AbsoluteDate reference,
double apparentOffset,
TimeScale timeScale)
Build an instance from an apparent clock offset with respect to another
instant in the perspective of a specific
time scale. |
FieldAbsoluteDate(AbsoluteDate since,
T elapsedDuration)
Build an instance from an elapsed duration since to another instant.
|
FieldAbsoluteDate(Field<T> field,
AbsoluteDate date)
Build an instance from an AbsoluteDate.
|
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
TimeStampedPVCoordinates.getDate()
Get the date.
|
AbsoluteDate |
FieldTimeSpanMap.Transition.getDate()
Get the transition absolute date.
|
AbsoluteDate |
TimeStampedAngularCoordinates.getDate()
Get the date.
|
AbsoluteDate |
TimeSpanMap.Transition.getDate()
Get the transition date.
|
AbsoluteDate |
IERSConventions.getNutationReferenceEpoch()
Get the reference epoch for fundamental nutation arguments.
|
AbsoluteDate |
ParameterDriver.getReferenceDate()
Get current reference date.
|
AbsoluteDate |
SecularAndHarmonic.getReferenceDate()
Get the reference date.
|
| Modifier and Type | Method and Description |
|---|---|
void |
SecularAndHarmonic.addPoint(AbsoluteDate date,
double osculatingValue)
Add a fitting point.
|
void |
TimeSpanMap.addValidAfter(T entry,
AbsoluteDate earliestValidityDate)
Add an entry valid after a limit date.
|
void |
TimeSpanMap.addValidBefore(T entry,
AbsoluteDate latestValidityDate)
Add an entry valid before a limit date.
|
double[] |
SecularAndHarmonic.approximateAsPolynomialOnly(int combinedDegree,
AbsoluteDate combinedReference,
int meanDegree,
int meanHarmonics,
AbsoluteDate start,
AbsoluteDate end,
double step)
Approximate an already fitted model to polynomial only terms.
|
double |
IERSConventions.evaluateTC(AbsoluteDate date)
Evaluate the date offset between the current date and the
reference date. |
List<T> |
TimeStampedGenerator.generate(AbsoluteDate existingDate,
AbsoluteDate date)
Generate a chronologically sorted list of entries to be cached.
|
T |
TimeSpanMap.get(AbsoluteDate date)
Get the entry valid at a specified date.
|
Stream<T> |
TimeStampedCache.getNeighbors(AbsoluteDate central)
Get the entries surrounding a central date.
|
Stream<T> |
GenericTimeStampedCache.getNeighbors(AbsoluteDate central)
Get the entries surrounding a central date.
|
Stream<T> |
ImmutableTimeStampedCache.getNeighbors(AbsoluteDate central)
Get the entries surrounding a central date.
|
TimeStampedPVCoordinates |
PVCoordinatesProvider.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
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 TimeStampedPVCoordinates |
TimeStampedPVCoordinates.interpolate(AbsoluteDate date,
CartesianDerivativesFilter filter,
Collection<TimeStampedPVCoordinates> sample)
Interpolate position-velocity.
|
static TimeStampedPVCoordinates |
TimeStampedPVCoordinates.interpolate(AbsoluteDate date,
CartesianDerivativesFilter filter,
Stream<TimeStampedPVCoordinates> sample)
Interpolate position-velocity.
|
double |
SecularAndHarmonic.meanDerivative(AbsoluteDate date,
int degree,
int harmonics)
Get mean derivative, truncated to first components.
|
double |
SecularAndHarmonic.meanSecondDerivative(AbsoluteDate date,
int degree,
int harmonics)
Get mean second derivative, truncated to first components.
|
double |
SecularAndHarmonic.meanValue(AbsoluteDate date,
int degree,
int harmonics)
Get mean value, truncated to first components.
|
double |
SecularAndHarmonic.osculatingDerivative(AbsoluteDate date)
Get fitted osculating derivative.
|
double |
SecularAndHarmonic.osculatingSecondDerivative(AbsoluteDate date)
Get fitted osculating second derivative.
|
double |
SecularAndHarmonic.osculatingValue(AbsoluteDate date)
Get fitted osculating value.
|
default void |
ParameterObserver.referenceDateChanged(AbsoluteDate previousReferenceDate,
ParameterDriver driver)
Notify that a parameter reference date has been changed.
|
void |
SecularAndHarmonic.resetFitting(AbsoluteDate date,
double... initialGuess)
Reset fitting.
|
void |
ParameterDriver.setReferenceDate(AbsoluteDate newReferenceDate)
Set reference date.
|
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.
|
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