Uses of Class
org.orekit.errors.OrekitException

Packages that use OrekitException

fr.cnes.sirius.patrius.assembly
fr.cnes.sirius.patrius.assembly.models
fr.cnes.sirius.patrius.assembly.properties
fr.cnes.sirius.patrius.bodies
fr.cnes.sirius.patrius.events
fr.cnes.sirius.patrius.events.multi
fr.cnes.sirius.patrius.events.postprocessing
fr.cnes.sirius.patrius.events.sensor
fr.cnes.sirius.patrius.forces
fr.cnes.sirius.patrius.forces.radiation
fr.cnes.sirius.patrius.groundstation
fr.cnes.sirius.patrius.guidance
fr.cnes.sirius.patrius.projections
fr.cnes.sirius.patrius.propagation
fr.cnes.sirius.patrius.propagation.events.multi
fr.cnes.sirius.patrius.propagation.numerical.multi
fr.cnes.sirius.patrius.propagation.precomputed.multi
fr.cnes.sirius.patrius.propagation.sampling.multi
fr.cnes.sirius.patrius.signalpropagation
fr.cnes.sirius.patrius.signalpropagation.iono
fr.cnes.sirius.patrius.stela
fr.cnes.sirius.patrius.stela.bodies
fr.cnes.sirius.patrius.stela.forces
fr.cnes.sirius.patrius.stela.forces.atmospheres
fr.cnes.sirius.patrius.stela.forces.drag
fr.cnes.sirius.patrius.stela.forces.gravity
fr.cnes.sirius.patrius.stela.forces.noninertial
fr.cnes.sirius.patrius.stela.forces.radiation
fr.cnes.sirius.patrius.stela.orbits
fr.cnes.sirius.patrius.stela.propagation
fr.cnes.sirius.patrius.stela.propagation.data
fr.cnes.sirius.patrius.tools.force.validation
fr.cnes.sirius.patrius.utils
fr.cnes.sirius.patrius.wrenches
fr.cnes.sirius.validate.files
fr.cnes.sirius.validate.mocks.ephemeris
org.orekit.attitudes	This package provides classes to represent simple attitudes.
org.orekit.attitudes.directions
org.orekit.attitudes.kinematics
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.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.atmospheres
org.orekit.forces.atmospheres.solarActivity
org.orekit.forces.atmospheres.solarActivity.specialized
org.orekit.forces.drag
org.orekit.forces.gravity
org.orekit.forces.gravity.potential	This package provides classes to read gravity field files and supports several different formats.
org.orekit.forces.gravity.tides
org.orekit.forces.gravity.tides.coefficients
org.orekit.forces.gravity.variations
org.orekit.forces.gravity.variations.coefficients
org.orekit.forces.maneuvers	This package provides models of simple maneuvers.
org.orekit.forces.radiation
org.orekit.forces.relativistic
org.orekit.frames	This package provides classes to handle frames and transforms between them.
org.orekit.frames.configuration
org.orekit.frames.configuration.eop
org.orekit.frames.configuration.libration
org.orekit.frames.transformations
org.orekit.models.earth	This package provides models that simulate certain physical phenomena experienced in the atmosphere of the earth.
org.orekit.orbits	This package provides classes to represent orbits.
org.orekit.parameter
org.orekit.propagation	This package provides tools to propagate orbital states with different methods.
org.orekit.propagation.analytical
org.orekit.propagation.analytical.covariance
org.orekit.propagation.analytical.tle	This package provides classes to read and extrapolate tle's.
org.orekit.propagation.analytical.twod
org.orekit.propagation.events	This package provides interfaces and classes dealing with events occurring during propagation.
org.orekit.propagation.events.multi
org.orekit.propagation.numerical
org.orekit.propagation.precomputed
org.orekit.propagation.sampling	This package provides interfaces and classes dealing with step handling during propagation.
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.
org.orekit.wrenches

Uses of OrekitException in fr.cnes.sirius.patrius.assembly

Methods in fr.cnes.sirius.patrius.assembly that throw OrekitException

void	Assembly.initMainPartFrame(SpacecraftState state) Initialize the main part's frame using a SpacecraftState as an input argument.
void	AssemblyBuilder.initMainPartFrame(SpacecraftState state) Sets up the main frame of the assembly from a "SpacecraftState" object.
void	MainPart.updateFrame(Transform inTransform)
void	Part.updateFrame(Transform inTransform)
void	IPart.updateFrame(Transform inTransform)
void	Assembly.updateMainPartFrame(SpacecraftState state) Updates the main part frame's transformation to its parent frame using a Transform as an input argument.
void	Assembly.updateMainPartFrame(Transform transform) Updates the main part frame's transformation to its parent frame using a Transform as an input argument.

Uses of OrekitException in fr.cnes.sirius.patrius.assembly.models

Methods in fr.cnes.sirius.patrius.assembly.models that throw OrekitException

void	RediffusedRadiativeModel.addDAccDParamRediffusedRadiativePressure(SpacecraftState s, Parameter param, double[] dAccdParam)
void	RediffusedRadiativeModel.addDAccDStateRediffusedRadiativePressure(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel)
void	DragLiftModel.addDDragAccDParam(SpacecraftState s, Parameter param, double density, Vector3D relativeVelocity, double[] dAccdParam) Compute acceleration derivatives with respect to ballistic coefficient.
void	GlobalAeroModel.addDDragAccDParam(SpacecraftState s, Parameter param, double density, Vector3D relativeVelocity, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters (the ballistic coefficient).
void	AeroModel.addDDragAccDParam(SpacecraftState s, Parameter param, double density, Vector3D relativeVelocity, double[] dAccdParam) Compute acceleration derivatives with respect to ballistic coefficient.
void	DragLiftModel.addDDragAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, double density, Vector3D acceleration, Vector3D relativeVelocity, boolean computeGradientPosition, boolean computeGradientVelocity) Compute acceleration derivatives with respect to state parameters (position and velocity).
void	GlobalAeroModel.addDDragAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, double density, Vector3D acceleration, Vector3D relativeVelocity, boolean computeGradientPosition, boolean computeGradientVelocity) Compute acceleration derivatives with respect to state parameters (position and velocity).
void	AeroModel.addDDragAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, double density, Vector3D acceleration, Vector3D relativeVelocity, boolean computeGradientPosition, boolean computeGradientVelocity) Compute acceleration derivatives with respect to state parameters (position and velocity).
void	DirectRadiativeModel.addDSRPAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam, Vector3D satSunVector) Compute acceleration derivatives with respect to additional parameters.
void	DirectRadiativeModel.addDSRPAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, Vector3D satSunVector) Compute acceleration derivatives with respect to state parameters.
double	SensorModel.celestialBodiesMaskingDistance(AbsoluteDate date) Computes the minimal euclidian distance to the celestial body shapes.
double	RFLinkBudgetModel.computeLinkBudget(AbsoluteDate date) Computes the link budget at a given date.
Vector3D	DragLiftModel.dragAcceleration(SpacecraftState state, double density, Vector3D relativeVelocity) Method to compute the aero acceleration, based on the assembly.
Vector3D	GlobalAeroModel.dragAcceleration(SpacecraftState state, double density, Vector3D relativeVelocity) Method to compute the aero acceleration, based on the assembly.
Vector3D	AeroModel.dragAcceleration(SpacecraftState state, double density, Vector3D relativeVelocity) Method to compute the aero acceleration, based on the assembly.
Wrench	AeroWrenchModel.dragWrench(SpacecraftState state, double density, Vector3D relativeVelocity)
Wrench	AeroWrenchModel.dragWrench(SpacecraftState state, double density, Vector3D relativeVelocity, Vector3D origin, Frame frame) Compute the torque due to radiation pressire.
protected static Vector3D	AeroModel.forceOnFacet(SpacecraftState state, IPart part, Assembly assembly, double density, Vector3D relativeVelocity) Method to compute the force for a plane model.
protected static Vector3D	DirectRadiativeModel.forceOnFacet(SpacecraftState state, IPart part, Vector3D flux) Method to compute the force for a plane model.
protected static Vector3D	AeroModel.forceOnSphere(SpacecraftState state, IPart part, double density, Vector3D relativeVelocity) Method to compute the force for a sphere model.
Matrix3D	InertiaSimpleModel.getInertiaMatrix(Frame frame, AbsoluteDate date) Getter for the inertia matrix of the spacecraft, expressed with respect to the MASS CENTER in a given frame.
Matrix3D	InertiaComputedModel.getInertiaMatrix(Frame frame, AbsoluteDate date) Getter for the inertia matrix of the spacecraft, expressed with respect to the MASS CENTER in a given frame.
Matrix3D	IInertiaModel.getInertiaMatrix(Frame frame, AbsoluteDate date) Getter for the inertia matrix of the spacecraft, expressed with respect to the MASS CENTER in a given frame.
Matrix3D	InertiaSimpleModel.getInertiaMatrix(Frame frame, AbsoluteDate date, Vector3D inertiaReferencePoint) Getter for the inertia matrix of the spacecraft, once expressed with respect to a point that can be different from the mass center.
Matrix3D	InertiaComputedModel.getInertiaMatrix(Frame frame, AbsoluteDate date, Vector3D inertiaReferencePoint) Getter for the inertia matrix of the spacecraft, once expressed with respect to a point that can be different from the mass center.
Matrix3D	IInertiaModel.getInertiaMatrix(Frame frame, AbsoluteDate date, Vector3D inertiaReferencePoint) Getter for the inertia matrix of the spacecraft, once expressed with respect to a point that can be different from the mass center.
double	SensorModel.getInhibitionTargetAngularRadius(AbsoluteDate date, int inhibitionFieldNumber) Computes the angular radius from the sensor of the main target at a date.
double	SensorModel.getInhibitTargetCenterToFieldAngle(AbsoluteDate date, int inhibitionFieldNumber) Computes the angular distance of the CENTER of an inhibition target to the border of the associated inhibition field at a date.
double	SensorModel.getMainTargetAngularRadius(AbsoluteDate date) Computes the angular radius from the sensor of the main target at a date.
Vector3D	InertiaSimpleModel.getMassCenter(Frame frame, AbsoluteDate date) Getter for the mass center.
Vector3D	InertiaComputedModel.getMassCenter(Frame frame, AbsoluteDate date) Getter for the mass center.
Vector3D	IInertiaModel.getMassCenter(Frame frame, AbsoluteDate date) Getter for the mass center.
Vector3D	SensorModel.getNormalisedTargetVectorInSensorFrame(AbsoluteDate date) Computes the target vector at a date in the sensor's frame.
PVCoordinates	SensorModel.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the sensor part in the selected frame.
Vector3D[]	SensorModel.getRefrenceAxis(Frame frame, AbsoluteDate date) Computes the reference axis of the sensor in a given frame at a date
Vector3D	SensorModel.getSightAxis(Frame frame, AbsoluteDate date) Computes the sight axis of the sensor in a given frame at a date
double	SensorModel.getTargetCenterFOVAngle(AbsoluteDate date) Computes the angular distance of the CENTER of the main target to the border of the main field of view at a date.
double[]	SensorModel.getTargetDihedralAngles(AbsoluteDate date) Computes the dihedral angles of the target at a date in the sensor's frame.
double	SensorModel.getTargetRefAxisAngle(AbsoluteDate date, int axisNumber)
double	SensorModel.getTargetRefAxisElevation(AbsoluteDate date, int axisNumber)
double	SensorModel.getTargetSightAxisAngle(AbsoluteDate date)
double	SensorModel.getTargetSightAxisElevation(AbsoluteDate date)
Vector3D	SensorModel.getTargetVectorInSensorFrame(AbsoluteDate date) Computes the target vector at a date in the sensor's frame.
boolean	SensorModel.isMainTargetInField(AbsoluteDate date) Checks if the main target at least partially is in the field of view at a date
boolean	SensorModel.noInhibition(AbsoluteDate date) Checks if at least an inhibition target is at least partially in its associated inhibition field at a date
Vector3D	DirectRadiativeModel.radiationPressureAcceleration(SpacecraftState state, Vector3D flux) Method to compute the radiation pressure acceleration, based on the assembly.
Wrench	DirectRadiativeWrenchModel.radiationWrench(SpacecraftState state, Vector3D flux) Compute the torque due to radiation pressire.
Wrench	DirectRadiativeWrenchModel.radiationWrench(SpacecraftState state, Vector3D flux, Vector3D origin, Frame frame) Compute the torque due to radiation pressire.
Vector3D	RediffusedRadiativeModel.rediffusedRadiationPressureAcceleration(SpacecraftState state, ElementaryFlux flux) Method to compute the rediffused radiation pressure acceleration, based on the assembly.
double	SensorModel.spacecraftsMaskingDistance(AbsoluteDate date) Computes the minimal euclidian distance to the spacecraft's shapes (GEOMERTY properties).
void	InertiaSimpleModel.updateMass(String part, double mass) Update the mass of the given part.
void	InertiaComputedModel.updateMass(String partName, double mass) Update the mass of the given part.
void	MassModel.updateMass(String partName, double newMass) Update the mass of the given part.
boolean	SensorModel.visibilityOk(AbsoluteDate date) Checks if the main target is in the field of view and no inhibition target in its inhibition field at a given date.

Constructors in fr.cnes.sirius.patrius.assembly.models that throw OrekitException

InertiaSimpleModel(double mass, Vector3D massCenter, Matrix3D inertiaMatrix, Frame frame, String partName)
Constructor for a simple inertia model.

InertiaSimpleModel(double mass, Vector3D massCenter, Matrix3D inertiaMatrix, Vector3D inertiaReferencePoint, Frame frame, String partName)
Constructor for a simple inertia model; the inertia matrix is expressed with respect to a point that can be different from the mass center.

RediffusedRadiativeModel(boolean inAlbedo, boolean inIr, double inK0Albedo, double inK0Ir, Assembly inAssembly)
Rediffused radiative model (the acceleration is computed from all the sub parts of the vehicle).

RediffusedRadiativeModel(boolean inAlbedo, boolean inIr, Parameter inK0Albedo, Parameter inK0Ir, Assembly inAssembly)
Rediffused radiative model (the acceleration is computed from all the sub parts of the vehicle).

Uses of OrekitException in fr.cnes.sirius.patrius.assembly.properties

Methods in fr.cnes.sirius.patrius.assembly.properties that throw OrekitException

void	MassEquation.computeDerivatives(SpacecraftState s, TimeDerivativesEquations adder) Compute the derivatives related to the additional state parameters.
double	AeroSphereProperty.getSphereRadius() Get the sphere radius.
double	RadiativeSphereProperty.getSphereRadius() Get the sphere radius.
void	MassProperty.updateMass(double newMass) Updates the mass of the part.

Constructors in fr.cnes.sirius.patrius.assembly.properties that throw OrekitException

AeroSphereProperty(Parameter inSphereArea, double dragCoef)
Constructor of this property giving the drag coef without the atmospheric height scale.

MassProperty(double inMass)
Constructor of this property.

MassProperty(Parameter inMass)
Constructor of this property using a Parameter.

RadiativeSphereProperty(Parameter inSphereArea)
Constructor with area.

Uses of OrekitException in fr.cnes.sirius.patrius.bodies

Methods in fr.cnes.sirius.patrius.bodies that throw OrekitException

Line	BasicBoardSun.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Get the line from the position in pvCoord to the Sun.
Vector3D	BasicBoardSun.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Get the direction of the sun.

Uses of OrekitException in fr.cnes.sirius.patrius.events

Methods in fr.cnes.sirius.patrius.events that throw OrekitException

Map<CodingEventDetector,PhenomenaList>	CodedEventsLogger.buildPhenomenaListMap(AbsoluteDateInterval definitionInterval, SpacecraftState duringState) Builds a map of PhenomenaList, one list per CodingEventDetector instance.
EventDetector.Action	EarthZoneDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle the event and choose what to do next.
EventDetector.Action	CombinedPhenomenaDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
EventDetector.Action	GenericCodingEventDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
EventDetector.Action	CentralBodyMaskCircularFOVDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a target in field of view outside eclipse reaching event and choose what to do next.
double	EarthZoneDetector.g(SpacecraftState s) Compute the value of the switching function.
double	CombinedPhenomenaDetector.g(SpacecraftState s) Compute the value of the switching function for a combination (AND or OR) of two phenomena. After computing the switching function of each detector and, if necessary, changing its sign to apply a general convention (g>0 if the phenomenon associated to an event is active), it returns one between the two g functions, according to the boolean operator.
double	GenericCodingEventDetector.g(SpacecraftState s)
double	CentralBodyMaskCircularFOVDetector.g(SpacecraftState s) The switching function is the minimum value between the eclipse detector g function and the circularFOVDetector
boolean	GenericCodingEventDetector.isStateActive(SpacecraftState state) Tells if the event state is "active" for the given input.
SpacecraftState	GenericCodingEventDetector.resetState(SpacecraftState oldState)

Uses of OrekitException in fr.cnes.sirius.patrius.events.multi

Methods in fr.cnes.sirius.patrius.events.multi that throw OrekitException

Map<MultiCodingEventDetector,PhenomenaList>	MultiCodedEventsLogger.buildPhenomenaListMap(AbsoluteDateInterval definitionInterval, Map<String,SpacecraftState> duringState) Builds a map of PhenomenaList, one list per MultiCodingEventDetector instance.
EventDetector.Action	MultiGenericCodingEventDetector.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward)
double	MultiGenericCodingEventDetector.g(Map<String,SpacecraftState> s)
boolean	MultiGenericCodingEventDetector.isStateActive(Map<String,SpacecraftState> states) Tells if the multi event state is "active" for the given input.
Map<String,SpacecraftState>	MultiGenericCodingEventDetector.resetStates(Map<String,SpacecraftState> oldStates)

Uses of OrekitException in fr.cnes.sirius.patrius.events.postprocessing

Constructors in fr.cnes.sirius.patrius.events.postprocessing that throw OrekitException

Timeline(CodedEventsLogger logger, AbsoluteDateInterval interval)
Builds an instance of the timeline from a CodedEventsLogger, generating the list of detected events and the list of corresponding phenomena.
These events and phenomena are the output of a propagation with events detector; the coherence between events and phenomena should be guaranteed by the detection process during propagation.

Uses of OrekitException in fr.cnes.sirius.patrius.events.sensor

Methods in fr.cnes.sirius.patrius.events.sensor that throw OrekitException

EventDetector.Action	SatToSatMutualVisibilityDetector.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward)
EventDetector.Action	SatToSatMutualVisibilityDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
EventDetector.Action	MaskingDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle "masking" event and choose what to do next.
EventDetector.Action	ExtremaSightAxisDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
EventDetector.Action	SensorVisibilityDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
EventDetector.Action	VisibilityFromStationDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle "visibility from station" event and choose what to do next.
EventDetector.Action	TargetInFieldOfViewDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
EventDetector.Action	RFVisibilityDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
EventDetector.Action	StationToSatMutualVisibilityDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
EventDetector.Action	SensorInhibitionDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward)
double	SatToSatMutualVisibilityDetector.g(Map<String,SpacecraftState> s)
double	SatToSatMutualVisibilityDetector.g(SpacecraftState s)
double	MaskingDetector.g(SpacecraftState s) Compute the value of the switching function.
double	ExtremaSightAxisDetector.g(SpacecraftState s) The switching function is specific case of the extrema three bodies angle detector.
double	SensorVisibilityDetector.g(SpacecraftState s)
double	VisibilityFromStationDetector.g(SpacecraftState s) Compute the value of the switching function.
double	TargetInFieldOfViewDetector.g(SpacecraftState s)
double	RFVisibilityDetector.g(SpacecraftState s) Compute the value of the switching function.
double	StationToSatMutualVisibilityDetector.g(SpacecraftState s)
double	SensorInhibitionDetector.g(SpacecraftState s)
protected Vector3D	AbstractDetectorWithTropoCorrection.getCorrectedVector(SpacecraftState s) Compute the apparent vector from the station to the spacecraft with tropospheric effects.
Map<String,SpacecraftState>	SatToSatMutualVisibilityDetector.resetStates(Map<String,SpacecraftState> oldStates)
void	SecondarySpacecraft.updateSpacecraftState(AbsoluteDate date) Updates the assembly frames at a given date from the orbit and attitude information provided by the propagator.

Uses of OrekitException in fr.cnes.sirius.patrius.forces

Methods in fr.cnes.sirius.patrius.forces that throw OrekitException

void	EmpiricalForce.addContribution(SpacecraftState state, TimeDerivativesEquations adder)
void	EmpiricalForce.addDAccDParam(SpacecraftState state, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	EmpiricalForce.addDAccDState(SpacecraftState state, double[][] dAccdPos, double[][] dAccdVel)
Vector3D	EmpiricalForce.computeAcceleration(PVCoordinates pv, LocalOrbitalFrame localFrameValidation, Vector3D vectorS, Frame frame, SpacecraftState state) Method to compute the acceleration.
Vector3D	EmpiricalForce.computeAcceleration(SpacecraftState state)

Uses of OrekitException in fr.cnes.sirius.patrius.forces.radiation

Methods in fr.cnes.sirius.patrius.forces.radiation that throw OrekitException

void	PatriusSolarRadiationPressure.addContribution(SpacecraftState s, TimeDerivativesEquations adder)
void	PatriusSolarRadiationPressure.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam)
void	PatriusSolarRadiationPressure.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel)
Vector3D	PatriusSolarRadiationPressure.computeAcceleration(SpacecraftState s)
static double	PatriusSolarRadiationPressure.getLightningRatio(PVCoordinatesProvider sun, Vector3D satSunVector, GeometricBodyShape earthModel, Vector3D position, Frame frame, AbsoluteDate date) Get the lightning ratio ([0-1]).

Uses of OrekitException in fr.cnes.sirius.patrius.groundstation

Methods in fr.cnes.sirius.patrius.groundstation that throw OrekitException

PVCoordinates	RFStationAntenna.getPVCoordinates(AbsoluteDate date, Frame frame)
PVCoordinates	GeometricStationAntenna.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the station antenna in the selected frame.

Uses of OrekitException in fr.cnes.sirius.patrius.guidance

Methods in fr.cnes.sirius.patrius.guidance that throw OrekitException

void	GuidanceProfile.checkDate(AbsoluteDate userDate) Check date validity
static AngularVelocitiesHarmonicProfile	GuidanceProfileBuilder.computeAngularVelocitiesHarmonicProfile(AttitudeLawLeg attitude, PVCoordinatesProvider provider, Frame frame, AbsoluteDate tref, double period, int order, KinematicsToolkit.IntegrationType integType, double integStep) Compute the angular velocities harmonic guidance profile.
Attitude	GuidanceProfile.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	QuaternionHarmonicProfile.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	QuaternionPolynomialProfile.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	AngularVelocitiesPolynomialProfile.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate userDate, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	AngularVelocitiesHarmonicProfile.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate userDate, Frame frame) Compute the attitude corresponding to an orbital state.
void	QuaternionHarmonicProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	QuaternionPolynomialProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	AngularVelocitiesPolynomialProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	AngularVelocitiesHarmonicProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.

Uses of OrekitException in fr.cnes.sirius.patrius.projections

Methods in fr.cnes.sirius.patrius.projections that throw OrekitException

List<GeodeticPoint>	AbstractProjection.applyInverseTo(double[] x, double[] y) Inversion transformation of arrays of x and y projected coordinates.
GeodeticPoint	Mercator.applyInverseTo(double x, double y) Inverse projection.
GeodeticPoint	GeneralizedFlamsteedSamson.applyInverseTo(double x, double y) Inverse projection.
GeodeticPoint	IProjection.applyInverseTo(double x, double y) Inverse projection.
GeodeticPoint	Mercator.applyInverseTo(double x, double y, double alt) This is the Two standard parallel Mercator Projection model.
GeodeticPoint	GeneralizedFlamsteedSamson.applyInverseTo(double x, double y, double alt) This is the Two standard parallel Mercator Projection model.
GeodeticPoint	IProjection.applyInverseTo(double x, double y, double alt) This is the Two standard parallel Mercator Projection model.
List<GeodeticPoint>	AbstractProjection.applyInverseTo(List<Vector2D> list) Inverse Projects a list of Vector2D (projected points) with a given projection.
Vector2D	Mercator.applyTo(double lat, double lon) Returns Easting value and Northing value in meters from latitude and longitude coordinates.
Vector2D	GeneralizedFlamsteedSamson.applyTo(double lat, double lon) Returns Easting value and Northing value in meters from latitude and longitude coordinates.
Vector2D	IProjection.applyTo(double lat, double lon) Returns Easting value and Northing value in meters from latitude and longitude coordinates.
Vector2D	Mercator.applyTo(GeodeticPoint geodeticPoint) Returns Easting value and Northing value in meters from geodetic coordinates.
Vector2D	GeneralizedFlamsteedSamson.applyTo(GeodeticPoint geodeticPoint) Returns Easting value and Northing value in meters from geodetic coordinates.
Vector2D	IProjection.applyTo(GeodeticPoint geodeticPoint) Returns Easting value and Northing value in meters from geodetic coordinates.
List<Vector2D>	AbstractProjection.applyTo(List<GeodeticPoint> list) Project a list of GeodeticPoints with a given projection.
List<Vector2D>	AbstractProjection.applyToAndDiscretize(GeodeticPoint from, GeodeticPoint to, double maxLength, boolean lastIncluded) Project two points, then discretize 2D the line.
double	ProjectionEllipsoid.computeBearing(GeodeticPoint gv1, GeodeticPoint gv2) Compute the bearing (azimuth) between two geodetic Points.
double	ProjectionEllipsoid.computeLoxodromicDistance(GeodeticPoint p1, GeodeticPoint p2) Loxodromic distance between P1 and P2.This is the distance of constant bearing (or along a line in Mercator).
GeodeticPoint	ProjectionEllipsoid.computePointAlongLoxodrome(GeodeticPoint p1, double distance, double azimuth) Compute the point coordinates from an origin point, an azimuth and a distance along the rhumb line (Loxodrome).
List<Vector2D>	AbstractProjection.discretizeAndApplyTo(List<GeodeticPoint> list, EnumLineProperty ltype, double maxLength) Discretizes a polygon conforming to a line property directive, and a maximum length of discretization.
List<Vector2D>	AbstractProjection.discretizeCircleAndApplyTo(List<GeodeticPoint> list, double maxLength) Discretize following great circle lines between vertices of polygon and project obtained points.
List<Vector2D>	AbstractProjection.discretizeRhumbAndApplyTo(List<GeodeticPoint> list, double maxLength) Project a rhumb line polygon, with the given projection.
List<GeodeticPoint>	ProjectionEllipsoid.discretizeRhumbLine(GeodeticPoint from, GeodeticPoint to, double maxLength) Discretize a rhumb line into N segments, between two points.

Uses of OrekitException in fr.cnes.sirius.patrius.propagation

Methods in fr.cnes.sirius.patrius.propagation that throw OrekitException

void	MultiPropagator.addInitialState(SpacecraftState initialState, String satId) Add a new spacecraft state to be propagated.
Map<String,SpacecraftState>	MultiPropagator.getInitialStates() Get the propagator initial states.

Constructors in fr.cnes.sirius.patrius.propagation that throw OrekitException

PVCoordinatePropagator(PVCoordinatesProvider pvCoordProvider, AbsoluteDate initDate, double mu, Frame frame)
Creates an instance of PVCoordinatePropagator without attitude and additional state providers

PVCoordinatePropagator(PVCoordinatesProvider pvCoordProvider, AbsoluteDate initDate, double mu, Frame frame, AttitudeProvider attProviderForces, AttitudeProvider attProviderEvents, List<AdditionalStateProvider> additionalStateProviders)
Creates an instance of PVCoordinatePropagator with PV, attitude for forces, attitude for events, and additional state providers given by the user.

Uses of OrekitException in fr.cnes.sirius.patrius.propagation.events.multi

Methods in fr.cnes.sirius.patrius.propagation.events.multi that throw OrekitException

EventDetector.Action	OneSatEventDetectorWrapper.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward)
abstract EventDetector.Action	MultiAbstractDetector.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward)
double	OneSatEventDetectorWrapper.g(Map<String,SpacecraftState> s)
abstract double	MultiAbstractDetector.g(Map<String,SpacecraftState> s)
Map<String,SpacecraftState>	OneSatEventDetectorWrapper.resetStates(Map<String,SpacecraftState> oldStates)
Map<String,SpacecraftState>	MultiAbstractDetector.resetStates(Map<String,SpacecraftState> oldStates)

Uses of OrekitException in fr.cnes.sirius.patrius.propagation.numerical.multi

Methods in fr.cnes.sirius.patrius.propagation.numerical.multi that throw OrekitException

void	MultiNumericalPropagator.addInitialState(SpacecraftState initialState, String satId)
PVCoordinates	MultiNumericalPropagator.getPVCoordinates(AbsoluteDate date, Frame frame, String satId) Get the PVCoordinates of the body in the selected frame.
SpacecraftState	MultiStateVectorInfo.mapArrayToState(double[] y, AbsoluteDate currentDate, OrbitType orbitType, PositionAngle angleType, AttitudeProvider attProviderForces, AttitudeProvider attProviderEvents, double mu, Frame integrationFrame, String satId) Extract a given SpacecraftState from the state vector.
Map<String,SpacecraftState>	MultiStateVectorInfo.mapArrayToStates(double[] y, AbsoluteDate currentDate, OrbitType orbitType, PositionAngle angleType, Map<String,AttitudeProvider> attProvidersForces, Map<String,AttitudeProvider> attProvidersEvents, Map<String,Double> mu, Map<String,Frame> integrationFrame) Convert state vector into a Map of SpacecraftState
void	MultiNumericalPropagator.setAdditionalStateTolerance(String name, double[] absTol, double[] relTol, String satId) Add additional state tolerances.
void	MultiNumericalPropagator.setOrbitFrame(String satId, Frame frame) Set a frame for propagation The initial state must have first been added using the MultiNumericalPropagator.addInitialState(SpacecraftState, String) method before defining the associated integration frame.
void	MultiNumericalPropagator.setOrbitTolerance(double[] absoluteTolerance, double[] relativeTolerance, String satId) Set the orbit tolerance of a defined state.

Uses of OrekitException in fr.cnes.sirius.patrius.propagation.precomputed.multi

Methods in fr.cnes.sirius.patrius.propagation.precomputed.multi that throw OrekitException

SpacecraftState	MultiIntegratedEphemeris.getInitialState() Get the propagator initial state.
PVCoordinates	MultiIntegratedEphemeris.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.

Constructors in fr.cnes.sirius.patrius.propagation.precomputed.multi that throw OrekitException

MultiIntegratedEphemeris(List<AbsoluteDate> startDatesIn, List<AbsoluteDate> minDatesIn, List<AbsoluteDate> maxDatesIn, OrbitType orbitTypeIn, PositionAngle angleTypeIn, AttitudeProvider attitudeProviderForces, AttitudeProvider attitudeProviderEvents, MultiStateVectorInfo stateInfos, List<ContinuousOutputModel> modelsIn, Frame referenceFrameIn, double muIn, String satIdIn)
Creates a new instance of IntegratedEphemeris.

Uses of OrekitException in fr.cnes.sirius.patrius.propagation.sampling.multi

Methods in fr.cnes.sirius.patrius.propagation.sampling.multi that throw OrekitException

Map<String,SpacecraftState>	MultiOrekitStepInterpolator.getInterpolatedStates() Get all the interpolated states.
Map<String,SpacecraftState>	MultiAdaptedStepHandler.getInterpolatedStates() Get all the interpolated states.

Uses of OrekitException in fr.cnes.sirius.patrius.signalpropagation

Methods in fr.cnes.sirius.patrius.signalpropagation that throw OrekitException

SignalPropagation	SignalPropagationModel.computeSignalPropagation(PVCoordinatesProvider transmitter, PVCoordinatesProvider receiver, AbsoluteDate date, SignalPropagationModel.FixedDate fixedDateType) Computes the signal propagation object in the void at a particular date
double	SignalPropagationModel.getSignalTropoCorrection(TroposphericCorrection correction, SignalPropagation signal, TopocentricFrame topo) Computes the tropospheric effects corrections to be applied to a given PropagationSignal object.
Vector3D	SignalPropagation.getVector(Frame expressionFrame)

Uses of OrekitException in fr.cnes.sirius.patrius.signalpropagation.iono

Methods in fr.cnes.sirius.patrius.signalpropagation.iono that throw OrekitException

double	BentModel.computeElectronicCont(AbsoluteDate date, Vector3D satellite, Frame frameSat) Computation of the electric content between the station and the satellite at a date.
double	BentModel.computeSignalDelay(AbsoluteDate date, Vector3D satellite, Frame frameSat)
double	IonosphericCorrection.computeSignalDelay(AbsoluteDate date, Vector3D satellite, Frame satFrame) Calculates the ionospheric signal delay for the signal path from the position of the transmitter and the receiver and the current date.
fr.cnes.sirius.patrius.signalpropagation.iono.USKData	USKProvider.getData(AbsoluteDate date, double r12) Returns the USK data for the Bent model.
fr.cnes.sirius.patrius.signalpropagation.iono.USKData	USKLoader.getData(AbsoluteDate date, double r12) Returns the USK data for the Bent model.
double	R12Provider.getR12(AbsoluteDate date) Provides the R12 value for the Bent model.
double	R12Loader.getR12(AbsoluteDate date)
void	USKLoader.loadData(InputStream input, String name) Load data from a stream.
void	R12Loader.loadData(InputStream input, String name)

Constructors in fr.cnes.sirius.patrius.signalpropagation.iono that throw OrekitException

R12Loader(String supportedFileName)
Constructor.

USKLoader(String fileName)
Creates a USK data file reader and load the file.

Uses of OrekitException in fr.cnes.sirius.patrius.stela

Methods in fr.cnes.sirius.patrius.stela that throw OrekitException

static Assembly	StelaSpacecraftFactory.createStelaCompatibleSpacecraft(String mainPartName, double mass, double dragArea, double dragCoefficient, double srpArea, double srpReflectionCoefficient) Utility method to create a STELA Assembly, made of a sphere with both radiative and aerodynamic properties.
static Assembly	StelaSpacecraftFactory.createStelaRadiativeSpacecraft(String mainPartName, double mass, double srpArea, double srpReflectionCoefficient) Utility method to create a STELA Assembly, made of a sphere with only radiative properties.
EventDetector.Action	PerigeeAltitudeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an altitude event and choose what to do next.
double	PerigeeAltitudeDetector.g(SpacecraftState s) Compute the value of the switching function.
static double[][]	JavaMathAdapter.matrixAdd(double[][] m1, double[][] m2) Add 2 matrices.
static double[][]	JavaMathAdapter.threeDMatrixVectorMultiply(double[][][] mat, double[] vect) Multiply an automatically-generated-3-dimensional matrix with a vector. Automatically generated 3D matrices have their rows and wideness inverted.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.bodies

Methods in fr.cnes.sirius.patrius.stela.bodies that throw OrekitException

double	GeodPosition.getGeodeticAltitude(Vector3D position) Compute geodetic altitude.
double	GeodPosition.getGeodeticLatitude(Vector3D position) Compute geodetic latitude.
double	GeodPosition.getGeodeticLongitude(Vector3D position, AbsoluteDate date) Compute the geodetic longitude at a given date.
PVCoordinates	MeeusSunStela.getPVCoordinates(AbsoluteDate date, Frame frame) Deprecated.
PVCoordinates	MeeusMoonStela.getPVCoordinates(AbsoluteDate date, Frame frame)
static void	MeeusSunStela.updateTransform(AbsoluteDate date, Frame frame) Deprecated. Update cached transform from FramesFactory.getMOD(boolean) to provided frame.
static void	MeeusMoonStela.updateTransform(AbsoluteDate date, Frame frame) Update cached transform from FramesFactory.getMOD(boolean) to provided frame.

Constructors in fr.cnes.sirius.patrius.stela.bodies that throw OrekitException

MeeusMoonStela(double inEarthRadius)
Simple constructor.

MeeusSunStela()
Deprecated. Simple constructor.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.forces

Methods in fr.cnes.sirius.patrius.stela.forces that throw OrekitException

double[][]	StelaForceModel.computePartialDerivatives(StelaEquinoctialOrbit orbit) Compute the partial derivatives for a given spacecraft state.
abstract double[]	AbstractStelaLagrangeContribution.computePerturbation(StelaEquinoctialOrbit orbit) Compute the dE/dt force derivatives for a given spacecraft state.
abstract double[]	AbstractStelaGaussContribution.computePerturbation(StelaEquinoctialOrbit orbit, OrbitNatureConverter converter) Compute the dE/dt force derivatives for a given spacecraft state.
double[]	StelaForceModel.computeShortPeriods(StelaEquinoctialOrbit orbit) Compute the short periodic variations for a given spacecraft state.
static double[][]	Squaring.computeSquaringPoints(int numPoints, StelaEquinoctialOrbit orbit, double startPoint, double endPoint) Computation of squaring points equally distributed according to true anomaly.
static StelaEquinoctialOrbit[]	Squaring.computeSquaringPointsEccentric(int numPoints, StelaEquinoctialOrbit orbit) Computation of squaring points equally distributed according to eccentric anomaly.
static double	Squaring.simpsonMean(double[] y) Simpson's rule.
static double	Squaring.simpsonMean(double[] y, double deltaEi) Simpson's rule when the integration is not done on the entire orbit, but only on one specific part.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.forces.atmospheres

Methods in fr.cnes.sirius.patrius.stela.forces.atmospheres that throw OrekitException

AtmosphereData	MSIS00Adapter.getData(AbsoluteDate date, Vector3D position, Frame frame) Get detailed atmospheric data.
double	MSIS00Adapter.getDensity(AbsoluteDate date, Vector3D position, Frame frame) Get the local density.
double	MSIS00Adapter.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
Vector3D	MSIS00Adapter.getVelocity(AbsoluteDate date, Vector3D position, Frame frame)

Uses of OrekitException in fr.cnes.sirius.patrius.stela.forces.drag

Methods in fr.cnes.sirius.patrius.stela.forces.drag that throw OrekitException

void	StelaAeroModel.addDDragAccDParam(SpacecraftState s, Parameter param, double density, Vector3D relativeVelocity, double[] dAccdParam)
void	StelaAeroModel.addDDragAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, double density, Vector3D acceleration, Vector3D relativeVelocity, boolean computeGradientPosition, boolean computeGradientVelocity)
double[][]	StelaAtmosphericDrag.computePartialDerivatives(StelaEquinoctialOrbit orbit)
double[]	StelaAtmosphericDrag.computePerturbation(StelaEquinoctialOrbit orbit, OrbitNatureConverter converter)
double[]	StelaAtmosphericDrag.computeShortPeriods(StelaEquinoctialOrbit orbit)
Vector3D	StelaAeroModel.dragAcceleration(SpacecraftState state, double density, Vector3D relativeVelocity) Return the drag acceleration in the CIRF frame.
double	StelaCd.getCd(Vector3D position) Compute the value of the Cd coefficient depending on spacecraft altitude.

Constructors in fr.cnes.sirius.patrius.stela.forces.drag that throw OrekitException

StelaAeroModel(double inMass, StelaCd inCd, double inSurface)
Constructor to be used when partial derivatives should not be computed.

StelaAeroModel(double inMass, StelaCd inCd, double inSurface, Atmosphere inAtmosphere, double atmosDX)
Constructor to be used when partial derivatives are computed using the full finite differences method.

StelaAeroModel(double inMass, StelaCd inCd, double inSurface, Atmosphere inAtmosphere, double atmosDH, GeodPosition inGeodPosition)
Constructor to be used when partial derivatives are computed using the altitude finite differences method.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.forces.gravity

Methods in fr.cnes.sirius.patrius.stela.forces.gravity that throw OrekitException

double[][]	SolidTidesAcc.computePartialDerivatives(StelaEquinoctialOrbit orbit) Compute the partial derivatives for a given spacecraft state.
double[][]	StelaTesseralAttraction.computePartialDerivatives(StelaEquinoctialOrbit orbit) Compute the partial derivatives for a given spacecraft state.
double[][]	StelaThirdBodyAttraction.computePartialDerivatives(StelaEquinoctialOrbit orbit)
double[]	SolidTidesAcc.computePerturbation(StelaEquinoctialOrbit orbit) Compute the dE/dt force derivatives for a given spacecraft state.
double[]	StelaTesseralAttraction.computePerturbation(StelaEquinoctialOrbit orbit) Compute the dE/dt force derivatives for a given spacecraft state.
double[]	StelaThirdBodyAttraction.computePerturbation(StelaEquinoctialOrbit orbit)
double[]	SolidTidesAcc.computeShortPeriods(StelaEquinoctialOrbit orbit) Compute the short periodic variations for a given spacecraft state.
double[]	StelaTesseralAttraction.computeShortPeriods(StelaEquinoctialOrbit orbit) Compute the short periodic variations for a given spacecraft state.
double[]	StelaThirdBodyAttraction.computeShortPeriods(StelaEquinoctialOrbit orbit)

Constructors in fr.cnes.sirius.patrius.stela.forces.gravity that throw OrekitException

StelaThirdBodyAttraction(CelestialBody thirdBody, int inThirdBodyDegreeMaxPerturbation, int inThirdBodyDegreeMaxShortPeriods, int inThirdBodyDegreeMaxPD)
Creates a Stela third body attraction force model.

StelaZonalAttraction(PotentialCoefficientsProvider provider, int inZonalDegreeMaxPerturbation, boolean inIsJ2SquareComputed, int inZonalDegreeMaxSP, int inZonalDegreeMaxPD, boolean inIsJ2SquareParDerComputed)
Constructor.

TesseralQuad(PotentialCoefficientsProvider provider, int coefN, int coefM, int coefP, int coefQ, Orbit orbit)
Constructor.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.forces.noninertial

Methods in fr.cnes.sirius.patrius.stela.forces.noninertial that throw OrekitException

Vector3D	NonInertialContribution.computeOmega(AbsoluteDate date, Frame frame1, Frame frame2) Compute rotation vector of frame2 with respect to frame1 expressed in frame2, which is the rotation vector from frame1 to frame2.
Vector3D	NonInertialContribution.computeOmegaDerivative(AbsoluteDate date, Frame frame1, Frame frame2, double dt) Compute rotation vector derivative from frame1 to frame2 using finite differences.
double[][]	NonInertialContribution.computePartialDerivatives(StelaEquinoctialOrbit orbit)
double[]	NonInertialContribution.computePerturbation(StelaEquinoctialOrbit orbit, OrbitNatureConverter converter)
double[]	NonInertialContribution.computeShortPeriods(StelaEquinoctialOrbit orbit)

Uses of OrekitException in fr.cnes.sirius.patrius.stela.forces.radiation

Methods in fr.cnes.sirius.patrius.stela.forces.radiation that throw OrekitException

Vector3D	SRPSquaring.computeAcceleration(StelaEquinoctialOrbit orbit, PVCoordinates satSunVector) Compute the acceleration due to the force.
protected double[]	SRPSquaring.computeInOutTrueAnom(StelaEquinoctialOrbit orbit, PVCoordinates sunPV) Computation of in and out true anomalies of the shadowed part of the orbit.
double[][]	SRPPotential.computePartialDerivatives(StelaEquinoctialOrbit orbit) Compute the partial derivatives for a given spacecraft state.
double[][]	SRPSquaring.computePartialDerivatives(StelaEquinoctialOrbit orbit) Compute the partial derivatives for a given spacecraft state.
double[][]	StelaSRPSquaring.computePartialDerivatives(StelaEquinoctialOrbit orbit) Compute the partial derivatives for a given spacecraft state.
double[]	SRPPotential.computePerturbation(StelaEquinoctialOrbit orbit) Compute the dE/dt force derivatives for a given spacecraft state.
double[]	SRPSquaring.computePerturbation(StelaEquinoctialOrbit orbit, OrbitNatureConverter converter) Compute the dE/dt force derivatives for a given spacecraft state.
double[]	StelaSRPSquaring.computePerturbation(StelaEquinoctialOrbit orbit, OrbitNatureConverter converter) Compute the dE/dt force derivatives for a given spacecraft state.
double[]	StelaSRPSquaring.computePotentialPerturbation(StelaEquinoctialOrbit orbit)
double[]	SRPSquaring.computeShortPeriods(StelaEquinoctialOrbit orbit) Compute the short periodic variations for a given spacecraft state.
double[]	StelaSRPSquaring.computeShortPeriods(StelaEquinoctialOrbit orbit) Compute the short periodic variations for a given spacecraft state.
protected double[]	SRPSquaring.computeSunBetaPhi(StelaEquinoctialOrbit orbit, PVCoordinates sunPV) Computation of Sun's right ascension (φ) and declination (β) wrt the orbit plane.

Constructors in fr.cnes.sirius.patrius.stela.forces.radiation that throw OrekitException

StelaSRPSquaring(double mass, double surface, double reflectionCoef, int quadraturePoints, CelestialBody sunBody)
Create an instance of the SRP force Stela model.

StelaSRPSquaring(double mass, double surface, double reflectionCoef, int quadraturePoints, CelestialBody sunBody, double earthRadius, double dRef, double pRef)
Create an instance of the SRP force Stela model.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.orbits

Methods in fr.cnes.sirius.patrius.stela.orbits that throw OrekitException

PVCoordinates	StelaEquinoctialOrbit.getPVCoordinates(AbsoluteDate otherDate, Frame otherFrame)
StelaEquinoctialOrbit	OrbitNatureConverter.toMean(StelaEquinoctialOrbit oscOrbit) Converts an osculating StelaEquinoctialOrbit to a mean one.
StelaEquinoctialOrbit	OrbitNatureConverter.toOsculating(StelaEquinoctialOrbit meanOrbit) Converts a mean StelaEquinoctialOrbit to an osculating one.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.propagation

Methods in fr.cnes.sirius.patrius.stela.propagation that throw OrekitException

protected SpacecraftState	StelaAbstractPropagator.acceptStep(AbsoluteDate target, double epsilon) Accept a step, triggering events and step handlers.
void	StelaAbstractPropagator.addAdditionalStateProvider(AdditionalStateProvider additionalStateProvider) Add a set of user-specified state parameters to be computed along with the orbit propagation.
SpacecraftState	StelaPartialDerivativesEquations.addInitialAdditionalState(SpacecraftState state)
SpacecraftState	StelaAdditionalEquations.addInitialAdditionalState(SpacecraftState state)
void	StelaPartialDerivativesEquations.computeDerivatives(StelaEquinoctialOrbit orbit, double[] p, double[] pDot)
void	StelaAdditionalEquations.computeDerivatives(StelaEquinoctialOrbit o, double[] p, double[] pDot) Compute the derivatives related to the additional state parameters.
SpacecraftState	StelaAbstractPropagator.getInitialState() Get the propagator initial state.
SpacecraftState	StelaBasicInterpolator.getInterpolatedState() Get the interpolated state.
PVCoordinates	StelaAbstractPropagator.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
protected abstract SpacecraftState	StelaAbstractPropagator.propagateSpacecraftState(AbsoluteDate date) Extrapolate a spacecraftState up to a specific target date.
protected SpacecraftState	StelaGTOPropagator.propagateSpacecraftState(AbsoluteDate date)
SpacecraftState	StelaAbstractPropagator.propagationManagement(SpacecraftState state, double stepSize, double dt, AbsoluteDate target) Manages the current step, method to override when user wants to deal with exceptions during the propagation.
SpacecraftState	StelaGTOPropagator.propagationManagement(SpacecraftState state, double stepSize, double dt, AbsoluteDate target)
void	StelaGTOPropagator.setInitialState(SpacecraftState initialState, double massIn, boolean isOsculatingIn) Set the initial state.
void	StelaAbstractPropagator.setOrbitFrame(Frame frame) Set propagation frame.

Constructors in fr.cnes.sirius.patrius.stela.propagation that throw OrekitException

StelaDifferentialEquations(StelaGTOPropagator inStelaPropagator)
Build a new instance of the Stela differential equations.

StelaGTOPropagator(FirstOrderIntegrator integr)
Build a StelaGTOPropagator.

StelaGTOPropagator(FirstOrderIntegrator integr, AttitudeProvider inAttitudeProviderForces, AttitudeProvider inAttitudeProviderEvents, StelaBasicInterpolator inInter, double maxShiftIn, double minStepSizeIn)
Build a StelaGTOPropagator.

StelaGTOPropagator(FirstOrderIntegrator integr, AttitudeProvider inAttitudeProvider, StelaBasicInterpolator inInter, double maxShiftIn, double minStepSizeIn)
Build a StelaGTOPropagator.

StelaGTOPropagator(FirstOrderIntegrator integr, double maxShiftIn, double minStepSizeIn)
Build a StelaGTOPropagator.

Uses of OrekitException in fr.cnes.sirius.patrius.stela.propagation.data

Methods in fr.cnes.sirius.patrius.stela.propagation.data that throw OrekitException

double[]	TimeDerivativeData.getTotalContribution() Getter for the sum of all contributions to dE'/dt (E' = mean orbital parameters).
double[][]	TimeDerivativeData.getTotalContributionSTM() Getter for the sum of all contributions to dSTM/dt (STM = state transition matrix).

Uses of OrekitException in fr.cnes.sirius.patrius.tools.force.validation

Methods in fr.cnes.sirius.patrius.tools.force.validation that throw OrekitException

PVCoordinates	BasicPVCoordinatesProvider.getPVCoordinates(AbsoluteDate date, Frame inFrame) Get the PVCoordinates of the body in the selected frame.
void	PVEphemerisLoader.loadData(InputStream input, String name)

Uses of OrekitException in fr.cnes.sirius.patrius.utils

Methods in fr.cnes.sirius.patrius.utils that throw OrekitException

PVCoordinates

AlmanacPVCoordinates.getPVCoordinates(AbsoluteDate date, Frame frame) Geometric computation of the position to a date.

Uses of OrekitException in fr.cnes.sirius.patrius.wrenches

Methods in fr.cnes.sirius.patrius.wrenches that throw OrekitException

Vector3D	SolarRadiationWrench.computeTorque(SpacecraftState s) Compute the resulting torque at the mass centre of the spacecraft in the frame of the main part.
Vector3D	GravitationalAttractionWrench.computeTorque(SpacecraftState s) Compute the resulting torque at the mass centre of the spacecraft in the frame of the main part.
Vector3D	DragWrench.computeTorque(SpacecraftState s) Compute the resulting torque at the mass centre of the spacecraft in the frame of the main part.
Vector3D	GenericWrenchModel.computeTorque(SpacecraftState s) Compute the resulting torque at the mass centre of the spacecraft in the frame of the main part.
Vector3D	MagneticWrench.computeTorque(SpacecraftState s) Compute the resulting torque at the mass centre of the spacecraft in the frame of the main part.
Vector3D	SolarRadiationWrench.computeTorque(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Vector3D	GravitationalAttractionWrench.computeTorque(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Vector3D	DragWrench.computeTorque(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Vector3D	GenericWrenchModel.computeTorque(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Vector3D	MagneticWrench.computeTorque(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	SolarRadiationWrench.computeWrench(SpacecraftState s) Compute the resulting wrench at the mass centre of the spacecraft in the frame of the main part.
Wrench	GravitationalAttractionWrench.computeWrench(SpacecraftState s) Compute the resulting wrench at the mass centre of the spacecraft in the frame of the main part.
Wrench	DragWrench.computeWrench(SpacecraftState s) Compute the resulting wrench at the mass centre of the spacecraft in the frame of the main part.
Wrench	GenericWrenchModel.computeWrench(SpacecraftState s) Compute the resulting wrench at the mass centre of the spacecraft in the frame of the main part.
Wrench	MagneticWrench.computeWrench(SpacecraftState s) Compute the resulting wrench at the mass centre of the spacecraft in the frame of the main part.
Wrench	SolarRadiationWrench.computeWrench(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	GravitationalAttractionWrench.computeWrench(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	DragWrench.computeWrench(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	GenericWrenchModel.computeWrench(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	MagneticWrench.computeWrench(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	DragWrenchSensitive.dragWrench(SpacecraftState state, double density, Vector3D relativeVelocity) Compute the torque due to radiation pressire.
Wrench	DragWrenchSensitive.dragWrench(SpacecraftState state, double density, Vector3D relativeVelocity, Vector3D origin, Frame frame) Compute the torque due to radiation pressire.
Wrench	RadiationWrenchSensitive.radiationWrench(SpacecraftState state, Vector3D flux) Compute the torque due to radiation pressire.
Wrench	RadiationWrenchSensitive.radiationWrench(SpacecraftState state, Vector3D flux, Vector3D origin, Frame frame) Compute the torque due to radiation pressire.

Uses of OrekitException in fr.cnes.sirius.validate.files

Methods in fr.cnes.sirius.validate.files that throw OrekitException

static void	ResultsFileWriter.writeLoggedEventsToVTS(String thematic, String useCase, String comment, String eventType, List<EventsLogger.LoggedEvent> results) Method to write a visualisation tool VTS MEM file.
static void	ResultsFileWriter.writeResultsToVTS(String thematic, String useCase, String orbitComment, String attitudeComment, List<SpacecraftState> results) Method to write the attitudes and positions files used by VTS visualization tool.

Uses of OrekitException in fr.cnes.sirius.validate.mocks.ephemeris

Methods in fr.cnes.sirius.validate.mocks.ephemeris that throw OrekitException

PVCoordinates	UserCelestialBody.getPVCoordinates(AbsoluteDate date, Frame frame) Computes the PV coordinates at a date using linear interpolator.
PVCoordinates	UserCelestialBody.getPVCoordinatesLagrange4(AbsoluteDate date, Frame frame) Computes the PV coordinates at a date using Lagrange 4 interpolator.

Uses of OrekitException in org.orekit.attitudes

Methods in org.orekit.attitudes that throw OrekitException

void	AttitudeLegsSequence.add(String code, AttitudeLeg leg) Adds an AttitudeLeg instance to the sequence, with conditions. The conditions are : A new instance can only be inserted at the beginning or the end of the sequence (except the first one of course).
void	Slew.compute(PVCoordinatesProvider pvProv) Compute the slew corresponding to an orbital state.
void	IsisNumericalSpinBiasSlew.compute(PVCoordinatesProvider pvProv) Compute the slew corresponding to an orbital state.
void	IsisAnalyticalSpinBiasSlew.compute(PVCoordinatesProvider pvProv) Compute the slew corresponding to an orbital state.
void	ConstantSpinSlew.compute(PVCoordinatesProvider pvProv)
void	TwoSpinBiasSlew.compute(PVCoordinatesProvider pvProv)
double	AbstractIsisSpinBiasSlew.computeDuration(PVCoordinatesProvider pvProv) Computes the slew duration.
protected double	FixedStepAttitudeEphemerisGenerator.computeStep(AbsoluteDate date, AbsoluteDateInterval ephemerisInterval)
protected abstract double	AbstractAttitudeEphemerisGenerator.computeStep(AbsoluteDate date, AbsoluteDateInterval ephemerisInterval) Computes the step used during attitude ephemeris generation.
protected double	VariableStepAttitudeEphemerisGenerator.computeStep(AbsoluteDate date, AbsoluteDateInterval ephemerisInterval) Computes the step used during the variable step ephemeris generation.
SortedSet<Attitude>	AbstractAttitudeEphemerisGenerator.generateEphemeris(AbsoluteDateInterval ephemerisInterval, Frame frame) Computes attitude ephemeris using a fixed or variable time step and choosing the interval of validity.
SortedSet<Attitude>	AbstractAttitudeEphemerisGenerator.generateEphemeris(Frame frame) Computes attitude ephemeris using a fixed or variable time step.
Attitude	Slew.getAttitude(AbsoluteDate date, Frame frame) Compute the attitude.
Attitude	ConstantSpinSlew.getAttitude(AbsoluteDate date, Frame frame)
Attitude	AbstractIsisSpinBiasSlew.getAttitude(AbsoluteDate date, Frame frame) Compute the attitude.
Attitude	TwoSpinBiasSlew.getAttitude(AbsoluteDate date, Frame frame)
Attitude	AbstractAttitudeLaw.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	AttitudeLawLeg.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	RelativeTabulatedAttitudeLaw.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	AttitudeLegLaw.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	AttitudeProvider.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	TabulatedAttitude.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	AbstractSlew.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	AttitudeLegsSequence.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	RelativeTabulatedAttitudeLeg.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	AttitudesSequence.getAttitude(Orbit orbit) Compute the attitude corresponding to an orbital state.
Attitude	ComposedAttitudeLaw.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame)
Attitude	BodyCenterPointing.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	TargetPointing.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	ConstantAttitudeLaw.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	AttitudeLawLeg.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame)
Attitude	YawCompensation.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	LofOffset.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	RelativeTabulatedAttitudeLaw.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	IsisSunPointing.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	TwoDirectionsAttitude.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the attitude corresponding to an orbital state.
Attitude	AttitudeLegLaw.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	TabulatedAttitude.getAttitude(PVCoordinatesProvider pvProvider, AbsoluteDate date, Frame inFrame) 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	AbstractSlew.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame)
Attitude	AttitudeLegsSequence.getAttitude(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Gets the attitude from the sequence. The AttitudeLeg matching the date is called to compute the attitude.
Attitude	RelativeTabulatedAttitudeLeg.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	SpinStabilized.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	GroundPointing.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.
AttitudeLeg	AttitudeLegsSequence.getAttitudeLeg(AbsoluteDate date) Gets the AttitudeLeg corresponding to the date.
AttitudeLeg	AttitudeLegsSequence.getAttitudeLeg(String code) Get the attitude leg in the sequence with the selected code.
Attitude	GroundPointingWrapper.getBaseState(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the base system state at given date, without compensation.
abstract TimeStampedAngularCoordinates	GroundPointingWrapper.getCompensation(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame, Attitude base) Compute the TimeStampedAngularCoordinates at a given time.
TimeStampedAngularCoordinates	YawCompensation.getCompensation(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame orbitFrame, Attitude base) Compute the TimeStampedAngularCoordinates at a given time.
TimeStampedAngularCoordinates	YawSteering.getCompensation(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame orbitFrame, Attitude base) Compute the TimeStampedAngularCoordinates at a given time.
double	AbstractSlew.getDuration()
AttitudeLeg	AttitudeLegsSequence.getNextAttitudeLeg(AttitudeLeg law) Gets the next attitude law after the selected AttitudeLeg law.
AttitudeLeg	AttitudeLegsSequence.getNextAttitudeLeg(String law) Gets the next attitude law after the AttitudeLeg law with the selected code.
Attitude	AttitudeLegsSequence.getOldAttitudeOnTransition(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Returns the attitude at a transition date for the old attitude law (the law previous to the valid transition law).
Rotation	DirectionTrackingOrientation.getOrientation(AbsoluteDate date, Frame frame)
Rotation	IOrientationLaw.getOrientation(AbsoluteDate date, Frame frame) Gets the rotation defining the orientation with respect to a given frame at a given date.
AttitudeLeg	AttitudeLegsSequence.getPreviousAttitudeLeg(AttitudeLeg law) Gets the previous attitude law before the selected AttitudeLeg law.
AttitudeLeg	AttitudeLegsSequence.getPreviousAttitudeLeg(String law) Gets the previous attitude law before the selected AttitudeLeg law.
Vector3D	Attitude.getRotationAcceleration() Get the satellite rotation acceleration.
Vector3D	TwoSpinBiasSlew.getSpinDerivatives(AbsoluteDate date, Frame frame) get the spin derivatives (default implementation : finite differences differentiator).
protected Vector3D	GroundPointingWrapper.getTargetPoint(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the target point in specified frame.
protected Vector3D	TargetGroundPointing.getTargetPoint(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the target point in specified frame.
protected Vector3D	BodyCenterGroundPointing.getTargetPoint(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame)
protected Vector3D	LofOffsetPointing.getTargetPoint(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the target point in specified frame.
protected Vector3D	NadirPointing.getTargetPoint(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the target point in specified frame.
protected abstract Vector3D	GroundPointing.getTargetPoint(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the target point in specified frame.
protected TimeStampedPVCoordinates	GroundPointingWrapper.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)
protected TimeStampedPVCoordinates	GroundPointing.getTargetPV(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the target point position/velocity in specified frame.
AbsoluteDateInterval	AbstractSlew.getTimeInterval()
AbsoluteDateInterval	RelativeTabulatedAttitudeLeg.getTimeInterval()
AbsoluteDateInterval	AttitudeLeg.getTimeInterval() Return the time interval of validity
Transform	OrientationTransformProvider.getTransform(AbsoluteDate date) Get the Transform corresponding to specified date.
Transform	AttitudeTransformProvider.getTransform(AbsoluteDate date) Get the Transform corresponding to specified date.
Transform	OrientationTransformProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	AttitudeTransformProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	OrientationTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the Transform corresponding to specified date.
Transform	AttitudeTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the Transform corresponding to specified date.
Transform	OrientationTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	AttitudeTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
double	YawCompensation.getYawAngle(PVCoordinatesProvider pvProv, AbsoluteDate date, Frame frame) Compute the yaw compensation angle at date.
Attitude	Attitude.interpolate(AbsoluteDate interpolationDate, Collection<Attitude> sample) Get an interpolated instance.
Attitude	Attitude.interpolate(AbsoluteDate interpolationDate, Collection<Attitude> sample, boolean computeSpinDerivatives) Interpolates attitude.
void	AbstractAttitudeLaw.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	GroundPointingWrapper.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	AttitudeLawLeg.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	RelativeTabulatedAttitudeLaw.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	AttitudeLegLaw.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	AttitudeProvider.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	RelativeTabulatedAttitudeLeg.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	LofOffsetPointing.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
void	AttitudesSequence.setSpinDerivativesComputation(boolean computeSpinDerivatives) Method to activate spin derivative computation.
TabulatedAttitude	TabulatedAttitude.setTimeInterval(AbsoluteDateInterval interval) Return a new law with the specified interval.
static Attitude	Attitude.slerp(AbsoluteDate date, Attitude attitude1, Attitude attitude2, Frame frame, boolean computeSpinDerivative) The slerp interpolation method is efficient but is less accurate than the interpolate method.
Attitude	Attitude.withReferenceFrame(Frame newReferenceFrame) Get a similar attitude with a specific reference frame.
Attitude	Attitude.withReferenceFrame(Frame newReferenceFrame, boolean spinDerivativesComputation) Get a similar attitude with a specific reference frame.

Constructors in org.orekit.attitudes that throw OrekitException

AttitudeFrame(PVCoordinatesProvider pvProvider, AttitudeLaw attitudeLaw, Frame referenceFrame)
Constructor of the dynamic spacecraft frame.

IsisSunPointing(IDirection sunDir)
Build a new instance of the class.

LofOffset(Frame inertialFrame, LOFType type)
Create a LOF-aligned attitude.

LofOffset(Frame pInertialFrame, LOFType type, RotationOrder order, double alpha1, double alpha2, double alpha3)
Creates new instance.

RelativeTabulatedAttitudeLaw(AbsoluteDate refDate, List<Pair<Double,AngularCoordinates>> angularCoordinates, Frame frame, RelativeTabulatedAttitudeLaw.AroundAttitudeType lawBefore, RelativeTabulatedAttitudeLaw.AroundAttitudeType lawAfter)
Create a RelativeTabulatedAttitudeLaw object with list of Angular Coordinates (during the interval of validity), a law before the interval and a law after the interval.

RelativeTabulatedAttitudeLaw(Frame frame, AbsoluteDate refDate, List<Pair<Double,Rotation>> orientations, RelativeTabulatedAttitudeLaw.AroundAttitudeType lawBefore, RelativeTabulatedAttitudeLaw.AroundAttitudeType lawAfter)
Create a RelativeTabulatedAttitudeLaw object with list of rotations (during the interval of validity), a law before the interval and a law after the interval.

RelativeTabulatedAttitudeLeg(AbsoluteDate referenceDate, Frame frame, List<Pair<Double,AngularCoordinates>> angularCoordinates)
Build a RelativeTabulatedAttitudeLeg with a reference date, a list of angular coordinates associated with a double representing the time elapsed since the reference date.

RelativeTabulatedAttitudeLeg(AbsoluteDate referenceDate, List<Pair<Double,AngularCoordinates>> angularCoordinates, int nbInterpolationPoints, Frame frame)
Build a RelativeTabulatedAttitudeLeg with a reference date, a list of angular coordinates associated with a double representing the time elapsed since the reference date and a number of points used for interpolation.

RelativeTabulatedAttitudeLeg(AbsoluteDate referenceDate, List<Pair<Double,Rotation>> orientations, Frame frame)
Build a RelativeTabulatedAttitudeLeg with a reference date, a list of Rotations associated with a double representing the time elapsed since the reference date.

RelativeTabulatedAttitudeLeg(AbsoluteDate referenceDate, List<Pair<Double,Rotation>> orientations, Frame frame, int nbInterpolationPoints)
Build a RelativeTabulatedAttitudeLeg with a reference date, a list of Rotations associated with a double representing the time elapsed since the reference date and a number of points used for interpolation.

SunPointing(CelestialBody body, Vector3D firstAxis, Vector3D secondAxis)
Constructor of the sun pointing attitude law.

SunPointing(Vector3D firstAxis, Vector3D secondAxis)
Constructor of the sun pointing attitude law.

SunPointing(Vector3D firstAxis, Vector3D secondAxis, CelestialBody sun)
Constructor of the sun pointing attitude law.

TabulatedAttitude(List<Attitude> inAttitudes)
Constructor with default number N of points used for interpolation.

TabulatedAttitude(List<Attitude> inAttitudes, int nbInterpolationPoints)
Constructor with number of points used for interpolation

TwoSpinBiasSlew(AttitudeProvider initialLaw, AttitudeProvider targetLaw, AbsoluteDate initialDate, double dtSCAOIn, double thetaMaxIn, double tauIn, double epsInRall, double omegaHigh, double thetaSwitch, double epsOutRall, double omegaLow, double tStab)
This class extends the AbstractSlew.

Uses of OrekitException in org.orekit.attitudes.directions

Methods in org.orekit.attitudes.directions that throw OrekitException

Line	IDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	NadirDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame)
Line	MomentumDirection.getLine(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	GlintApproximatePointingDirection.getLine(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	EarthToCelestialBodyCenterDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	GenericTargetDirection.getLine(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	ConstantVectorDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the line containing the given origin point and directed by the direction vector
Line	EarthCenterDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	CrossProductDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the cross product of directions.
Line	GroundVelocityDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame)
Line	CelestialBodyPolesAxisDirection.getLine(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	VelocityDirection.getLine(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
Line	ToCelestialBodyCenterDirection.getLine(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
PVCoordinates	ITargetDirection.getTargetPVCoordinates(AbsoluteDate date, Frame frame) Provides the target point at a given date in a given frame, represented by the associated PVCoordinates object
PVCoordinates	EarthToCelestialBodyCenterDirection.getTargetPVCoordinates(AbsoluteDate date, Frame frame) Provides the target point at a given date in a given frame, represented by the associated PVCoordinates object
PVCoordinates	GenericTargetDirection.getTargetPVCoordinates(AbsoluteDate date, Frame frame) Provides the target point at a given date in a given frame, represented by the associated PVCoordinates object
PVCoordinates	EarthCenterDirection.getTargetPVCoordinates(AbsoluteDate date, Frame frame) Provides the target point at a given date in a given frame, represented by the associated PVCoordinates object
PVCoordinates	ToCelestialBodyCenterDirection.getTargetPVCoordinates(AbsoluteDate date, Frame frame) Provides the target point at a given date in a given frame, represented by the associated PVCoordinates object
Vector3D	IDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	NadirDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame)
Vector3D	MomentumDirection.getVector(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	GlintApproximatePointingDirection.getVector(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	EarthToCelestialBodyCenterDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	GenericTargetDirection.getVector(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	ConstantVectorDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	EarthCenterDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	CrossProductDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the cross product of direction1 vector and dirction2 vector.
Vector3D	GroundVelocityDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame)
Vector3D	CelestialBodyPolesAxisDirection.getVector(PVCoordinatesProvider pvCoord, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	VelocityDirection.getVector(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.
Vector3D	ToCelestialBodyCenterDirection.getVector(PVCoordinatesProvider origin, AbsoluteDate date, Frame frame) Provides the direction vector at a given date in a given frame.

Uses of OrekitException in org.orekit.attitudes.kinematics

Methods in org.orekit.attitudes.kinematics that throw OrekitException

Vector3D	AbstractOrientationFunction.computeSpin(AbsoluteDate date) Estimate the spin at a given date from the current OrientationFunction using the quaternions formula: Ω = 2 * Q' dQ, where Q' is the conjugate of the quaternion and dQ is the derivative of the quaternion at the given date.
static Vector3D	KinematicsToolkit.computeSpin(double[] ang, double[] angd, RotationOrder order) Compute spin knowing the instantaneous quaternion and its derivative.
Vector3D	AbstractOrientationFunction.estimateRate(AbsoluteDate date, double dt) Estimate the spin at a given date from the current OrientationFunction using the AngularCoordinates.estimateRate(Rotation, Rotation, double) method.
abstract Rotation	AbstractOrientationFunction.getOrientation(AbsoluteDate date) Get the orientation at a given date.
Rotation	OrientationFunction.getOrientation(AbsoluteDate date) Get the orientation at a given date.
abstract Vector3D	AbstractVector3DFunction.getVector3D(AbsoluteDate date) Get the vector at a given date.
Vector3D	Vector3DFunction.getVector3D(AbsoluteDate date) Get the vector at a given date.

Uses of OrekitException in org.orekit.bodies

Methods in org.orekit.bodies that throw OrekitException

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.
double	GeometricBodyShape.distanceTo(Line line, Frame frame, AbsoluteDate date) Computes the distance to a line.
double	ExtendedOneAxisEllipsoid.distanceTo(Line line, Frame frame, AbsoluteDate date)
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.
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	ExtendedOneAxisEllipsoid.getIntersectionPoint(Line line, Vector3D close, Frame frame, AbsoluteDate date)
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[]	GeometricBodyShape.getIntersectionPoints(Line line, Frame frame, AbsoluteDate date) Compute the intersection points with a line.
Vector3D[]	ExtendedOneAxisEllipsoid.getIntersectionPoints(Line line, Frame frame, AbsoluteDate date)
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.
double	GeometricBodyShape.getLocalRadius(Vector3D position, Frame frame, AbsoluteDate date, PVCoordinatesProvider occultedBody) Calculate the apparent radius.
double	ExtendedOneAxisEllipsoid.getLocalRadius(Vector3D position, Frame frame, AbsoluteDate date, PVCoordinatesProvider occultedBody)
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.
PVCoordinates	MeeusMoon.getPVCoordinates(AbsoluteDate date, Frame frame)
PVCoordinates	ExtendedOneAxisEllipsoid.getPVCoordinates(AbsoluteDate date, Frame frame)
PVCoordinates	MeeusSun.getPVCoordinates(AbsoluteDate date, Frame frame)
abstract PVCoordinates	AbstractCelestialBody.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
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.
GeodeticPoint	BodyShape.transform(Vector3D point, Frame frame, AbsoluteDate date) Transform a cartesian point to a surface-relative point.
GeodeticPoint	ExtendedOneAxisEllipsoid.transform(Vector3D point, Frame frame, AbsoluteDate date)
GeodeticPoint	OneAxisEllipsoid.transform(Vector3D point, Frame frame, AbsoluteDate date) Transform a cartesian point to a surface-relative point.
Vector3D	OneAxisEllipsoid.transformAndComputeJacobian(GeodeticPoint geodeticPoint, double[][] jacobian) Transform a surface-relative point to a cartesian point and compute the jacobian of the transformation.
GeodeticPoint	OneAxisEllipsoid.transformAndComputeJacobian(Vector3D point, Frame frame, AbsoluteDate date, double[][] jacobian) Transform a cartesian point to a surface-relative point and compute the jacobian of the transformation.
static void	MeeusSun.updateTransform(AbsoluteDate date, Frame frame) Update cached transform from FramesFactory.getMOD(boolean) to provided frame.

Constructors in org.orekit.bodies that throw OrekitException

JPLEphemeridesLoader(String supportedNames, JPLEphemeridesLoader.EphemerisType generateType)
Create a loader for JPL ephemerides binary files.

MeeusMoon()
Simple constructor.

MeeusMoon(int numberOfLongitudeTerms, int numberOfLatitudeTerms, int numberOfDistanceTerms)
Simple constructor.

MeeusSun()
Simple constructor for standard Meeus model.

MeeusSun(MeeusSun.MODEL model)
Constructor to build wished Meeus model : standard model, STELA model or board model.

Uses of OrekitException in org.orekit.data

Methods in org.orekit.data that throw OrekitException

void	DataProvidersManager.addDefaultProviders() Add the default providers configuration.
boolean	ZipJarCrawler.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	DataProvider.feed(Pattern supported, DataLoader visitor) Feed a data file loader by browsing the data collection.
boolean	NetworkCrawler.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.

Constructors in org.orekit.data that throw OrekitException

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.

PoissonSeries(InputStream stream, double factor, String name)
Build a Poisson series 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.

Uses of OrekitException in org.orekit.errors

Subclasses of OrekitException in org.orekit.errors

class	FrameAncestorException This class is the base class for exception thrown by the UpdatableFrame.updateTransform(Frame, Frame, Transform, AbsoluteDate) method.
class	PropagationException This class is the base class for all specific exceptions thrown by during the propagation computation.
class	TimeStampedCacheException This class is the base class for all specific exceptions thrown by during the TimeStampedCache.

Methods in org.orekit.errors that return OrekitException

OrekitException

OrekitExceptionWrapper.getException() Get the wrapped exception.

Methods in org.orekit.errors with parameters of type OrekitException

static TimeStampedCacheException	TimeStampedCacheException.unwrap(OrekitException oe) Recover a PropagationException, possibly embedded in a OrekitException.
static PropagationException	PropagationException.unwrap(OrekitException oe) Recover a PropagationException, possibly embedded in a OrekitException.

Constructors in org.orekit.errors with parameters of type OrekitException

OrekitException(OrekitException exception)
Copy constructor.

OrekitExceptionWrapper(OrekitException wrappedException)
Simple constructor.

PropagationException(OrekitException exception)
Simple constructor.

TimeStampedCacheException(OrekitException exception)
Simple constructor.

Uses of OrekitException in org.orekit.files.general

Methods in org.orekit.files.general that throw OrekitException

OrbitFile	OrbitFileParser.parse(InputStream stream) Reads an orbit file from the given stream and returns a parsed OrbitFile.
OrbitFile	OrbitFileParser.parse(String fileName) Reads the orbit file and returns a parsed OrbitFile.

Uses of OrekitException in org.orekit.files.sp3

Methods in org.orekit.files.sp3 that throw OrekitException

SP3File	SP3Parser.parse(InputStream stream) Reads an orbit file from the given stream and returns a parsed OrbitFile.
SP3File	SP3Parser.parse(String fileName) Reads the orbit file and returns a parsed OrbitFile.

Uses of OrekitException in org.orekit.forces

Methods in org.orekit.forces that throw OrekitException

void	ForceModel.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
Vector3D	ForceModel.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.

Uses of OrekitException in org.orekit.forces.atmospheres

Methods in org.orekit.forces.atmospheres that throw OrekitException

double	DTM2000InputParameters.get24HoursKp(AbsoluteDate date) Get the last 24H mean geomagnetic index.
double	JB2006InputParameters.getAp(AbsoluteDate date) Get the Geomagnetic planetary 3-hour index Ap.
double[]	MSISE2000InputParameters.getApValues(AbsoluteDate date) Get the array containing the 7 ap values
AtmosphereData	DTM2000.getData(AbsoluteDate date, Vector3D position, Frame frame) Get detailed atmospheric data.
AtmosphereData	ExtendedAtmosphere.getData(AbsoluteDate date, Vector3D position, Frame frame) Get detailed atmospheric data.
AtmosphereData	MSISE2000.getData(AbsoluteDate date, Vector3D position, Frame frame) Get detailed atmospheric data.
double	DTM2000.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	HarrisPriester.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	MSISE2000.getDensity(AbsoluteDate date, Vector3D position, Frame frame) Get the local density.
double	US76.getDensity(AbsoluteDate date, Vector3D position, Frame frame) Get the local density for altitude in interval [0, 1E6] m
double	JB2006.getDensity(AbsoluteDate date, Vector3D position, Frame frame) Get the local density.
double	DTM2000.getDensity(double day, double alti, double lon, double lat, double hl, double f, double fbar, double akp3, double akp24) Deprecated. use DTM2000.getDensity(AbsoluteDate, Vector3D, Frame) instead
double	HarrisPriester.getDensity(double sunRAsc, double sunDecl, Vector3D satPos, double satAlt) Get the local density.
double	JB2006InputParameters.getF10(AbsoluteDate date) Get the value of the instantaneous solar flux index (1e-22*Watt/(m2*Hertz)).
double	JB2006InputParameters.getF10B(AbsoluteDate date) Get the value of the mean solar flux.
double	MSISE2000InputParameters.getInstantFlux(AbsoluteDate date) Get the value of the instantaneous solar flux.
double	DTM2000InputParameters.getInstantFlux(AbsoluteDate date) Get the value of the instantaneous solar flux.
double	MSISE2000InputParameters.getMeanFlux(AbsoluteDate date) Get the 81 day average of F10.7 flux.
double	DTM2000InputParameters.getMeanFlux(AbsoluteDate date) Get the value of the mean solar flux.
double	US76.getPress(AbsoluteDate date, Vector3D position, Frame frame) Get the local pressure for altitude in interval [0, 1E6] m
double	JB2006InputParameters.getS10(AbsoluteDate date) Get the EUV index (26-34 nm) scaled to F10.
double	JB2006InputParameters.getS10B(AbsoluteDate date) Get the EUV 81-day averaged centered index.
double	DTM2000.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	Atmosphere.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	HarrisPriester.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	SimpleExponentialAtmosphere.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	MSISE2000.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	US76.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	JB2006.getSpeedOfSound(AbsoluteDate date, Vector3D position, Frame frame) Get the local speed of sound.
double	US76.getTemp(AbsoluteDate date, Vector3D position, Frame frame) Get the local temperature for altitude in interval [0, 1E6] m
double	DTM2000InputParameters.getThreeHourlyKP(AbsoluteDate date) Get the value of the 3 hours geomagnetic index.
Vector3D	DTM2000.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the inertial velocity of atmosphere molecules.
Vector3D	Atmosphere.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the spacecraft velocity relative to the atmosphere.
Vector3D	HarrisPriester.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the inertial velocity of atmosphere molecules.
Vector3D	SimpleExponentialAtmosphere.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the spacecraft velocity relative to the atmosphere.
Vector3D	MSISE2000.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the spacecraft velocity relative to the atmosphere.
Vector3D	US76.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the spacecraft velocity relative to the atmosphere.
Vector3D	JB2006.getVelocity(AbsoluteDate date, Vector3D position, Frame frame) Get the inertial velocity of atmosphere molecules.
double	JB2006InputParameters.getXM10(AbsoluteDate date) Get the MG2 index scaled to F10.
double	JB2006InputParameters.getXM10B(AbsoluteDate date) Get the MG2 81-day average centered index.

Constructors in org.orekit.forces.atmospheres that throw OrekitException

DTM2000(DTM2000InputParameters parameters, PVCoordinatesProvider sun, BodyShape earth)
Simple constructor for independent computation.

Uses of OrekitException in org.orekit.forces.atmospheres.solarActivity

Methods in org.orekit.forces.atmospheres.solarActivity that throw OrekitException

static void	SolarActivityDataFactory.addDefaultSolarActivityDataReaders() Add the default READERS for solar activity
double	ExtendedSolarActivityWrapper.getAp(AbsoluteDate date) Get Ap value at given user date
double	SolarActivityDataProvider.getAp(AbsoluteDate date) Get Ap value at given user date
SortedMap<AbsoluteDate,Double[]>	ExtendedSolarActivityWrapper.getApKpValues(AbsoluteDate date1, AbsoluteDate date2) Get ap / kp values between the given dates
SortedMap<AbsoluteDate,Double[]>	SolarActivityDataProvider.getApKpValues(AbsoluteDate date1, AbsoluteDate date2) Get ap / kp values between the given dates
static double	SolarActivityToolbox.getAverageFlux(AbsoluteDate date1, AbsoluteDate date2, SolarActivityDataProvider data) Compute mean flux between given dates.
double	ExtendedSolarActivityWrapper.getInstantFluxValue(AbsoluteDate date) Get instant flux values at the given dates (possibly interpolated)
double	ConstantSolarActivity.getInstantFluxValue(AbsoluteDate date) Get instant flux values at the given dates (possibly interpolated)
double	SolarActivityDataProvider.getInstantFluxValue(AbsoluteDate date) Get instant flux values at the given dates (possibly interpolated)
SortedMap<AbsoluteDate,Double>	ExtendedSolarActivityWrapper.getInstantFluxValues(AbsoluteDate date1, AbsoluteDate date2) Get raw instant flux values between the given dates
SortedMap<AbsoluteDate,Double>	SolarActivityDataProvider.getInstantFluxValues(AbsoluteDate date1, AbsoluteDate date2) Get raw instant flux values between the given dates
double	ExtendedSolarActivityWrapper.getKp(AbsoluteDate date) Get Kp value at given user date
double	SolarActivityDataProvider.getKp(AbsoluteDate date) Get Kp value at given user date
static double	SolarActivityToolbox.getMeanAp(AbsoluteDate minDate, AbsoluteDate maxDate, SolarActivityDataProvider data) Compute mean flux between given dates (rectangular rule)
static double	SolarActivityToolbox.getMeanFlux(AbsoluteDate date1, AbsoluteDate date2, SolarActivityDataProvider data) Compute mean flux between given dates using trapezoidal rule
static SolarActivityDataProvider	SolarActivityDataFactory.getSolarActivityDataProvider() Get the solar activity provider from the first supported file.
abstract void	SolarActivityDataReader.loadData(InputStream input, String name) Load data from a stream.
void	ACSOLFormatReader.loadData(InputStream input, String name) Load data from a stream.
void	NOAAFormatReader.loadData(InputStream input, String name) Load data from a stream.

Constructors in org.orekit.forces.atmospheres.solarActivity that throw OrekitException

ACSOLFormatReader(String supportedNames)
Constructor.

NOAAFormatReader(String supportedNames)
Constructor.

Uses of OrekitException in org.orekit.forces.atmospheres.solarActivity.specialized

Methods in org.orekit.forces.atmospheres.solarActivity.specialized that throw OrekitException

double	DTM2000SolarData.get24HoursKp(AbsoluteDate date) Get the last 24H mean geomagnetic index.
double	MarshallSolarActivityFutureEstimation.get24HoursKp(AbsoluteDate date) The Kp index is derived from the Ap index.
abstract double[]	AbstractMSISE2000SolarData.getApValues(AbsoluteDate date) Get the array containing the 7 ap values
double[]	ContinuousMSISE2000SolarData.getApValues(AbsoluteDate date) Get the array containing the 7 ap values
double[]	ClassicalMSISE2000SolarData.getApValues(AbsoluteDate date) Get the array containing the 7 ap values
DateComponents	MarshallSolarActivityFutureEstimation.getFileDate(AbsoluteDate date) Get the date of the file from which data at the specified date comes from.
double	AbstractMSISE2000SolarData.getInstantFlux(AbsoluteDate date) Get the value of the instantaneous solar flux.
double	DTM2000SolarData.getInstantFlux(AbsoluteDate date) Get the value of the instantaneous solar flux.
double	MarshallSolarActivityFutureEstimation.getInstantFlux(AbsoluteDate date) Get the value of the instantaneous solar flux.
double	ContinuousMSISE2000SolarData.getInstantFlux(AbsoluteDate date) Get the value of the instantaneous solar flux.
double	AbstractMSISE2000SolarData.getMeanFlux(AbsoluteDate date) Get the 81 day average of F10.7 flux.
double	DTM2000SolarData.getMeanFlux(AbsoluteDate date) Get the value of the mean solar flux.
double	MarshallSolarActivityFutureEstimation.getMeanFlux(AbsoluteDate date) Get the value of the mean solar flux.
double	ContinuousMSISE2000SolarData.getMeanFlux(AbsoluteDate date) Get the 81 day average of F10.7 flux.
double	DTM2000SolarData.getThreeHourlyKP(AbsoluteDate date) Get the value of the 3 hours geomagnetic index.
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.

Uses of OrekitException in org.orekit.forces.drag

Methods in org.orekit.forces.drag that throw OrekitException

void	DragForce.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the drag to the perturbing acceleration.
void	DragForce.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	DragForce.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	DragSensitive.addDDragAccDParam(SpacecraftState s, Parameter param, double density, Vector3D relativeVelocity, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters (the ballistic coefficient).
void	DragSensitive.addDDragAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, double density, Vector3D acceleration, Vector3D relativeVelocity, boolean computeGradientPosition, boolean computeGradientVelocity) Compute acceleration derivatives with respect to state parameters (position and velocity).
static Vector3D	DragForce.computeAcceleration(PVCoordinates pv, Frame frame, Atmosphere atm, AbsoluteDate date, double kD, double mass) Method to compute the acceleration.
Vector3D	DragForce.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	DragSensitive.dragAcceleration(SpacecraftState state, double density, Vector3D relativeVelocity) Compute the acceleration due to drag and the lift.

Uses of OrekitException in org.orekit.forces.gravity

Methods in org.orekit.forces.gravity that throw OrekitException

void	BalminoAttractionModel.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	DrozinerAttractionModel.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	ThirdBodyAttraction.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	CunninghamAttractionModel.addContribution(SpacecraftState s, TimeDerivativesEquations 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.
void	BalminoAttractionModel.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	ThirdBodyAttraction.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	CunninghamAttractionModel.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	NewtonianAttraction.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	BalminoAttractionModel.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	ThirdBodyAttraction.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	CunninghamAttractionModel.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	NewtonianAttraction.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
Vector3D	DrozinerAttractionModel.computeAcceleration(PVCoordinates pv, AbsoluteDate date) Method to compute the acceleration.
Vector3D	ThirdBodyAttraction.computeAcceleration(PVCoordinates pv, Frame frame, AbsoluteDate date) Method to compute the acceleration.
Vector3D	NewtonianAttraction.computeAcceleration(PVCoordinates pv, Frame frame, AbsoluteDate date) Method to compute the acceleration.
Vector3D	BalminoAttractionModel.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	DrozinerAttractionModel.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	ThirdBodyAttraction.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	CunninghamAttractionModel.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	NewtonianAttraction.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
static Vector3D	GravityToolbox.computeDrozinerAcceleration(PVCoordinates pv, Frame frame, double[][] coefficientsC, double[][] coefficientsS, double muc, double eqRadius, double threshold, int degree, int order) Method to compute the acceleration, from Droziner algorithm (see DrozinerAttractionModel).
static ForceModel	EarthGravitationalModelFactory.getBalmino(EarthGravitationalModelFactory.GravityFieldNames potentialFileName, int n, int m) Create an instance of a central body attraction with normalized coefficients and Helmholtz Polynomials (Balmino model).
static ForceModel	EarthGravitationalModelFactory.getCunningham(EarthGravitationalModelFactory.GravityFieldNames potentialFileName, int n, int m) Create an instance of the gravitational field of a celestial body using Cunningham model.
static ForceModel	EarthGravitationalModelFactory.getDroziner(EarthGravitationalModelFactory.GravityFieldNames potentialFileName, int n, int m) Create an instance of the gravitational field of a celestial body using Droziner model.
static ForceModel	EarthGravitationalModelFactory.getGravitationalModel(EarthGravitationalModelFactory.GravityFieldNames potentialFileName, int n, int m) Create an default instance of a gravitational field of a celestial body using Balmino model.

Uses of OrekitException in org.orekit.forces.gravity.potential

Methods in org.orekit.forces.gravity.potential that throw OrekitException

double[][]	PotentialCoefficientsReader.getC(int n, int m, boolean normalized) Get the tesseral-sectorial and zonal coefficients.
double[][]	PotentialCoefficientsProvider.getC(int n, int m, boolean normalized) Get the tesseral-sectorial and zonal coefficients.
double[]	PotentialCoefficientsReader.getJ(boolean normalized, int n) Get the zonal coefficients.
double[]	PotentialCoefficientsProvider.getJ(boolean normalized, int n) Get the zonal coefficients.
static PotentialCoefficientsProvider	GravityFieldFactory.getPotentialProvider() Get the gravity field coefficients provider from the first supported file.
double[][]	PotentialCoefficientsReader.getS(int n, int m, boolean normalized) Get tesseral-sectorial coefficients.
double[][]	PotentialCoefficientsProvider.getS(int n, int m, boolean normalized) Get tesseral-sectorial coefficients.
void	GRGSFormatReader.loadData(InputStream input, String name) Load data from a stream.
abstract void	PotentialCoefficientsReader.loadData(InputStream input, String name) Load data from a stream.
void	SHMFormatReader.loadData(InputStream input, String name) Load data from a stream.
void	EGMFormatReader.loadData(InputStream input, String name) Load data from a stream.
void	ICGEMFormatReader.loadData(InputStream input, String name) Load data from a stream.

Uses of OrekitException in org.orekit.forces.gravity.tides

Methods in org.orekit.forces.gravity.tides that throw OrekitException

void	AbstractTides.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	AbstractTides.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	AbstractTides.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
Vector3D	AbstractTides.computeAcceleration(PVCoordinates pv, Frame frame, AbsoluteDate date) Method to compute the acceleration, from Balmino algorithm (see BalminoAttractionModel class).
Vector3D	AbstractTides.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
static double[][]	TidesToolbox.computeFundamentalArguments(AbsoluteDate date, TidesStandards.TidesStandard standard) Method to compute the Doodson fundamental arguments.
double[][]	OceanTides.getDenormalizedCCoefs(AbsoluteDate date) Get denormalized C coefficients table
double[][]	OceanTides.getDenormalizedSCoefs(AbsoluteDate date) Get denormalized S coefficients table
double[][]	OceanTides.getNormalizedCCoefs(AbsoluteDate date) Get normalized C coefficients table
double[][]	OceanTides.getNormalizedSCoefs(AbsoluteDate date) Get normalized S coefficients table
abstract void	AbstractTides.updateCoefficientsCandS(AbsoluteDate date) Update the C and the S coefficients for acceleration computation.
void	PotentialTimeVariations.updateCoefficientsCandS(AbsoluteDate date) Update the C and the S coefficients for acceleration computation.
void	OceanTides.updateCoefficientsCandS(AbsoluteDate date) Update the C and the S coefficients for acceleration computation.
void	TerrestrialTides.updateCoefficientsCandS(AbsoluteDate date) Update the C and the S coefficients for acceleration computation.
abstract void	AbstractTides.updateCoefficientsCandSPD(AbsoluteDate date) Update the C and the S coefficients for partial derivatives computation.
void	PotentialTimeVariations.updateCoefficientsCandSPD(AbsoluteDate date) Update the C and the S coefficients for partial derivatives computation.
void	OceanTides.updateCoefficientsCandSPD(AbsoluteDate date) Update the C and the S coefficients for partial derivatives computation.
void	TerrestrialTides.updateCoefficientsCandSPD(AbsoluteDate date) Update the C and the S coefficients for partial derivatives computation.

Constructors in org.orekit.forces.gravity.tides that throw OrekitException

TerrestrialTides(Frame centralBodyFrame, double equatorialRadius, double mu)
Creates a new instance.

TerrestrialTides(Frame centralBodyFrame, double equatorialRadius, double mu, boolean computePD)
Creates a new instance.

TerrestrialTides(Frame centralBodyFrame, double equatorialRadius, double mu, List<CelestialBody> bodies, boolean thirdBodyAttDegree3, boolean frequencyCorr, boolean ellipticityCorr, ITerrestrialTidesDataProvider terrestrialData)
Creates a new instance.

TerrestrialTides(Frame centralBodyFrame, double equatorialRadius, double mu, List<CelestialBody> bodies, boolean thirdBodyAttDegree3, boolean frequencyCorr, boolean ellipticityCorr, ITerrestrialTidesDataProvider terrestrialData, boolean computePD)
Creates a new instance.

TerrestrialTides(Frame centralBodyFrame, Parameter equatorialRadius, Parameter mu)
Creates a new instance.

TerrestrialTides(Frame centralBodyFrame, Parameter equatorialRadius, Parameter mu, boolean computePD)
Creates a new instance.

TerrestrialTides(Frame centralBodyFrame, Parameter equatorialRadius, Parameter mu, List<CelestialBody> bodies, boolean thirdBodyAttDegree3, boolean frequencyCorr, boolean ellipticityCorr, ITerrestrialTidesDataProvider terrestrialData)
Creates a new instance using Parameter.

TerrestrialTides(Frame centralBodyFrame, Parameter equatorialRadius, Parameter mu, List<CelestialBody> bodies, boolean thirdBodyAttDegree3, boolean frequencyCorr, boolean ellipticityCorr, ITerrestrialTidesDataProvider terrestrialData, boolean computePD)
Creates a new instance using Parameter.

TerrestrialTidesDataProvider()
Simple constructor.

TerrestrialTidesDataProvider(TidesStandards.TidesStandard tideStandard)
Simple constructor.

Uses of OrekitException in org.orekit.forces.gravity.tides.coefficients

Methods in org.orekit.forces.gravity.tides.coefficients that throw OrekitException

static OceanTidesCoefficientsProvider	OceanTidesCoefficientsFactory.getCoefficientsProvider() Get the ocean tides coefficients provider from the first supported file.
void	FES2004FormatReader.loadData(InputStream input, String name) Load data from a stream.
abstract void	OceanTidesCoefficientsReader.loadData(InputStream input, String name) Load data from a stream.

Uses of OrekitException in org.orekit.forces.gravity.variations

Methods in org.orekit.forces.gravity.variations that throw OrekitException

void	VariablePotentialAttractionModel.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	VariablePotentialAttractionModel.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	VariablePotentialAttractionModel.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel)
Vector3D	VariablePotentialAttractionModel.computeAcceleration(AbsoluteDate date, PVCoordinates pv) Compute acceleration in rotating frame
Vector3D	VariablePotentialAttractionModel.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
void	VariablePotentialAttractionModel.updateCoefficientsCandS(AbsoluteDate date) Update the C and the S coefficients for acceleration computation.
void	VariablePotentialAttractionModel.updateCoefficientsCandSPD(AbsoluteDate date) Update the C and the S coefficients for partial derivatives computation.

Constructors in org.orekit.forces.gravity.variations that throw OrekitException

VariablePotentialAttractionModel(Frame centralBodyFrame, VariablePotentialCoefficientsProvider provider, int degree, int order)
Variable gravity field force model constructor (static part only).

VariablePotentialAttractionModel(Frame centralBodyFrame, VariablePotentialCoefficientsProvider provider, int degree, int order, int degreePD, int orderPD)
Variable gravity field force model constructor (static part only).

VariablePotentialAttractionModel(Frame centralBodyFrame, VariablePotentialCoefficientsProvider provider, int degree, int order, int degreeOptional, int orderOptional, boolean computeOptionalOnce)
Variable gravity field force model constructor.

VariablePotentialAttractionModel(Frame centralBodyFrame, VariablePotentialCoefficientsProvider provider, int degree, int order, int degreePD, int orderPD, int degreeOptional, int orderOptional, int degreeOptionalPD, int orderOptionalPD, boolean computeOptionalOnce)
Variable gravity field force model constructor.

Uses of OrekitException in org.orekit.forces.gravity.variations.coefficients

Methods in org.orekit.forces.gravity.variations.coefficients that throw OrekitException

static VariablePotentialCoefficientsProvider	VariableGravityFieldFactory.getVariablePotentialProvider() Get the variable gravity field coefficients provider from the first supported file.
void	GRGSRL02FormatReader.loadData(InputStream input, String name) Load data from a stream.
protected void	VariablePotentialCoefficientsReader.setYear(int fileYear) Set file year

Uses of OrekitException in org.orekit.forces.maneuvers

Methods in org.orekit.forces.maneuvers that throw OrekitException

void	ConstantThrustManeuver.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	ConstantThrustError.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the constant thrust error model to the perturbing acceleration.
void	VariableThrustManeuver.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	ConstantThrustManeuver.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	ConstantThrustError.addDAccDParam(SpacecraftState state, Parameter param, double[] dAccdParam)
void	VariableThrustManeuver.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	ConstantThrustManeuver.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	ConstantThrustError.addDAccDState(SpacecraftState state, double[][] dAccdPos, double[][] dAccdVel)
void	VariableThrustManeuver.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
SpacecraftState	SmallManeuverAnalyticalModel.apply(SpacecraftState state1) Compute the effect of the maneuver on a spacecraft state.
Vector3D	ConstantThrustManeuver.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	ConstantThrustError.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	VariableThrustManeuver.computeAcceleration(SpacecraftState s)
EventDetector.Action	ImpulseManeuver.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
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.
SpacecraftState	ImpulseManeuver.resetState(SpacecraftState oldState) Reset the state (including additional states) prior to continue propagation.

Constructors in org.orekit.forces.maneuvers that throw OrekitException

SmallManeuverAnalyticalModel(SpacecraftState state0, Frame frame, Vector3D dV, double isp, String partName)
Build a maneuver defined in user-specified frame.

SmallManeuverAnalyticalModel(SpacecraftState state0, Vector3D dV, double isp, String partName)
Build a maneuver defined in spacecraft frame.

Uses of OrekitException in org.orekit.forces.radiation

Methods in org.orekit.forces.radiation that throw OrekitException

void	RediffusedRadiationPressure.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	SolarRadiationPressure.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	RediffusedRadiationPressure.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	SolarRadiationPressure.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	RediffusedRadiationSensitive.addDAccDParamRediffusedRadiativePressure(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives.
void	RediffusedRadiationPressure.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	SolarRadiationPressure.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	RediffusedRadiationSensitive.addDAccDStateRediffusedRadiativePressure(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives.
void	RadiationSensitive.addDSRPAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam, Vector3D satSunVector) Compute acceleration derivatives with respect to additional parameters.
void	RadiationSensitive.addDSRPAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel, Vector3D satSunVector) Compute acceleration derivatives with respect to state parameters.
Vector3D	RediffusedRadiationPressure.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	SolarRadiationPressure.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
double	SolarRadiationPressure.getLightningRatio(Vector3D position, Frame frame, AbsoluteDate date) Get the lightning ratio ([0-1]).
Vector3D	RadiationSensitive.radiationPressureAcceleration(SpacecraftState state, Vector3D flux) Compute the acceleration due to radiation pressure.
Vector3D	RediffusedRadiationSensitive.rediffusedRadiationPressureAcceleration(SpacecraftState state, ElementaryFlux flux) rediffused radiative pressure acceleration

Constructors in org.orekit.forces.radiation that throw OrekitException

RediffusedFlux(int nCorona, int nMeridian, Frame bodyFrame, CelestialBody sunProvider, PVCoordinatesProvider satProvider, AbsoluteDate d, IEmissivityModel model)
Default constructor of rediffused flux.

RediffusedFlux(int nCorona, int nMeridian, Frame bodyFrame, CelestialBody sun, PVCoordinatesProvider satProvider, AbsoluteDate dDate, IEmissivityModel model, boolean inIr, boolean inAlbedo)
Generic constructor of rediffused flux.

RediffusedRadiationPressure(CelestialBody inSun, Frame inBodyFrame, int inCorona, int inMeridian, IEmissivityModel inEmissivityModel, RediffusedRadiationSensitive inModel)
Constructor.

RediffusedRadiationPressure(CelestialBody inSun, Frame inBodyFrame, int inCorona, int inMeridian, IEmissivityModel inEmissivityModel, RediffusedRadiationSensitive inModel, boolean computePD)
Constructor.

Uses of OrekitException in org.orekit.forces.relativistic

Methods in org.orekit.forces.relativistic that throw OrekitException

void	SchwarzschildRelativisticEffect.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	LenseThirringRelativisticEffect.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	CoriolisRelativisticEffect.addContribution(SpacecraftState s, TimeDerivativesEquations adder) Compute the contribution of the force model to the perturbing acceleration.
void	SchwarzschildRelativisticEffect.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	LenseThirringRelativisticEffect.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	CoriolisRelativisticEffect.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	SchwarzschildRelativisticEffect.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	LenseThirringRelativisticEffect.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
void	CoriolisRelativisticEffect.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.
Vector3D	SchwarzschildRelativisticEffect.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	LenseThirringRelativisticEffect.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.
Vector3D	CoriolisRelativisticEffect.computeAcceleration(SpacecraftState s) Compute the acceleration due to the force.

Uses of OrekitException in org.orekit.frames

Methods in org.orekit.frames that throw OrekitException

GeodeticPoint	TopocentricFrame.computeLimitVisibilityPoint(double radius, double azimuth, double elevation) Compute the limit visibility point for a satellite in a given direction.
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.getAzimuthRate(PVCoordinates extPV, Frame frame, AbsoluteDate date) Get the azimuth rate of a point.
double	TopocentricFrame.getElevation(Vector3D extPoint, Frame frame, AbsoluteDate date) Get the elevation of a point with regards to the local point.
double	TopocentricFrame.getElevationRate(PVCoordinates extPV, Frame frame, AbsoluteDate date) Get the elevation rate of a point.
static FactoryManagedFrame	FramesFactory.getEODFrame(boolean applyEOPCorr) This class implements the EOD frame (mean ecliptic and equinox of the epoch).
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 Frame	FramesFactory.getH0MinusN(String name, AbsoluteDate h0MinusN, double longitude) Get the "H0 - n" reference frame.
static Frame	FramesFactory.getH0MinusN(String name, AbsoluteDate h0, double n, double longitude) Get the "H0 - n" reference frame.
static Frame	FramesFactory.getICRF() Get the unique ICRF frame.
static FactoryManagedFrame	FramesFactory.getITRF() Get the ITRF reference frame.
static FactoryManagedFrame	FramesFactory.getITRFEquinox() Get the equinox-based ITRF reference frame.
static FactoryManagedFrame	FramesFactory.getMOD(boolean applyEOPCorr) Get the MOD reference frame.
PVCoordinates	SpacecraftFrame.getPVCoordinates(AbsoluteDate date, Frame frame) Deprecated. Get the PVCoordinates of the spacecraft frame origin in the selected frame.
PVCoordinates	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() Get the TIRF reference frame.
static FactoryManagedFrame	FramesFactory.getTOD(boolean applyEOPCorr) Get the TOD reference frame.
RealMatrix	Frame.getTransformJacobian(Frame to, AbsoluteDate date) Compute the Jacobian from current frame to target frame at provided date.
Transform	Frame.getTransformTo(Frame destination, AbsoluteDate date) Get the transform from the instance to another frame.
Transform	Frame.getTransformTo(Frame destination, AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from the instance to another frame.
Transform	Frame.getTransformTo(Frame destination, AbsoluteDate date, FramesConfiguration config) Get the transform from the instance to another frame.
Transform	Frame.getTransformTo(Frame destination, AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from the instance to another frame.
static FactoryManagedFrame	FramesFactory.getVeis1950() Get the VEIS 1950 reference frame.
double	TopocentricFrame.getXangleCardan(Vector3D extPoint, Frame frame, AbsoluteDate date) Get the Cardan x angle of a point.
double	TopocentricFrame.getXangleCardanRate(PVCoordinates extPV, Frame frame, AbsoluteDate date) Get the Cardan x angle rate.
double	TopocentricFrame.getYangleCardan(Vector3D extPoint, Frame frame, AbsoluteDate date) Get the Cardan y angle of a point with regards to the projection point on the plane defined by the zenith and the west axis.
double	TopocentricFrame.getYangleCardanRate(PVCoordinates extPV, Frame frame, AbsoluteDate date) Get the Cardan y angle rate.
GeodeticPoint	TopocentricFrame.pointAtDistance(double azimuth, double elevation, double distance) Compute the point observed from the station at some specified distance.
CardanMountPosition	TopocentricFrame.transformFromPositionToCardan(Vector3D extPoint, Frame frame, AbsoluteDate date) Transform a Cartesian position coordinates into Cardan mounting in this local topocentric frame.
TopocentricPosition	TopocentricFrame.transformFromPositionToTopocentric(Vector3D extPoint, Frame frame, AbsoluteDate date) Transform a Cartesian position coordinates into topocentric coordinates in this local topocentric frame.
CardanMountPV	TopocentricFrame.transformFromPVToCardan(PVCoordinates extPV, Frame frame, AbsoluteDate date) Transform a Cartesian position coordinates into Cardan mounting in this local topocentric frame.
TopocentricPV	TopocentricFrame.transformFromPVToTopocentric(PVCoordinates extPV, Frame frame, AbsoluteDate date) Transform a Cartesian position and velocity coordinates into topocentric coordinates in this local topocentric frame.
void	UpdatableFrame.updateTransform(Frame f1, Frame f2, Transform f1Tof2, AbsoluteDate date) Update the transform from parent frame implicitly according to two other frames.

Uses of OrekitException in org.orekit.frames.configuration

Methods in org.orekit.frames.configuration that throw OrekitException

double[]	FramesConfigurationImplementation.getPolarMotion(AbsoluteDate date) Compute corrected polar motion.
double[]	FramesConfiguration.getPolarMotion(AbsoluteDate date) Compute corrected polar motion.
PoleCorrection	PolarMotion.getPoleCorrection(AbsoluteDate date) Compute pole correction.

Uses of OrekitException in org.orekit.frames.configuration.eop

Methods in org.orekit.frames.configuration.eop that throw OrekitException

void	AbstractEOPHistory.checkEOPContinuity(double maxGap) Check Earth orientation parameters continuity.
void	RapidDataAndPredictionColumnsLoader.fillHistory(EOP1980History history) Load celestial body.
void	RapidDataAndPredictionXMLLoader.fillHistory(EOP1980History history) Load celestial body.
void	BulletinBFilesLoader.fillHistory(EOP1980History history) Load celestial body.
void	EOP08C04FilesLoader.fillHistory(EOP1980History history) Load celestial body.
void	EOP1980HistoryLoader.fillHistory(EOP1980History history) Load celestial body.
void	EOP05C04FilesLoader.fillHistory(EOP1980History history) Load celestial body.
void	NoEOP1980HistoryLoader.fillHistory(EOP1980History history) History with zero orientation.
static void	EOP08C04FilesLoader.fillHistory(EOP1980History history, InputStream istream) Fills the history object directy with data from the InputStream, bypassing the Orekit data loaders.
static void	EOP05C04FilesLoader.fillHistory(EOP1980History history, InputStream istream) Fills the history object directy with data from the InputStream, bypassing the Orekit data loaders.
void	RapidDataAndPredictionColumnsLoader.fillHistory(EOP2000History history) Load celestial body.
void	RapidDataAndPredictionXMLLoader.fillHistory(EOP2000History history) Load celestial body.
void	EOP2000HistoryLoader.fillHistory(EOP2000History history) Load celestial body.
void	BulletinBFilesLoader.fillHistory(EOP2000History history) Load celestial body.
void	EOP08C04FilesLoader.fillHistory(EOP2000History history) Load celestial body.
void	EOP05C04FilesLoader.fillHistory(EOP2000History history) Load celestial body.
static void	EOP08C04FilesLoader.fillHistory(EOP2000History history, InputStream istream) Fills the history object directy with data from the InputStream, bypassing the Orekit data loaders.
static void	EOP05C04FilesLoader.fillHistory(EOP2000History history, InputStream istream) Fills the history object directy with data from the InputStream, bypassing the Orekit data loaders.
static EOP1980History	EOPHistoryFactory.getEOP1980History() Get Earth Orientation Parameters history (IAU1980) data.
static EOP1980History	EOPHistoryFactory.getEOP1980History(EOPInterpolators interpMethod) Get Earth Orientation Parameters history (IAU1980) data.
static EOP2000History	EOPHistoryFactory.getEOP2000History() Get Earth Orientation Parameters history (IAU2000) data.
static EOP2000History	EOPHistoryFactory.getEOP2000History(EOPInterpolators interpMethod) Get Earth Orientation Parameters history (IAU2000) data.
static EOP2000History	EOPHistoryFactory.getEOP2000History(EOPInterpolators interpMethod, EOP2000HistoryLoader loader) Get Earth Orientation Parameters history (IAU2000) data using a specific loader.
static EOP2000HistoryConstantOutsideInterval	EOPHistoryFactory.getEOP2000HistoryConstant() Get Earth Orientation Parameters history (IAU2000) data.
static EOP2000HistoryConstantOutsideInterval	EOPHistoryFactory.getEOP2000HistoryConstant(EOPInterpolators interpMethod) Get Earth Orientation Parameters history (IAU2000) data.
static EOP2000HistoryConstantOutsideInterval	EOPHistoryFactory.getEOP2000HistoryConstant(EOPInterpolators interpMethod, EOP2000HistoryLoader loader) Get Earth Orientation Parameters history (IAU2000) data using a specific loader.
void	RapidDataAndPredictionColumnsLoader.loadData(InputStream input, String name) Load data from a stream.
void	RapidDataAndPredictionXMLLoader.loadData(InputStream input, String name) Load data from a stream.
void	BulletinBFilesLoader.loadData(InputStream input, String name) Load data from a stream.
void	EOP08C04FilesLoader.loadData(InputStream input, String name) Load data from a stream.
void	EOP05C04FilesLoader.loadData(InputStream input, String name) Load data from a stream.
void	NoEOP1980HistoryLoader.loadData(InputStream input, String name) Load data from a stream.

Constructors in org.orekit.frames.configuration.eop that throw OrekitException

EOP1980Entry(AbsoluteDate adate, double dt, double lod, double x, double y, double ddPsi, double ddEps)
Constructor with an AbsoluteDate parameter.

EOP1980Entry(AbsoluteDate adate, double dt, double lod, double x, double y, double ddPsi, double ddEps, EOPEntry.DtType type)
Constructor with an AbsoluteDate parameter.

EOP1980Entry(DateComponents datec, double dt, double lod, double x, double y, double ddPsi, double ddEps)
Constructor with DateComponents parameter.

EOP1980Entry(DateComponents datec, double dt, double lod, double x, double y, double ddPsi, double ddEps, EOPEntry.DtType type)
Constructor with DateComponents parameter.

EOP1980Entry(int mjd, double dt, double lod, double x, double y, double ddPsi, double ddEps)
Simple constructor.

EOP1980Entry(int mjd, double dt, double lod, double x, double y, double ddPsi, double ddEps, EOPEntry.DtType type)
Simple constructor.

EOP2000Entry(AbsoluteDate adate, double dt, double lod, double x, double y, double dx, double dy)
Constructor with an AbsoluteDate parameter.

EOP2000Entry(AbsoluteDate adate, double dt, double lod, double x, double y, double dx, double dy, EOPEntry.DtType type)
Constructor with an AbsoluteDate parameter.

EOP2000Entry(DateComponents datec, double dt, double lod, double x, double y, double dx, double dy)
Constructor with DateComponents parameter.

EOP2000Entry(DateComponents datec, double dt, double lod, double x, double y, double dx, double dy, EOPEntry.DtType type)
Constructor with DateComponents parameter.

EOP2000Entry(int mjd, double dt, double lod, double x, double y, double dx, double dy)
Simple constructor.

EOP2000Entry(int mjd, double dt, double lod, double x, double y, double dx, double dy, EOPEntry.DtType type)
Simple constructor.

EOPEntry(AbsoluteDate adate, double dt, double lod, double x, double y, double dx, double dy)
Constructor with an AbsoluteDate parameter.

EOPEntry(AbsoluteDate adate, double dt, double lod, double x, double y, double dx, double dy, EOPEntry.DtType type)
Constructor with an AbsoluteDate parameter.

EOPEntry(DateComponents datec, double dt, double lod, double x, double y, double dx, double dy)
Constructor with DateComponents parameter.

EOPEntry(DateComponents datec, double dt, double lod, double x, double y, double dx, double dy, EOPEntry.DtType type)
Constructor with DateComponents parameter.

EOPEntry(int mjd, double dt, double lod, double x, double y, double dx, double dy)
Simple constructor.

EOPEntry(int mjd, double dt, double lod, double x, double y, double dx, double dy, EOPEntry.DtType type)
Simple constructor.

Uses of OrekitException in org.orekit.frames.configuration.libration

Methods in org.orekit.frames.configuration.libration that throw OrekitException

PoleCorrection	LibrationCorrectionPerThread.getPoleCorrection(AbsoluteDate date) Compute the pole corrections at a given date.
PoleCorrection	LibrationCorrectionModel.getPoleCorrection(AbsoluteDate t) Compute the pole corrections at a given date.
PoleCorrection	IERS2010LibrationCorrection.getPoleCorrection(AbsoluteDate date) This method provides the diurnal lunisolar effect on polar motion in time domain.

Uses of OrekitException in org.orekit.frames.transformations

Methods in org.orekit.frames.transformations that throw OrekitException

static double	TIRFProvider.getEarthRotationAngle(AbsoluteDate date) Get the Earth Rotation Angle at the current date.
static double	TODProvider.getEquationOfEquinoxes(AbsoluteDate date) Get the Equation of the Equinoxes at the current date.
double	GTODProvider.getGAST(AbsoluteDate date) Get the Greenwich apparent sidereal time, in radians.
static double	GTODProvider.getGMST(AbsoluteDate date) Get the Greenwich mean sidereal time, in radians.
static double	GTODProvider.getRotationRate(AbsoluteDate date) Get the rotation rate of the Earth.
Transform	VEISProvider.getTransform(AbsoluteDate date) Get the transform from GTOD at specified date.
Transform	MODProvider.getTransform(AbsoluteDate date) Get the transfrom from parent frame.
Transform	CIRFProvider.getTransform(AbsoluteDate date) Get the transform from GCRF to CIRF2000 at the specified date.
Transform	TEMEProvider.getTransform(AbsoluteDate date) Get the transform from True Of Date date.
Transform	ITRFProvider.getTransform(AbsoluteDate date) Get the transform from TIRF 2000 at specified date.
Transform	TIRFProvider.getTransform(AbsoluteDate date) Get the transform from CIRF 2000 at specified date.
Transform	EODProvider.getTransform(AbsoluteDate date) Get the Transform corresponding to specified date.
Transform	ITRFEquinoxProvider.getTransform(AbsoluteDate date) Get the transform from GTOD at specified date.
Transform	GTODProvider.getTransform(AbsoluteDate date) Get the transform from TOD at specified date.
Transform	TransformProvider.getTransform(AbsoluteDate date) Get the Transform corresponding to specified date.
Transform	HelmertTransformation.getTransform(AbsoluteDate date) Compute the transform at some date.
Transform	TODProvider.getTransform(AbsoluteDate date) Get the transform from Mean Of Date at specified date.
Transform	InterpolatingTransformProvider.getTransform(AbsoluteDate date) Get the Transform corresponding to specified date.
Transform	VEISProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from GTOD at specified date.
Transform	MODProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transfrom from parent frame.
Transform	CIRFProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from GCRF to CIRF2000 at the specified date.
Transform	TEMEProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from True Of Date date.
Transform	ITRFProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from TIRF 2000 at specified date.
Transform	TIRFProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from CIRF 2000 at specified date.
Transform	EODProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	FixedTransformProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	ITRFEquinoxProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from GTOD at specified date.
Transform	GTODProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from TOD at specified date.
Transform	TransformProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	HelmertTransformation.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Compute the transform at some date.
Transform	TODProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the transform from Mean Of Date at specified date.
Transform	InterpolatingTransformProvider.getTransform(AbsoluteDate date, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	VEISProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the Transform corresponding to specified date.
Transform	CIRFProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from GCRF to CIRF2000 at the specified date.
Transform	TEMEProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from True Of Date date.
Transform	ITRFProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from TIRF 2000 at specified date.
Transform	TIRFProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from CIRF 2000 at specified date.
Transform	EODProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the Transform corresponding to specified date.
Transform	ITRFEquinoxProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from GTOD at specified date.
Transform	GTODProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from TOD at specified date.
Transform	TransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the Transform corresponding to specified date.
Transform	HelmertTransformation.getTransform(AbsoluteDate date, FramesConfiguration config) Compute the transform at some date.
Transform	TODProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the transform from Mean Of Date at specified date.
Transform	InterpolatingTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config) Get the Transform corresponding to specified date.
Transform	VEISProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	CIRFProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from GCRF to CIRF2000 at the specified date.
Transform	TEMEProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from True Of Date date.
Transform	ITRFProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from TIRF 2000 at specified date.
Transform	TIRFProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from CIRF 2000 at specified date.
Transform	EODProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	FixedTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	ITRFEquinoxProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from GTOD at specified date.
Transform	GTODProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from TOD at specified date.
Transform	TransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
Transform	HelmertTransformation.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Compute the transform at some date.
Transform	TODProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the transform from Mean Of Date at specified date.
Transform	InterpolatingTransformProvider.getTransform(AbsoluteDate date, FramesConfiguration config, boolean computeSpinDerivatives) Get the Transform corresponding to specified date.
static Transform	Transform.interpolate(AbsoluteDate date, boolean useVelocities, boolean useRotationRates, Collection<Transform> sample) Interpolate a transform from a sample set of existing transforms.
static Transform	Transform.interpolate(AbsoluteDate date, boolean useVelocities, boolean useRotationRates, Collection<Transform> sample, boolean computeSpinDerivative) Interpolate a transform from a sample set of existing transforms.
Transform	Transform.interpolate(AbsoluteDate interpolationDate, Collection<Transform> sample) Get an interpolated instance.
Transform	Transform.interpolate(AbsoluteDate interpolationDate, Collection<Transform> sample, boolean computeSpinDerivative) Get an interpolated instance.

Constructors in org.orekit.frames.transformations that throw OrekitException

GTODProvider()
Simple constructor.

H0MinusNProvider(AbsoluteDate h0MinusN, double longitude)
Simple constructor.

ITRFEquinoxProvider()
Simple constructor.

TODProvider(boolean applyEOPCorr)
Simple constructor.

VEISProvider()
Constructor for the singleton.

Uses of OrekitException in org.orekit.models.earth

Methods in org.orekit.models.earth that throw OrekitException

static void	GeoMagneticFieldFactory.addDefaultGeoMagneticModelReader(GeoMagneticFieldFactory.FieldModel type) Add a default reader for geomagnetic models.
static void	GeoMagneticFieldFactory.addGeoMagneticModelReader(GeoMagneticFieldFactory.FieldModel type, GeoMagneticModelReader reader) Add a reader for geomagnetic models.
GeoMagneticElements	GeoMagneticField.calculateField(Vector3D point, Frame frame, AbsoluteDate date) Calculate the magnetic field at the specified point identified by the coordinates of the point and the reference point.
static double	GeoMagneticField.getDecimalYear(AbsoluteDate date) Utility function to get a decimal year for a given AbsoluteDate.
static FixedDelayModel	FixedDelayModel.getDefaultModel(double height) Returns the default model, loading delay values from the file "tropospheric-delay.txt".
static GeoMagneticField	GeoMagneticFieldFactory.getField(GeoMagneticFieldFactory.FieldModel type, AbsoluteDate year) Get the GeoMagneticField for the given model type and year.
static GeoMagneticField	GeoMagneticFieldFactory.getField(GeoMagneticFieldFactory.FieldModel type, double year) Get the GeoMagneticField for the given model type and year.
static GeoMagneticField	GeoMagneticFieldFactory.getIGRF(AbsoluteDate year) Get the IGRF model for the given year.
static GeoMagneticField	GeoMagneticFieldFactory.getIGRF(double year) Get the IGRF model for the given year.
static GeoMagneticField	GeoMagneticFieldFactory.getWMM(AbsoluteDate year) Get the WMM model for the given year.
static GeoMagneticField	GeoMagneticFieldFactory.getWMM(double year) Get the WMM model for the given year.
abstract void	GeoMagneticModelReader.loadData(InputStream input, String name) Load data from a stream.
void	COFFileFormatReader.loadData(InputStream input, String name)
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.

Constructors in org.orekit.models.earth that throw OrekitException

FixedDelayModel(String supportedName, double height)
Creates a new FixedDelayModel instance, and loads the delay values from the given resource via the DataProvidersManager.

Uses of OrekitException in org.orekit.orbits

Methods in org.orekit.orbits that throw OrekitException

abstract Orbit	OrbitType.convertOrbit(Orbit initOrbit, Frame frame) Convert an orbit from a given orbit type to an other in a wished frame.
PVCoordinates	Orbit.getPVCoordinates(AbsoluteDate otherDate, Frame otherFrame) Get the PVCoordinates of the body in the selected frame.
PVCoordinates	Orbit.getPVCoordinates(Frame outputFrame) Get the PVCoordinates in a specified frame.

Uses of OrekitException in org.orekit.parameter

Methods in org.orekit.parameter that throw OrekitException

void	IJacobiansParameterizable.addDAccDParam(SpacecraftState s, Parameter param, double[] dAccdParam) Compute acceleration derivatives with respect to additional parameters.
void	IJacobiansParameterizable.addDAccDState(SpacecraftState s, double[][] dAccdPos, double[][] dAccdVel) Compute acceleration derivatives with respect to state parameters.

Constructors in org.orekit.parameter that throw OrekitException

PiecewiseFunction(ArrayList<IParamDiffFunction> flist, ArrayList<AbsoluteDate> xlist)
Simple constructor with 2 lists (IParamDiffFunction and AbsoluteDate) where dates list represent the connection points between functions.

Uses of OrekitException in org.orekit.propagation

Methods in org.orekit.propagation that throw OrekitException

protected SpacecraftState	AbstractPropagator.acceptStep(AbsoluteDate target, double epsilon) Accept a step, triggering events and step handlers.
SpacecraftState	SpacecraftState.addAdditionalState(String name, double[] state) Add an additional state to the additional states map.
SpacecraftState	SpacecraftState.addAttitudeToAdditionalStates(AttitudeEquation.AttitudeType attitudeType) Add attitude to the additional states map.
SpacecraftState	OsculatingToMeanElementsConverter.convert() Convert an osculating orbit into a mean orbit, in DSST sense.
double[]	SpacecraftState.getAdditionalState(String name) Get one additional state.
Attitude	SpacecraftState.getAttitude(Frame outputFrame) Get the default attitude : the attitude for forces computation in given output frame.
Attitude	SpacecraftState.getAttitude(LOFType lofType) Get the default attitude : the attitude for forces computation in given local orbital frame.
Attitude	SpacecraftState.getAttitudeEvents(Frame outputFrame) Get the attitude for events computation in given output frame.
Attitude	SpacecraftState.getAttitudeEvents(LOFType lofType) Get the attitude for events computation in given local orbital frame.
Attitude	SpacecraftState.getAttitudeForces(Frame outputFrame) Get the attitude for forces computation in given output frame.
Attitude	SpacecraftState.getAttitudeForces(LOFType lofType) Get the attitude for forces computation in given local orbital frame.
SpacecraftState	Propagator.getInitialState() Get the propagator initial state.
SpacecraftState	AbstractPropagator.getInitialState() Get the propagator initial state.
double	SpacecraftState.getMass(String partName) Get the mass of the given part.
PVCoordinates	AbstractPropagator.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
PVCoordinates	SpacecraftState.getPVCoordinates(Frame outputFrame) Get the PVCoordinates in given output frame.
SpacecraftState	SpacecraftState.interpolate(AbsoluteDate date, Collection<SpacecraftState> sample) Get an interpolated instance.
protected void	AbstractPropagator.manageStateFrame() Manage the state frame : the orbit to propagate is converted in the propagation frame.
Orbit	MeanOsculatingElementsProvider.mean2osc(Orbit orbit) Convert provided mean orbit into osculating elements.
Orbit	MeanOsculatingElementsProvider.osc2mean(Orbit orbit) Convert provided osculating orbit into mean elements.
Orbit	MeanOsculatingElementsProvider.propagateMeanOrbit(AbsoluteDate date) Propagate mean orbit until provided date.
void	Propagator.setOrbitFrame(Frame frame) Set propagation frame.
void	AbstractPropagator.setOrbitFrame(Frame frame) Set propagation frame.
Transform	SpacecraftState.toTransform() Compute the transform from orbit/attitude reference frame to spacecraft frame.
Transform	SpacecraftState.toTransform(Frame frame) Compute the transform from specified frame to spacecraft frame.
Transform	SpacecraftState.toTransform(Frame frame, LOFType lofType) Compute the transform from specified frame to local orbital frame.
Transform	SpacecraftState.toTransformEvents() Compute the transform from orbit/attitude (for events computation) reference frame to spacecraft frame.
Transform	SpacecraftState.toTransformEvents(Frame frame) Compute the transform from specified reference frame to spacecraft frame.
Transform	SpacecraftState.toTransformForces() Compute the transform from orbit/attitude (for forces computation) reference frame to spacecraft frame.
Transform	SpacecraftState.toTransformForces(Frame frame) Compute the transform from specified frame to spacecraft frame.
void	MassProvider.updateMass(String partName, double mass) Update the mass of the given part.
SpacecraftState	SpacecraftState.updateMass(String partName, double newMass) Update the mass of the given part.

Constructors in org.orekit.propagation that throw OrekitException

SpacecraftState(double[] y, OrbitType orbitType, PositionAngle angleType, AbsoluteDate date, double mu, Frame frame, Map<String,AdditionalStateInfo> addStatesInfo, AttitudeProvider attProviderForces, AttitudeProvider attProviderEvents)
Build a spacecraft from an array (a state vector) and an additional states informations map.

Uses of OrekitException in org.orekit.propagation.analytical

Methods in org.orekit.propagation.analytical that throw OrekitException

SpacecraftState	AdapterPropagator.DifferentialEffect.apply(SpacecraftState original) Apply the effect to a spacecraft state.
SpacecraftState	J2DifferentialEffect.apply(SpacecraftState state1) Apply the effect to a spacecraft state.
Orbit	EcksteinHechlerPropagator.computeMeanOrbit(Orbit osculating) Deprecated. use EcksteinHechlerPropagator.osc2mean(Orbit) instead
protected Orbit	AbstractLyddanePropagator.computeSecular(Orbit orbit, AbstractLyddanePropagator.LyddaneParametersType fromType) Compute secular orbit in body frame from provided orbit.
protected Orbit	AbstractLyddanePropagator.convertFrame(Orbit orbit, Frame outputFrame) Convert provided orbit in output frame.
SpacecraftState	AdapterPropagator.getInitialState() Get the propagator initial state.
protected void	KeplerianPropagator.manageStateFrame() Manage the state frame : the orbit to propagate is converted in the propagation frame.
Orbit	EcksteinHechlerPropagator.mean2osc(Orbit orbit) Convert provided mean orbit into osculating elements.
Orbit	LyddaneLongPeriodPropagator.mean2osc(Orbit orbit) Convert provided mean orbit into osculating elements.
Orbit	LyddaneSecularPropagator.mean2osc(Orbit orbit) Convert provided mean orbit into osculating elements.
Orbit	EcksteinHechlerPropagator.osc2mean(Orbit orbit) Convert provided osculating orbit into mean elements.
Orbit	LyddaneLongPeriodPropagator.osc2mean(Orbit orbit) Convert provided osculating orbit into mean elements.
Orbit	LyddaneSecularPropagator.osc2mean(Orbit orbit) Convert provided osculating orbit into mean elements.
Orbit	EcksteinHechlerPropagator.propagateMeanOrbit(AbsoluteDate date) Propagate mean orbit until provided date.
Orbit	AbstractLyddanePropagator.propagateMeanOrbit(AbsoluteDate date) Propagate mean orbit until provided date.
Orbit	LyddaneSecularPropagator.propagateMeanOrbit(AbsoluteDate date) Propagate mean orbit until provided date.
protected void	AbstractLyddanePropagator.updateSecularOrbit(Orbit secularOrbit) Update for secular Orbit.

Constructors in org.orekit.propagation.analytical that throw OrekitException

AbstractLyddanePropagator(Orbit secularOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents, MassProvider massProvider)
Generic constructor.

J2DifferentialEffect(Orbit orbit0, Orbit orbit1, boolean applyBefore, PotentialCoefficientsProvider 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, PotentialCoefficientsProvider gravityField)
Simple constructor.

J2SecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, Frame frameIn)
Constructor without attitude provider and mass provider.

J2SecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, Frame frameIn, AttitudeProvider attitudeProvider)
Constructor without mass provider.

J2SecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, Frame frameIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents)
Constructor without mass provider.

J2SecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, Frame frameIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents, MassProvider massProvider)
Generic constructor.

J2SecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, Frame frameIn, AttitudeProvider attitudeProvider, MassProvider massProvider)
Generic constructor.

J2SecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, Frame frameIn, MassProvider massProvider)
Constructor without attitude provider.

LyddaneLongPeriodPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn)
Constructor without attitude provider and mass provider.

LyddaneLongPeriodPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvider)
Constructor without mass provider.

LyddaneLongPeriodPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents)
Constructor without mass provider.

LyddaneLongPeriodPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents, MassProvider massProvider)
Generic constructor.

LyddaneLongPeriodPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvider, MassProvider massProvider)
Generic constructor.

LyddaneLongPeriodPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, MassProvider massProvider)
Constructor without attitude provider.

LyddaneSecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn)
Constructor without attitude provider and mass provider.

LyddaneSecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvider)
Constructor without mass provider.

LyddaneSecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents)
Constructor without mass provider.

LyddaneSecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvForces, AttitudeProvider attitudeProvEvents, MassProvider massProvider)
Generic constructor.

LyddaneSecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, AttitudeProvider attitudeProvider, MassProvider massProvider)
Generic constructor.

LyddaneSecularPropagator(Orbit initialOrbit, double referenceRadiusIn, double muIn, double c20In, double c30In, double c40In, double c50In, Frame frameIn, ParametersType parametersTypeIn, MassProvider massProvider)
Constructor without attitude provider.

Uses of OrekitException in org.orekit.propagation.analytical.covariance

Methods in org.orekit.propagation.analytical.covariance that throw OrekitException

SymmetricMatrix	OrbitCovariance.getCovarianceMatrix(Orbit refOrbit, OrbitType outCovType, Frame outFrame) Covariance matrix getter.
RealMatrix	CovarianceInterpolation.interpolate(AbsoluteDate t) Computes the interpolation of a covariance matrix based on its two surrounding covariance matrices which define the interpolation interval allowed.
double[][]	CovarianceInterpolation.interpolateArray(AbsoluteDate t) Computes the interpolation of a covariance matrix based on its two surrounding covariance matrices which define the interpolation interval allowed.

Constructors in org.orekit.propagation.analytical.covariance that throw OrekitException

CovarianceInterpolation(AbsoluteDate t1In, double[][] matrix1, AbsoluteDate t2In, double[][] matrix2, int order, Orbit orbitSatellite, double muValue)
Constructor of the class CovarianceInterpolation

CovarianceInterpolation(AbsoluteDate t1In, RealMatrix matrix1, AbsoluteDate t2In, RealMatrix matrix2, int order, Orbit orbitSatellite, double muValue)
Constructor of the class CovarianceInterpolation

Uses of OrekitException in org.orekit.propagation.analytical.tle

Methods in org.orekit.propagation.analytical.tle that throw OrekitException

protected double[]	LevenbergMarquardtOrbitConverter.fit(double[] initial) Find the TLE elements that minimize the mean square error for a sample of states.
protected double[]	DifferentialOrbitConverter.fit(double[] initial) Find the TLE elements that minimize the mean square error for a sample of states.
protected abstract double[]	AbstractTLEFitter.fit(double[] initial) Find the TLE elements that minimize the mean square error for a sample of states.
Set<Integer>	TLESeries.getAvailableSatelliteNumbers() Get the available satellite numbers.
String	TLE.getLine1() Get the first line.
PVCoordinates	TLEPropagator.getPVCoordinates(AbsoluteDate date) Get the extrapolated position and velocity from an initial TLE.
PVCoordinates	TLESeries.getPVCoordinates(AbsoluteDate date) Get the extrapolated position and velocity from an initial date.
protected double[]	AbstractTLEFitter.getResiduals(double[] parameters) Get the residuals for a given position/velocity/B* parameters set.
protected double	AbstractTLEFitter.getRMS(double[] parameters) Get the RMS for a given position/velocity/B* parameters set.
static boolean	TLE.isFormatOK(String line1, String line2) Check the lines format validity.
void	TLESeries.loadData(InputStream input, String name) Load data from a stream.
void	TLESeries.loadTLEData() Load TLE data for a specified object.
void	TLESeries.loadTLEData(int satelliteNumber) Load TLE data for a specified object.
void	TLESeries.loadTLEData(int launchYear, int launchNumber, String launchPiece) Load TLE data for a specified object.
static TLEPropagator	TLEPropagator.selectExtrapolator(TLE tle) Selects the extrapolator to use with the selected TLE.
static TLEPropagator	TLEPropagator.selectExtrapolator(TLE tle, AttitudeProvider attitudeProviderForces, AttitudeProvider attitudeProviderEvents, MassProvider mass) Selects the extrapolator to use with the selected TLE.
static TLEPropagator	TLEPropagator.selectExtrapolator(TLE tle, AttitudeProvider attitudeProvider, MassProvider 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).
TLE	AbstractTLEFitter.toTLE(List<SpacecraftState> states, double positionTolerance, boolean positionOnly, boolean withBStar) Find the TLE elements that minimize the mean square error for a sample of states.

Constructors in org.orekit.propagation.analytical.tle that throw OrekitException

TLE(String line1, String line2)
Simple constructor from unparsed two lines.

TLEPropagator(TLE initialTLE, AttitudeProvider attitudeProviderForces, AttitudeProvider attitudeProviderEvents, MassProvider mass)
Protected constructor for derived classes.

TLEPropagator(TLE initialTLE, AttitudeProvider attitudeProvider, MassProvider mass)
Protected constructor for derived classes.

Uses of OrekitException in org.orekit.propagation.analytical.twod

Methods in org.orekit.propagation.analytical.twod that throw OrekitException

double[]	Analytical2DOrbitModel.propagateModel(AbsoluteDate date) Propagate each parameter model to specified date and return an array of 6 values.
double[]	Analytical2DOrbitModel.propagateModel(AbsoluteDate date, int[] orders) Propagate each parameter model to specified date and return an array of 6 values.

Uses of OrekitException in org.orekit.propagation.events

Methods in org.orekit.propagation.events that throw OrekitException

boolean	EventState.evaluateStep(OrekitStepInterpolator interpolator) Evaluate the impact of the proposed step on the event detector. See Orekit issue 110 for more information.
boolean	EventState.evaluateStep(SpacecraftState state) Evaluate the impact of the proposed step on the event handler.
EventDetector.Action	ThreeBodiesAngleDetector.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward)
EventDetector.Action	ExtremaThreeBodiesAngleDetector.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward)
EventDetector.Action	ThreeBodiesAngleDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an angle event and choose what to do next.
EventDetector.Action	ExtremaDistanceDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an extrema distance event and choose what to do next.
abstract EventDetector.Action	AbstractDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
EventDetector.Action	EventDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
EventDetector.Action	NodeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a node crossing event and choose what to do next.
EventDetector.Action	NthOccurrenceDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
EventDetector.Action	LatitudeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a latitude reaching event and choose what to do next.
EventDetector.Action	DistanceDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a distance event and choose what to do next.
EventDetector.Action	DihedralFieldOfViewDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an fov event and choose what to do next.
EventDetector.Action	AltitudeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an altitude event and choose what to do next.
EventDetector.Action	ApsideDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an apside crossing event and choose what to do next.
EventDetector.Action	EventShifter.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an event and choose what to do next.
EventDetector.Action	ExtremaThreeBodiesAngleDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a min or max angle event and choose what to do next.
EventDetector.Action	CircularFieldOfViewDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an fov event and choose what to do next.
EventDetector.Action	BetaAngleDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handles a beta angle event and chooses what to do next.
EventDetector.Action	EclipseDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an eclipse event and choose what to do next.
EventDetector.Action	LocalTimeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a local time event and choose what to do next.
EventDetector.Action	NadirSolarIncidenceDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a solar incidence event and choose what to do next.
EventDetector.Action	ElevationDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an elevation event and choose what to do next.
EventDetector.Action	DateDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a date event and choose what to do next.
EventDetector.Action	ExtremaLatitudeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an extrema latitude event and choose what to do next.
EventDetector.Action	GroundMaskElevationDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an azimuth-elevation event and choose what to do next.
EventDetector.Action	AOLDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an AOL event and choose what to do next.
EventDetector.Action	AlignmentDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an alignment event and choose what to do next.
EventDetector.Action	ExtremaElevationDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an extrema distance event and choose what to do next.
EventDetector.Action	SolarTimeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a solar time event and choose what to do next.
EventDetector.Action	ExtremaLongitudeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an extrema distance event and choose what to do next.
EventDetector.Action	LongitudeDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle a longitude reaching event and choose what to do next.
EventDetector.Action	ApparentElevationDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an apparent elevation event and choose what to do next.
EventDetector.Action	AnomalyDetector.eventOccurred(SpacecraftState s, boolean increasing, boolean forward) Handle an anomaly event and choose what to do next.
double	ThreeBodiesAngleDetector.g(Map<String,SpacecraftState> s)
double	ExtremaThreeBodiesAngleDetector.g(Map<String,SpacecraftState> s)
double	ThreeBodiesAngleDetector.g(SpacecraftState s) Compute the value of the switching function.
double	ExtremaDistanceDetector.g(SpacecraftState s)
abstract double	AbstractDetector.g(SpacecraftState s) Compute the value of the switching function.
double	EventDetector.g(SpacecraftState s) Compute the value of the switching function.
double	NodeDetector.g(SpacecraftState s) Compute the value of the switching function.
double	NthOccurrenceDetector.g(SpacecraftState s) Compute the value of the switching function.
double	LatitudeDetector.g(SpacecraftState s)
double	DistanceDetector.g(SpacecraftState s)
double	DihedralFieldOfViewDetector.g(SpacecraftState s) Compute the value of the switching function.
double	AltitudeDetector.g(SpacecraftState s) Compute the value of the switching function.
double	ApsideDetector.g(SpacecraftState s) Compute the value of the switching function.
double	EventShifter.g(SpacecraftState s) Compute the value of the switching function.
double	ExtremaThreeBodiesAngleDetector.g(SpacecraftState s)
double	CircularFieldOfViewDetector.g(SpacecraftState s) Compute the value of the switching function.
double	BetaAngleDetector.g(SpacecraftState s)
double	EclipseDetector.g(SpacecraftState s) Compute the value of the switching function.
double	LocalTimeDetector.g(SpacecraftState s)
double	NadirSolarIncidenceDetector.g(SpacecraftState s)
double	ElevationDetector.g(SpacecraftState s) Compute the value of the switching function.
double	DateDetector.g(SpacecraftState s) Compute the value of the switching function.
double	ExtremaLatitudeDetector.g(SpacecraftState s)
double	GroundMaskElevationDetector.g(SpacecraftState s) Compute the value of the switching function.
double	AOLDetector.g(SpacecraftState s)
double	AlignmentDetector.g(SpacecraftState s) Compute the value of the switching function.
double	ExtremaElevationDetector.g(SpacecraftState s)
double	NullMassDetector.g(SpacecraftState s) Compute the value of the switching function.
double	SolarTimeDetector.g(SpacecraftState s)
double	ExtremaLongitudeDetector.g(SpacecraftState s)
double	LongitudeDetector.g(SpacecraftState s)
double	ApparentElevationDetector.g(SpacecraftState s) Compute the value of the switching function.
double	NullMassPartDetector.g(SpacecraftState s) Compute the value of the switching function.
double	AnomalyDetector.g(SpacecraftState s)
void	EventState.reinitializeBegin(SpacecraftState state0) Reinitialize the beginning of the step.
SpacecraftState	EventState.reset(SpacecraftState oldState) Let the event detector reset the state if it wants.
SpacecraftState	AbstractDetector.resetState(SpacecraftState oldState) Reset the state (including additional states) prior to continue propagation.
SpacecraftState	EventDetector.resetState(SpacecraftState oldState) Reset the state (including additional states) prior to continue propagation.
SpacecraftState	NthOccurrenceDetector.resetState(SpacecraftState oldState) Reset the state (including additional states) prior to continue propagation.
SpacecraftState	NullMassPartDetector.resetState(SpacecraftState oldState) Reset the state (including additional states) prior to continue propagation.
Map<String,SpacecraftState>	ThreeBodiesAngleDetector.resetStates(Map<String,SpacecraftState> oldStates)
Map<String,SpacecraftState>	ExtremaThreeBodiesAngleDetector.resetStates(Map<String,SpacecraftState> oldStates)
void	EventState.stepAccepted(SpacecraftState state) Acknowledge the fact the step has been accepted by the propagator.
void	EventState.storeState(SpacecraftState state0) Reinitialize event state with provided time and state.

Constructors in org.orekit.propagation.events that throw OrekitException

DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget, Vector3D center, Vector3D axis1, double halfAperture1, Vector3D axis2, double halfAperture2, double maxCheck)
Build a new instance.

DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget, Vector3D center, Vector3D axis1, double halfAperture1, Vector3D axis2, double halfAperture2, double maxCheck, double epsilon)
Build a new instance.

DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget, Vector3D center, Vector3D axis1, double halfAperture1, Vector3D axis2, double halfAperture2, double maxCheck, EventDetector.Action entry, EventDetector.Action exit)
Build a new instance.

DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget, Vector3D center, Vector3D axis1, double halfAperture1, Vector3D axis2, double halfAperture2, double maxCheck, EventDetector.Action entry, EventDetector.Action exit, boolean removeEntry, boolean removeExit)
Build a new instance.

DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget, Vector3D center, Vector3D axis1, double halfAperture1, Vector3D axis2, double halfAperture2, double maxCheck, EventDetector.Action entry, EventDetector.Action exit, boolean removeEntry, boolean removeExit, double epsilon)
Build a new instance.

DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget, Vector3D center, Vector3D axis1, double halfAperture1, Vector3D axis2, double halfAperture2, double maxCheck, EventDetector.Action entry, EventDetector.Action exit, double epsilon)
Build a new instance.

LocalTimeDetector(double localTime, Frame bodyFrame)
Constructor for a LocalTimeDetector instance.

LocalTimeDetector(double localTime, Frame bodyFrame, double maxCheck, double threshold)
Constructor for a LocalTimeDetector instance with complimentary parameters.

LocalTimeDetector(double localTime, Frame bodyFrame, double maxCheck, double threshold, EventDetector.Action action)
Constructor for a LocalTimeDetector instance with complimentary parameters.

LocalTimeDetector(double localTime, Frame bodyFrame, double maxCheck, double threshold, EventDetector.Action action, boolean remove)
Constructor for a LocalTimeDetector instance with complimentary parameters.

NadirSolarIncidenceDetector(double incidence, BodyShape earth, double maxCheck, double threshold)
Constructor for the nadir point solar incidence detector

NadirSolarIncidenceDetector(double incidence, BodyShape earth, double maxCheck, double threshold, EventDetector.Action action)
Constructor for the nadir point solar incidence detector

NadirSolarIncidenceDetector(double incidence, BodyShape earth, double maxCheck, double threshold, EventDetector.Action action, boolean remove)
Constructor for the nadir point solar incidence detector

SolarTimeDetector(double solarTime)
Constructor for a SolarTimeDetector instance.

SolarTimeDetector(double solarTime, double maxCheck, double threshold)
Constructor for a SolarTimeDetector instance with complimentary parameters.

SolarTimeDetector(double solarTime, double maxCheck, double threshold, EventDetector.Action action)
Constructor for a SolarTimeDetector instance with complimentary parameters.

SolarTimeDetector(double solarTime, double maxCheck, double threshold, EventDetector.Action action, boolean remove)
Constructor for a SolarTimeDetector instance with complimentary parameters.

Uses of OrekitException in org.orekit.propagation.events.multi

Methods in org.orekit.propagation.events.multi that throw OrekitException

EventDetector.Action	MultiEventDetector.eventOccurred(Map<String,SpacecraftState> s, boolean increasing, boolean forward) Handle an event and choose what to do next.
double	MultiEventDetector.g(Map<String,SpacecraftState> s) Compute the value of the switching function.
Map<String,SpacecraftState>	MultiEventDetector.resetStates(Map<String,SpacecraftState> oldStates) Reset the states (including additional states) prior to continue propagation.

Uses of OrekitException in org.orekit.propagation.numerical

Methods in org.orekit.propagation.numerical that throw OrekitException

void	TimeDerivativesEquations.addAcceleration(Vector3D gamma, Frame frame) Add the contribution of an acceleration expressed in some inertial frame.
void	PartialDerivativesEquations.computeDerivatives(SpacecraftState s, TimeDerivativesEquations adder) Compute the derivatives related to the additional state parameters.
void	AdditionalEquations.computeDerivatives(SpacecraftState s, TimeDerivativesEquations adder) Compute the derivatives related to the additional state parameters.
JacobiansMapper	PartialDerivativesEquations.getMapper() Get a mapper between two-dimensional Jacobians and one-dimensional additional state.
double[]	JacobiansMapper.getParametersJacobian(Parameter parameter, SpacecraftState state) Get the Jacobian with respect to provided parameter parameter.
void	JacobiansMapper.getParametersJacobian(Parameter parameter, SpacecraftState state, double[] dYdP) Get the Jacobian with respect to provided parameter parameter.
double[][]	JacobiansMapper.getParametersJacobian(SpacecraftState state) Get the Jacobian with respect to parameters.
void	JacobiansMapper.getParametersJacobian(SpacecraftState state, double[][] dYdP) Get the Jacobian with respect to parameters.
PVCoordinates	NumericalPropagator.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
double[][]	JacobiansMapper.getStateJacobian(SpacecraftState state) Get the Jacobian with respect to state.
void	JacobiansMapper.getStateJacobian(SpacecraftState state, double[][] dYdY0) Get the Jacobian with respect to state.
void	NumericalPropagator.setAdditionalStateTolerance(String name, double[] absTol, double[] relTol) Add additional state tolerances.
SpacecraftState	PartialDerivativesEquations.setInitialJacobians(SpacecraftState s, double[][] dY1dY0) Set the initial value of the Jacobian with respect to state.
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 paramDimension) Set the initial value of the Jacobian with respect to state and parameter.
SpacecraftState	PartialDerivativesEquations.setInitialJacobians(SpacecraftState s, Parameter parameter, double[] dY1dP) Set the initial value of the Jacobian with respect to state.
void	NumericalPropagator.setOrbitFrame(Frame frame) Set propagation frame.

Constructors in org.orekit.propagation.numerical that throw OrekitException

PartialDerivativesEquations(String name, NumericalPropagator propagator)
Simple constructor.

Uses of OrekitException in org.orekit.propagation.precomputed

Methods in org.orekit.propagation.precomputed that throw OrekitException

protected Attitude[]	AbstractEphemeris.attitudesInterpolation(SpacecraftState[] tab, int order, int i0, AbsoluteDate date) This method is called only if attitudes are supported.
protected double[]	AbstractEphemeris.convertTab(SpacecraftState[] tab, int order) Convert the SpacecraftState[] into a double[] which represents the duration from the first state date.
static SpacecraftState[]	AbstractEphemeris.generateSpacecraftState(PVCoordinatesProvider pvProv, AttitudeProvider attProvForces, AttitudeProvider attProvEvents, double step, AbsoluteDateInterval ptsInterval, Frame frame, double mu) Creates a spacecraft array with constant step size.
SpacecraftState	IntegratedEphemeris.getInitialState() Get the propagator initial state.
SpacecraftState	AbstractEphemeris.getInitialState() Get the propagator initial state.
protected SpacecraftState	LagrangeEphemeris.getInterpolatedSpacecraftState(AbsoluteDate date) Deprecated. Get the interpolated spacecraft state.
protected SpacecraftState	HermiteEphemeris.getInterpolatedSpacecraftState(AbsoluteDate date) Deprecated. Get the interpolated spacecraft state.
protected abstract SpacecraftState	AbstractEphemeris.getInterpolatedSpacecraftState(AbsoluteDate date) Get the interpolated spacecraft state.
PVCoordinates	IntegratedEphemeris.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
PVCoordinates	Ephemeris.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
protected int	AbstractEphemeris.indexValidity(SpacecraftState[] tab, int index) Checks if interpolation is valid : meaning if 0<= index +1 -interpOrder/2 or index + interpOrder/2 <= maximalIndex
protected AbsoluteDateInterval	AbstractEphemeris.intervalValidity(SpacecraftState[] tab) Corrects the min and max dates with the constraint of the number of interpolations points required.
void	IntegratedEphemeris.manageStateFrame() In this class, nothing as to be done in the frame managing before propagation because propagation will be performed in Frame referenceFrame It just throws an OrekitException if this frame is non inertial or pseudo-inertial.

Constructors in org.orekit.propagation.precomputed that throw OrekitException

AbstractEphemeris(SpacecraftState[] tab)
Simple constructor of class AbstractEphemeris, with interpolation order equal to 2.

AbstractEphemeris(SpacecraftState[] tab, int order)
Simple constructor of class AbstractEphemeris with defined interpolation order.

AbstractEphemeris(SpacecraftState[] tab, int order, ISearchIndex algo)
Constructor of class AbstractEphemeris where a search index algorithm is passed on the constructor.

AbstractEphemeris(SpacecraftState[] tab, ISearchIndex algo)
Constructor of class AbstractEphemeris where a search index algorithm is passed on the constructor and the order is by default equal to 2.

HermiteEphemeris(PVCoordinatesProvider pvProv, AttitudeProvider attProvForces, AttitudeProvider attProvEvents, double step, AbsoluteDateInterval intervalOfPoints, Frame frame2, double mu2)
Deprecated. Constructor n°1 using a PV coordinates provider and an AttitudeProvider to build the interpolation points.

HermiteEphemeris(SpacecraftState[] tab)
Deprecated. Constructor of class HermiteEphemeris, used by other constructors of this class.

HermiteEphemeris(SpacecraftState[] tab, ISearchIndex algo)
Deprecated. Constructor of class HermiteEphemeris, where a search index algorithm is passed on the constructor.

HermiteEphemeris(SpacecraftState[] tabStates, Vector3D[] tabAcc)
Deprecated. Constructor of class HermiteEphemeris, where a search index algorithm is passed on the constructor.

HermiteEphemeris(SpacecraftState[] tab, Vector3D[] tabAcc, ISearchIndex algo)
Deprecated. Constructor of class HermiteEphemeris, where a search index algorithm is passed on the constructor.

IntegratedEphemeris(List<AbsoluteDate> startDates, List<AbsoluteDate> minDates, List<AbsoluteDate> maxDates, OrbitType orbitType, PositionAngle angleType, AttitudeProvider attitudeForcesProvider, AttitudeProvider attitudeEventsProvider, Map<String,AdditionalStateInfo> additionalStateInfos, List<ContinuousOutputModel> models, Frame referenceFrame, double mu)
Creates a new instance of IntegratedEphemeris.

LagrangeEphemeris(PVCoordinatesProvider pvProv, AttitudeProvider attProvForces, AttitudeProvider attProvEvents, double step, AbsoluteDateInterval intervalOfPoints, Frame frame, double mu)
Deprecated. Constructor using a PV coordinates provider and an AttitudeProvider to build the interpolation points, defaulting to 8th order for the Lagrange interpolator.

LagrangeEphemeris(PVCoordinatesProvider pvProv, AttitudeProvider attProvForces, AttitudeProvider attProvEvents, double step, AbsoluteDateInterval intervalOfPoints, Frame frame, double mu, int order)
Deprecated. Constructor using a PV coordinates provider and an AttitudeProvider to build the interpolation points.

LagrangeEphemeris(SpacecraftState[] tabulatedStates)
Deprecated. Constructor for a 8th order Lagrange interpolator.

LagrangeEphemeris(SpacecraftState[] tab, int order)
Deprecated. Constructor of class LagrangeEphemeris, used by other constructors of this class.

LagrangeEphemeris(SpacecraftState[] tab, int order, ISearchIndex algo)
Deprecated. Constructor of class LagrangeEphemeris, where a search index algorithm is passed on the constructor.

LagrangeEphemeris(SpacecraftState[] tab, ISearchIndex algo)
Deprecated. Constructor for a 8th order Lagrange interpolator, where a search index algorithm is passed on the constructor.

Uses of OrekitException in org.orekit.propagation.sampling

Methods in org.orekit.propagation.sampling that throw OrekitException

SpacecraftState	BasicStepInterpolator.getInterpolatedState() Get the interpolated state.
SpacecraftState	OrekitStepInterpolator.getInterpolatedState() Get the interpolated state.
SpacecraftState	AdaptedStepHandler.getInterpolatedState() Get the interpolated state.

Uses of OrekitException in org.orekit.time

Methods in org.orekit.time that throw OrekitException

double	LocalTime.computeEquationOfTime(AbsoluteDate date) Compute equation of time in TIRF in the range ]-43200s; 43200s].
double	LocalTime.computeMeanLocalTime(AbsoluteDate date, Vector3D pos, Frame frame) Compute mean local time in TIRF frame in s.
double	LocalTime.computeMeanLocalTime(Orbit orbit) Compute mean local time in TIRF frame in s.
double	LocalTime.computeTrueLocalTime(AbsoluteDate date, Vector3D pos, Frame frame) Compute true local time in TIRF frame in the range [0s; 86400s[.
double	LocalTime.computeTrueLocalTime(Orbit orbit) Compute true local time in TIRF frame in the range [0s; 86400s[.
static GMSTScale	TimeScalesFactory.getGMST() Get the Greenwich Mean Sidereal Time scale.
static UT1Scale	TimeScalesFactory.getUT1() Get the Universal Time 1 scale.
static UTCScale	TimeScalesFactory.getUTC() Get the Universal Time Coordinate scale.
T	TimeInterpolable.interpolate(AbsoluteDate date, Collection<T> sample) Get an interpolated instance.
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).

Uses of OrekitException in org.orekit.utils

Methods in org.orekit.utils that throw OrekitException

PVCoordinates	PVCoordinatesProvider.getPVCoordinates(AbsoluteDate date, Frame frame) Get the PVCoordinates of the body in the selected frame.
PVCoordinates	EphemerisPvLagrange.getPVCoordinates(AbsoluteDate date, Frame frame) Frame can be null : by default the frame of expression is the frame used at instantiation (which is the frame of the first spacecraft state when instantiation is done from a table of spacecraft states).
PVCoordinates	EphemerisPvHermite.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 TimeStampedAngularCoordinates	TimeStampedAngularCoordinates.interpolate(AbsoluteDate date, AngularDerivativesFilter filter, Collection<TimeStampedAngularCoordinates> sample, boolean computeSpinDerivatives) Interpolate angular coordinates.
static AngularCoordinates	AngularCoordinates.interpolate(AbsoluteDate date, boolean useRotationRates, Collection<Pair<AbsoluteDate,AngularCoordinates>> sample) Deprecated. since 3.1 replaced with TimeStampedAngularCoordinates.interpolate(AbsoluteDate, AngularDerivativesFilter, Collection)
static AngularCoordinates	AngularCoordinates.interpolate(AbsoluteDate date, boolean useRotationRates, Collection<Pair<AbsoluteDate,AngularCoordinates>> sample, boolean computeSpinDerivative) Deprecated. since 3.1 replaced with TimeStampedAngularCoordinates.interpolate(AbsoluteDate, AngularDerivativesFilter, Collection, boolean)
void	InterpolationTableLoader.loadData(InputStream input, String name) Loads an bi-variate interpolation table from the given InputStream.
static Vector3D	ReferencePointsDisplacement.solidEarthTidesCorrections(AbsoluteDate date, Vector3D point, Vector3D sun, Vector3D moon) Computes the displacement of reference points due to the effect of the solid Earth tides.
FieldVector3D<DerivativeStructure>	PVCoordinates.toDerivativeStructureVector(int order) Transform the instance to a FieldVector3D<DerivativeStructure>.

Constructors in org.orekit.utils that throw OrekitException

AngularCoordinates(PVCoordinates u1, PVCoordinates u2, PVCoordinates v1, PVCoordinates v2, double tolerance)
Build the rotation that transforms a pair of pv coordinates into another one.

AngularCoordinates(PVCoordinates u1, PVCoordinates u2, PVCoordinates v1, PVCoordinates v2, double tolerance, boolean spinDerivativesComputation)
Build the rotation that transforms a pair of 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.

TimeStampedAngularCoordinates(AbsoluteDate date, PVCoordinates u1, PVCoordinates u2, PVCoordinates v1, PVCoordinates v2, double tolerance, boolean spinDerivativesComputation)
Build the rotation that transforms a pair of pv coordinates into another pair.

Uses of OrekitException in org.orekit.wrenches

Methods in org.orekit.wrenches that throw OrekitException

Vector3D	WrenchModel.computeTorque(SpacecraftState s) Compute the resulting torque at the mass centre of the spacecraft in the frame of the main part.
Vector3D	WrenchModel.computeTorque(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.
Wrench	WrenchModel.computeWrench(SpacecraftState s) Compute the resulting wrench at the mass centre of the spacecraft in the frame of the main part.
Wrench	WrenchModel.computeWrench(SpacecraftState s, Vector3D origin, Frame frame) Compute the resulting wrench.

Constructors in org.orekit.wrenches that throw OrekitException

Wrench(double[] data)
Constructor from an array.