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| Uses of OrekitException in fr.cnes.sirius.patrius.assembly |
|---|
| Methods in fr.cnes.sirius.patrius.assembly that throw OrekitException | |
|---|---|
Assembly |
Vehicle.createAssembly(Frame frame)
Create an Assembly. |
Assembly |
Vehicle.createAssembly(Frame frame,
double cMass,
double cDrag,
double cSRP)
Create an Assembly with multiplicative coefficients to take into
account the change in surface for drag or SRP during a propagation of the
change of dry mass. |
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,
Frame mainPartFrame)
Method to compute the force for the part model (cylinder, parallelepiped, sphere). |
protected static Vector3D |
DirectRadiativeModel.forceOnSphere(SpacecraftState state,
IPart part,
Vector3D flux,
Frame mainPartFrame)
Method to compute the force for a spherical 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 | |
|---|---|
GlobalDragCoefficientProvider(GlobalDragCoefficientProvider.INTERP method,
String filePath)
Constructor. |
|
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 |
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| 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 |
RadiativeSphereProperty.getCrossSection(SpacecraftState state,
Vector3D flux,
Frame mainPartFrame,
Frame partFrame)
|
double |
RadiativeCrossSectionProperty.getCrossSection(SpacecraftState state,
Vector3D flux,
Frame mainPartFrame,
Frame partFrame)
Compute the cross section of main shape using the relative velocity in the part (having the aero property) frame as the direction to provider to the CrossSectionProvider.getCrossSection(Vector3D). |
double |
AeroCrossSectionProperty.getCrossSection(SpacecraftState state,
Vector3D relativeVelocity,
Frame mainPartFrame,
Frame partFrame)
Compute the cross section of main shape using the relative velocity in the part (having the aero property) frame as the direction to provider to the CrossSectionProvider.getCrossSection(Vector3D). |
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. |
|
TankProperty(String nameIn,
double massIn)
Constructor. |
|
| 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 |
ConstantThrustManeuver.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing acceleration. |
void |
EmpiricalForce.addContribution(SpacecraftState state,
TimeDerivativesEquations adder)
|
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 |
EmpiricalForce.addDAccDParam(SpacecraftState state,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters. |
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 |
EmpiricalForce.addDAccDState(SpacecraftState state,
double[][] dAccdPos,
double[][] dAccdVel)
|
void |
VariableThrustManeuver.addDAccDState(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters. |
Vector3D |
EmpiricalForce.computeAcceleration(PVCoordinates pv,
LocalOrbitalFrame localFrameValidation,
Vector3D vectorS,
Frame frame,
SpacecraftState state)
Method to compute the acceleration. |
Vector3D |
ConstantThrustManeuver.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force. |
Vector3D |
EmpiricalForce.computeAcceleration(SpacecraftState state)
|
Vector3D |
VariableThrustManeuver.computeAcceleration(SpacecraftState s)
|
| Constructors in fr.cnes.sirius.patrius.forces that throw OrekitException | |
|---|---|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
double thrust,
double isp,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn)
Constructor for a constant direction in satellite frame and constant thrust. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
double thrust,
double isp,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn,
Frame frameIn)
Constructor for a constant direction in provided frame and constant thrust. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
double thrust,
double isp,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn,
LOFType lofTyp)
Constructor for a constant direction in provided local orbital frame and constant thrust. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
Parameter thrust,
Parameter flowRate,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn)
Constructor for a constant direction in satellite frame and constant thrust using Parameter. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
Parameter thrust,
Parameter flowRate,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn,
Frame frameIn)
Constructor for a constant dDirection in provided frame and constant thrust using Parameter. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
Parameter thrust,
Parameter flowRate,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn,
LOFType lofTyp)
Constructor for a constant direction in provided local orbital frame and constant thrust using Parameter
with a local orbital frame defined by its type. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
PropulsiveProperty engine,
Parameter flowRate,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank)
Constructor for a constant direction in satellite frame and constant thrust using Parameter,
PropulsiveProperty and TankProperty. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
PropulsiveProperty engine,
Parameter flowRate,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank,
Frame frameIn)
Constructor for a constant direction in satellite frame and constant thrust using Parameter,
PropulsiveProperty and TankProperty. |
|
ConstantThrustManeuver(AbsoluteDate date,
double duration,
PropulsiveProperty engine,
Parameter flowRate,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank,
LOFType lofTyp)
Constructor for a constant direction in satellite frame and constant thrust using Parameter,
PropulsiveProperty and TankProperty. |
|
ConstantThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
double thrust,
double isp,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn)
Constructor for a constant direction in satellite frame and constant thrust. |
|
ConstantThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
double thrust,
double isp,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn,
Frame frameIn)
Constructor for a constant direction in provided frame and constant thrust. |
|
ConstantThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
double thrust,
double isp,
Vector3D inDirection,
MassProvider massProvider,
String partNameIn,
LOFType lofTyp)
Constructor for a constant direction in provided local orbital frame and constant thrust. |
|
ConstantThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank)
Constructor for a constant direction in satellite frame and constant thrust with PropulsiveProperty and TankProperty properties. |
|
ConstantThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank,
Frame frameIn)
Constructor for a constant direction in provided frame and constant thrust with PropulsiveProperty and TankProperty properties. |
|
ConstantThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank,
LOFType lofTyp)
Constructor for a constant direction in local orbital frame and constant thrust with PropulsiveProperty and TankProperty properties. |
|
| 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. |
double |
TwoSpinBiasSlew.computeDuration(PVCoordinatesProvider pvProv)
Computes the actual 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 |
AeroAttitudeLaw.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state. |
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 | |
|---|---|
AeroAttitudeLaw(double angleofattack,
double sideslip,
double rollVel,
ExtendedOneAxisEllipsoid earthShape)
Constructor. |
|
AeroAttitudeLaw(double angleofattack,
double sideslip,
double rollVel,
ExtendedOneAxisEllipsoid earthShape,
double dtSpin,
double dtAcc)
Constructor with parameterizable delta-time for spin and acceleration computation. |
|
AeroAttitudeLaw(IParameterizableFunction angleofattack,
IParameterizableFunction sideslip,
IParameterizableFunction rollVel,
ExtendedOneAxisEllipsoid earthShape)
Constructor. |
|
AeroAttitudeLaw(IParameterizableFunction angleofattack,
IParameterizableFunction sideslip,
IParameterizableFunction rollVel,
ExtendedOneAxisEllipsoid earthShape,
double dtSpin,
double dtAcc)
Constructor with parameterizable delta-time for spin and acceleration computation. |
|
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)
Calculate the apparent radius. |
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 |
UserCelestialBody.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the PVCoordinates of the body in the selected 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,
String filename,
int n,
int m)
Create an instance of a central body attraction with normalized coefficients, Helmholtz Polynomials (Balmino model) and specific data. |
static ForceModel |
EarthGravitationalModelFactory.getCunningham(EarthGravitationalModelFactory.GravityFieldNames potentialFileName,
String filename,
int n,
int m)
Create an instance of the gravitational field of a celestial body using Cunningham model and specific data. |
static ForceModel |
EarthGravitationalModelFactory.getDroziner(EarthGravitationalModelFactory.GravityFieldNames potentialFileName,
String filename,
int n,
int m)
Create an instance of the gravitational field of a celestial body using Droziner model and specific data. |
static ForceModel |
EarthGravitationalModelFactory.getGravitationalModel(EarthGravitationalModelFactory.GravityFieldNames potentialFileName,
String filename,
int n,
int m)
Create an default instance of a gravitational field of a celestial body using Balmino model and specific data. |
| 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 |
ConstantThrustError.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the constant thrust error model to the perturbing acceleration. |
void |
ConstantThrustError.addDAccDParam(SpacecraftState state,
Parameter param,
double[] dAccdParam)
|
void |
ConstantThrustError.addDAccDState(SpacecraftState state,
double[][] dAccdPos,
double[][] dAccdVel)
|
SpacecraftState |
SmallManeuverAnalyticalModel.apply(SpacecraftState state1)
Compute the effect of the maneuver on a spacecraft state. |
Vector3D |
ConstantThrustError.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force. |
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. |
| Constructors in org.orekit.frames that throw OrekitException | |
|---|---|
H0MinusNFrame(String name,
AbsoluteDate h0In,
double nIn,
double longitudeIn)
Constructor. |
|
| 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 |
LocalTimeAngleDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle a local time angle 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 |
SolarTimeAngleDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle a solar time angle 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 |
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 |
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 |
LocalTimeAngleDetector.g(SpacecraftState s)
|
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 |
SolarTimeAngleDetector.g(SpacecraftState s)
|
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 |
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 |
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 state,
boolean forceUpdate)
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. |
|
LocalTimeAngleDetector(double localTimeAngle)
Constructor for a LocalTimeDetector instance. |
|
LocalTimeAngleDetector(double localTimeAngle,
double maxCheck,
double threshold)
Constructor for a LocalTimeDetector instance with complimentary parameters. |
|
LocalTimeAngleDetector(double localTimeAngle,
double maxCheck,
double threshold,
EventDetector.Action action)
Constructor for a LocalTimeDetector instance with complimentary parameters. |
|
LocalTimeAngleDetector(double localTimeAngle,
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 |
|
SolarTimeAngleDetector(double solarTimeAngle)
Constructor for a SolarTimeDetector instance. |
|
SolarTimeAngleDetector(double solarTimeAngle,
double maxCheck,
double threshold)
Constructor for a SolarTimeDetector instance with complimentary parameters. |
|
SolarTimeAngleDetector(double solarTimeAngle,
double maxCheck,
double threshold,
EventDetector.Action action)
Constructor for a SolarTimeDetector instance with complimentary parameters. |
|
SolarTimeAngleDetector(double solarTimeAngle,
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 |
Jacobianizer.addDAccDParam(SpacecraftState s,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters. |
void |
Jacobianizer.addDAccDState(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters. |
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 |
LocalTimeAngle.computeEquationOfTime(AbsoluteDate date)
Compute equation of time in TIRF in the range [-43200s; 43200s]. |
double |
LocalTimeAngle.computeMeanLocalTimeAngle(AbsoluteDate date,
Vector3D pos,
Frame frame)
Compute mean local time angle in TIRF frame in the range [-Π, Π[. |
double |
LocalTimeAngle.computeMeanLocalTimeAngle(Orbit orbit)
Compute mean local time angle in TIRF frame in the range [-Π, Π[. |
double |
LocalTimeAngle.computeTrueLocalTimeAngle(AbsoluteDate date,
Vector3D pos,
Frame frame)
Compute true local time angle in TIRF frame in the range [-Π, Π[. |
double |
LocalTimeAngle.computeTrueLocalTimeAngle(Orbit orbit)
Compute true local time angle in TIRF frame in the range [-Π, Π[. |
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. |
int |
AbstractBoundedPVProvider.indexValidity(int index)
Checks if interpolation is valid : meaning if 0<= index +1 -interpOrder/2 or index + interpOrder/2 <= maximalIndex |
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. |
|
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