| Modifier and Type | Method and Description |
|---|---|
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 |
AssemblyBuilder.initMainPartFrame(SpacecraftState state)
Sets up the main frame of the assembly from a "SpacecraftState" object.
|
void |
Assembly.initMainPartFrame(SpacecraftState state)
Initialize the main part's frame using a
SpacecraftState as an input argument. |
void |
Vehicle.setAerodynamicsProperties(AerodynamicCoefficient cx,
AerodynamicCoefficient cz)
Set aerodynamics properties only if possible : main shape must be a sphere or there must be
no solar panels.
|
void |
Vehicle.setAerodynamicsProperties(double cx,
double cz)
Set aerodynamics properties as constants.
|
void |
Vehicle.setDryMass(double mass)
Set vehicle dry mass.
|
void |
Vehicle.setRadiativeProperties(double ka,
double ks,
double kd,
double kaIr,
double ksIr,
double kdIr)
Set radiative properties.
|
void |
MainPart.updateFrame(Transform transform) |
void |
Part.updateFrame(Transform t) |
void |
IPart.updateFrame(Transform t) |
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. |
| Modifier and Type | Method and Description |
|---|---|
void |
RediffusedRadiativeModel.addDAccDParamRediffusedRadiativePressure(SpacecraftState s,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives.
|
void |
RediffusedRadiativeModel.addDAccDStateRediffusedRadiativePressure(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives.
|
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 |
DragLiftModel.addDDragAccDParam(SpacecraftState s,
Parameter param,
double density,
Vector3D relativeVelocity,
double[] dAccdParam)
Compute acceleration derivatives with respect to ballistic coefficient.
|
void |
AeroModel.addDDragAccDParam(SpacecraftState s,
Parameter param,
double density,
Vector3D relativeVelocity,
double[] dAccdParam)
Compute acceleration derivatives with respect to ballistic coefficient.
|
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 |
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 |
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 |
GlobalAeroModel.dragAcceleration(SpacecraftState state,
double density,
Vector3D relativeVelocity)
Method to compute the aero acceleration, based on the assembly.
|
Vector3D |
DragLiftModel.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)
Compute the torque due to radiation pressire.
|
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.
|
double[][] |
InterpolatedDragReader.readFile(String filePath)
Read the aero coefficients file by parsing it.
|
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 |
MassModel.updateMass(String partName,
double newMass)
Update the mass of the given part.
|
void |
InertiaComputedModel.updateMass(String partName,
double mass)
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.
|
| Constructor and Description |
|---|
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).
|
| Modifier and Type | Method and Description |
|---|---|
static double |
AeroCoeffByAoA.angleOfAttackFromSpacecraftState(SpacecraftState state,
EllipsoidBodyShape earthShape)
Computes the angle of attack from the spacecraft state and the Earth shape.
|
protected abstract double |
AbstractAeroCoeff1D.computeXVariable(SpacecraftState state)
Computes the x variable from the spacecraft state.
|
protected double |
AeroCoeffByAoA.computeXVariable(SpacecraftState state)
Computes the x variable from the spacecraft state.
|
protected double |
AeroCoeffByMach.computeXVariable(SpacecraftState state)
Computes the x variable from the spacecraft state.
|
static double |
AeroCoeffByMach.machFromSpacecraftState(SpacecraftState state,
Atmosphere atmosphere)
Computes the Mach number from the spacecraft state and an atmosphere model.
|
| Modifier and Type | Method and Description |
|---|---|
void |
MassEquation.computeDerivatives(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the derivatives related to the additional state parameters.
|
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 |
RadiativeSphereProperty.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 |
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 |
RadiativeSphereProperty.getSphereRadius()
Get the sphere radius.
|
double |
AeroSphereProperty.getSphereRadius()
Get the sphere radius.
|
void |
MassProperty.updateMass(double newMass)
Updates the mass of the part.
|
| Constructor and Description |
|---|
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. |
TankProperty(double massIn)
Constructor.
|
| Modifier and Type | Method and Description |
|---|---|
double |
AerodynamicProperties.getConstantDragCoef()
Get the drag coefficient.
|
double |
AerodynamicProperties.getConstantLiftCoef()
Get the lift coefficient.
|
AerodynamicCoefficientType |
AerodynamicProperties.getFunctionType()
Get the type of the aerodynamic coefficients functions among:
- Coefficients as a function of altitude, - Coefficients as a function of angle of attack - Coefficients as a function of Mach number - Coefficients as a function of Mach number and angle of attack. |
void |
RadiativeProperties.setRadiativeProperties(AssemblyBuilder builder,
String mainPartName,
double multiplicativeFactor)
Set radiative properties.
|
| Constructor and Description |
|---|
AerodynamicProperties(Sphere sphere,
AerodynamicCoefficient dragCoefIn,
AerodynamicCoefficient liftCoefIn)
Constructor.
|
AerodynamicProperties(Sphere sphere,
double dragCoefIn)
Constructor.
|
AerodynamicProperties(VehicleSurfaceModel vehicleSurface,
double dragCoefIn,
double liftCoefIn)
Constructor.
|
RadiativeProperties(RadiativeProperty radiativePropertyIn,
RadiativeIRProperty radiativeIRPropertyIn,
VehicleSurfaceModel vehicleSurfaceModelIn)
Constructor.
|
VehicleSurfaceModel(CrossSectionProvider vehicleShape)
Constructor without solar panels and with default multiplicative factor set to 1.0.
|
VehicleSurfaceModel(CrossSectionProvider vehicleShape,
RightParallelepiped solarPanels)
Constructor with default multiplicative factor set to 1.0.
|
VehicleSurfaceModel(CrossSectionProvider mainPart,
RightParallelepiped solarPanels,
double multiplicativeFactorIn)
Constructor.
|
| Modifier and Type | Method and Description |
|---|---|
static double[] |
AeroAttitudeLaw.aircraftToAero(double slopeVel,
double azimuthVel,
double yaw,
double pitch,
double roll)
Method to compute aerodynamic frame orientation angles with respect aircraft frame.
|
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 |
TwoSpinBiasSlew.compute(PVCoordinatesProvider pvProv)
Compute the slew corresponding to an orbital state.
|
void |
ConstantSpinSlew.compute(PVCoordinatesProvider pvProv)
Compute the slew corresponding to an orbital state.
|
void |
Slew.compute(PVCoordinatesProvider pvProv)
Compute the slew corresponding to an orbital state.
|
double |
TwoSpinBiasSlew.computeDuration(PVCoordinatesProvider pvProv)
Computes the actual slew duration.
|
double |
AbstractIsisSpinBiasSlew.computeDuration(PVCoordinatesProvider pvProv)
Computes the slew duration.
|
protected double |
VariableStepAttitudeEphemerisGenerator.computeStep(AbsoluteDate date,
AbsoluteDateInterval ephemerisInterval)
Computes the step used during the variable step ephemeris generation.
|
protected double |
FixedStepAttitudeEphemerisGenerator.computeStep(AbsoluteDate date,
AbsoluteDateInterval ephemerisInterval)
Computes the step used during attitude ephemeris generation.
|
protected abstract double |
AbstractAttitudeEphemerisGenerator.computeStep(AbsoluteDate date,
AbsoluteDateInterval ephemerisInterval)
Computes the step used during attitude 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 |
TwoSpinBiasSlew.getAttitude(AbsoluteDate date,
Frame frame)
Compute the attitude.
|
Attitude |
ConstantSpinSlew.getAttitude(AbsoluteDate date,
Frame frame)
Compute the attitude.
|
Attitude |
Slew.getAttitude(AbsoluteDate date,
Frame frame)
Compute the attitude.
|
Attitude |
AbstractIsisSpinBiasSlew.getAttitude(AbsoluteDate date,
Frame frame)
Compute the attitude.
|
Attitude |
AbstractSlew.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 |
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 |
TabulatedAttitude.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 |
AttitudeLawLeg.getAttitude(Orbit orbit)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AbstractAttitudeLaw.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 |
AttitudeLegsSequence.getAttitude(Orbit orbit)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AbstractSlew.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Warning: provided
PVCoordinatesProvider is here not used. |
Attitude |
LofOffsetPointing.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
YawSteering.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AttitudeProvider.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 |
AttitudeLegLaw.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
ComposedAttitudeLaw.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
TabulatedAttitude.getAttitude(PVCoordinatesProvider pvProvider,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AttitudesSequence.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
LofOffset.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
FixedRate.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AeroAttitudeLaw.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 |
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)
Compute the attitude corresponding to an orbital state.
|
Attitude |
BodyCenterPointing.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 |
YawCompensation.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 |
AbstractGroundPointing.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 |
CelestialBodyPointed.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
RelativeTabulatedAttitudeLeg.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.
|
Attitude |
AbstractGroundPointingWrapper.getBaseState(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the base system state at given date, without compensation.
|
TimeStampedAngularCoordinates |
YawSteering.getCompensation(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame orbitFrame,
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.
|
abstract TimeStampedAngularCoordinates |
AbstractGroundPointingWrapper.getCompensation(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame,
Attitude base)
Compute the TimeStampedAngularCoordinates at a given time.
|
double |
AbstractSlew.getDuration() |
Rotation |
DirectionTrackingOrientation.getOrientation(AbsoluteDate date,
Frame frame)
Gets the rotation defining the orientation with respect to a given frame at a given date.
|
Rotation |
IOrientationLaw.getOrientation(AbsoluteDate date,
Frame frame)
Gets the rotation defining the orientation with respect to a given frame at a given date.
|
Attitude |
AttitudeLegsSequence.getPreviousAttitude(PVCoordinatesProvider provider,
AbsoluteDate date,
Frame frame)
Returns attitude from previous leg (compared to leg matching provided date) from the
sequence.
|
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 |
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 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)
Compute the target point in specified frame.
|
protected abstract Vector3D |
AbstractGroundPointing.getTargetPoint(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point in specified frame.
|
protected Vector3D |
AbstractGroundPointingWrapper.getTargetPoint(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point in specified frame.
|
TimeStampedPVCoordinates |
NadirPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
protected TimeStampedPVCoordinates |
AbstractGroundPointing.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
protected TimeStampedPVCoordinates |
AbstractGroundPointingWrapper.getTargetPV(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Compute the target point position/velocity in specified frame.
|
AbsoluteDateInterval |
AbstractSlew.getTimeInterval()
Return the time interval of validity of the leg
|
AbsoluteDateInterval |
RelativeTabulatedAttitudeLeg.getTimeInterval()
Return the time interval of validity of the leg
|
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 |
LofOffsetPointing.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AttitudeProvider.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 |
AttitudesSequence.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AttitudeLawLeg.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AbstractAttitudeLaw.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
RelativeTabulatedAttitudeLeg.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AttitudeLegsSequence.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AbstractGroundPointingWrapper.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.
|
| Constructor and Description |
|---|
AeroAttitudeLaw(double angleofattack,
double sideslip,
double rollVel,
EllipsoidBodyShape earthShape)
Constructor.
|
AeroAttitudeLaw(double angleofattack,
double sideslip,
double rollVel,
EllipsoidBodyShape earthShape,
double dtSpin,
double dtAcc)
Constructor with parameterizable delta-time for spin and acceleration computation.
|
AeroAttitudeLaw(IParameterizableFunction angleofattack,
IParameterizableFunction sideslip,
IParameterizableFunction rollVel,
EllipsoidBodyShape earthShape)
Constructor.
|
AeroAttitudeLaw(IParameterizableFunction angleofattack,
IParameterizableFunction sideslip,
IParameterizableFunction rollVel,
EllipsoidBodyShape 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 inertialFrameIn,
LOFType typeIn)
Create a LOF-aligned attitude.
|
LofOffset(Frame pInertialFrame,
LOFType typeIn,
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,
Frame frame,
List<Pair<Double,AngularCoordinates>> angularCoordinates,
String natureIn)
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,AngularCoordinates>> angularCoordinates,
int nbInterpolationPoints,
Frame frame,
String natureIn)
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.
|
RelativeTabulatedAttitudeLeg(AbsoluteDate referenceDate,
List<Pair<Double,Rotation>> orientations,
Frame frame,
int nbInterpolationPoints,
String natureIn)
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.
|
RelativeTabulatedAttitudeLeg(AbsoluteDate referenceDate,
List<Pair<Double,Rotation>> orientations,
Frame frame,
String natureIn)
Build a RelativeTabulatedAttitudeLeg with a reference date, a list of Rotations associated
with a double representing the time elapsed since the reference date.
|
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
|
TabulatedAttitude(List<Attitude> inAttitudes,
int nbInterpolationPoints,
String natureIn)
Constructor with number of points used for interpolation
|
TabulatedAttitude(List<Attitude> inAttitudes,
String natureIn)
Constructor with default number N of points used for interpolation.
|
TargetGroundPointing(BodyShape shape,
Vector3D targetIn)
Creates a new instance from body shape and target expressed in cartesian coordinates.
|
TargetGroundPointing(BodyShape shape,
Vector3D targetIn,
Vector3D losInSatFrameVec,
Vector3D losNormalInSatFrameVec)
Creates a new instance from body shape and target expressed in cartesian coordinates with
specified los axis in satellite frame.
|
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.
|
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,
String natureIn)
This class extends the AbstractSlew.
|
| Modifier and Type | Method and Description |
|---|---|
Vector3D |
GlintApproximatePointingDirection.getGlintVectorPosition(PVCoordinatesProvider origin,
AbsoluteDate date,
Frame frame)
Get the position vector of the glint point
|
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 |
EarthCenterDirection.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 |
ConstantVectorDirection.getLine(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Provides the line containing the given origin point 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.
|
Line |
NadirDirection.getLine(PVCoordinatesProvider pvCoord,
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 |
IDirection.getLine(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
|
Line |
GroundVelocityDirection.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 |
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 |
CelestialBodyPolesAxisDirection.getLine(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Provides the line containing the origin (given PV coordinates) and directed by the direction vector.
|
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
|
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 |
ITargetDirection.getTargetPVCoordinates(AbsoluteDate date,
Frame frame)
Provides the target point at a given date in a given frame, represented by the
associated PVCoordinates object
|
Vector3D |
MomentumDirection.getVector(PVCoordinatesProvider origin,
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 |
VelocityDirection.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 |
ToCelestialBodyCenterDirection.getVector(PVCoordinatesProvider origin,
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)
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 |
IDirection.getVector(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Provides the direction vector at a given date in a given frame.
|
Vector3D |
GroundVelocityDirection.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 |
CrossProductDirection.getVector(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Provides the cross product of direction1 vector and dirction2 vector.
|
Vector3D |
CelestialBodyPolesAxisDirection.getVector(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Provides the direction vector at a given date in a given frame.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
Attitude |
MultiAttitudeProviderWrapper.getAttitude(Map<String,Orbit> orbits)
Computes the attitude corresponding to several orbital states.
|
Attitude |
MultiAttitudeProvider.getAttitude(Map<String,Orbit> orbits)
Computes the attitude corresponding to several orbital states.
|
Attitude |
MultiAttitudeProviderWrapper.getAttitude(Map<String,PVCoordinatesProvider> pvProvs,
AbsoluteDate date,
Frame frame)
Computes the attitude corresponding to several orbital states.
|
Attitude |
MultiAttitudeProvider.getAttitude(Map<String,PVCoordinatesProvider> pvProvs,
AbsoluteDate date,
Frame frame)
Computes the attitude corresponding to several orbital states.
|
| Modifier and Type | Method and Description |
|---|---|
double |
OrientationAngleProvider.getOrientationAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date)
Compute the orientation angle corresponding to an orbital state.
|
double |
ConstantOrientationAngleLeg.getOrientationAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date)
Compute the orientation angle corresponding to an orbital state.
|
double |
ConstantOrientationAngleLaw.getOrientationAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date)
Compute the orientation angle corresponding to an orbital state.
|
double |
OrientationAngleProfileSequence.getOrientationAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date)
Compute the orientation angle corresponding to an orbital state.
|
double |
OrientationAngleLawLeg.getOrientationAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date)
Compute the orientation angle corresponding to an orbital state.
|
double |
OrientationAngleLegsSequence.getOrientationAngle(PVCoordinatesProvider pvProv,
AbsoluteDate date)
Compute the orientation angle corresponding to an orbital state.
|
AbsoluteDateInterval |
AbstractOrientationAngleLeg.getTimeInterval()
Return the time interval of validity of the leg
|
| Modifier and Type | Method and Description |
|---|---|
void |
AbstractAttitudeProfile.checkDate(AbsoluteDate userDate)
Check date validity.
|
Attitude |
AttitudeProfilesSequence.getAttitude(Orbit orbit)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AbstractAttitudeProfile.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 |
AngularVelocitiesPolynomialProfile.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate userDate,
Frame frame)
Compute the attitude corresponding to an orbital state.
|
Attitude |
AttitudeProfilesSequence.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate date,
Frame frame)
Gets the attitude from the sequence.
|
Attitude |
AngularVelocitiesHarmonicProfile.getAttitude(PVCoordinatesProvider pvProv,
AbsoluteDate userDate,
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.
|
Vector3D |
AngularVelocitiesHarmonicProfile.getRotationAcceleration(AbsoluteDate dateIn)
Get the rotation acceleration from the vector3D at a given date.
|
void |
QuaternionHarmonicProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AngularVelocitiesPolynomialProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AttitudeProfilesSequence.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
AngularVelocitiesHarmonicProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
void |
QuaternionPolynomialProfile.setSpinDerivativesComputation(boolean computeSpinDerivatives)
Method to activate spin derivative computation.
|
AngularVelocitiesPolynomialSegment |
AngularVelocitiesPolynomialSegment.truncateSegment(AbsoluteDateInterval newInterval)
Given a sub-interval, truncate the original segment.
|
AngularVelocitiesPolynomialProfile |
AngularVelocitiesPolynomialProfile.truncateTimeInterval(AbsoluteDateInterval newInterval)
Truncate the validity interval of the profile
|
AngularVelocitiesHarmonicProfile |
AngularVelocitiesHarmonicProfile.truncateTimeInterval(AbsoluteDateInterval newInterval)
Truncate the validity interval of the profile
|
| Modifier and Type | Method and Description |
|---|---|
static void |
CelestialBodyFactory.addDefaultCelestialBodyLoader(String supportedNames)
Add the default loaders for all predefined celestial bodies.
|
static void |
CelestialBodyFactory.addDefaultCelestialBodyLoader(String name,
String supportedNames)
Add the default loaders for celestial bodies.
|
double |
GeometricBodyShape.distanceTo(Line line,
Frame frame,
AbsoluteDate date)
Computes the distance to a line.
|
double |
ExtendedOneAxisEllipsoid.distanceTo(Line line,
Frame frame,
AbsoluteDate date)
Computes the distance to a line.
|
double |
ExtendedOneAxisEllipsoid.getApparentRadius(Vector3D position,
Frame frame,
AbsoluteDate date,
PVCoordinatesProvider occultedBody)
Calculate the apparent radius.
|
double |
EllipsoidBodyShape.getApparentRadius(Vector3D position,
Frame frame,
AbsoluteDate date,
PVCoordinatesProvider occultedBody)
Calculate the apparent radius.
|
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 |
OneAxisEllipsoid.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)
Get the intersection point of a line with the surface of the body.
|
GeodeticPoint |
BodyShape.getIntersectionPoint(Line line,
Vector3D close,
Frame frame,
AbsoluteDate date)
Get the intersection point of a line with the surface of the body.
|
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)
Compute the intersection points with a line.
|
static CelestialBody |
CelestialBodyFactory.getJupiter()
Get the Jupiter singleton body.
|
Line |
BasicBoardSun.getLine(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Get the line from the position in pvCoord to the Sun.
|
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 |
MeeusSun.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
PVCoordinates |
UserCelestialBody.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
abstract PVCoordinates |
AbstractCelestialBody.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
PVCoordinates |
ExtendedOneAxisEllipsoid.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
PVCoordinates |
MeeusMoon.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.
|
Vector3D |
BasicBoardSun.getVector(PVCoordinatesProvider pvCoord,
AbsoluteDate date,
Frame frame)
Get the direction of the sun.
|
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 |
OneAxisEllipsoid.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)
Transform a cartesian point to a surface-relative point.
|
GeodeticPoint |
BodyShape.transform(Vector3D point,
Frame frame,
AbsoluteDate date)
Transform a cartesian point to a surface-relative point.
|
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. |
| Constructor and Description |
|---|
JPLEphemeridesLoader(String supportedNamesIn,
JPLEphemeridesLoader.EphemerisType generateTypeIn)
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.
|
| Modifier and Type | Method and Description |
|---|---|
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 |
ClasspathCrawler.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 |
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.
|
| Constructor and Description |
|---|
ClasspathCrawler(ClassLoader classLoaderIn,
String... list)
Build a data classpath crawler.
|
ClasspathCrawler(String... list)
Build a data classpath crawler.
|
DirectoryCrawler(File rootIn)
Build a data files crawler.
|
PoissonSeries(InputStream stream,
double factor,
String name)
Build a Poisson series from an IERS table file.
|
ZipJarCrawler(ClassLoader classLoaderIn,
String resourceIn)
Build a zip crawler for an archive file in classpath.
|
ZipJarCrawler(String resourceIn)
Build a zip crawler for an archive file in classpath.
|
ZipJarCrawler(URL urlIn)
Build a zip crawler for an archive file on network.
|
| Modifier and Type | Method and Description |
|---|---|
Map<CodingEventDetector,PhenomenaList> |
CodedEventsLogger.buildPhenomenaListMap(AbsoluteDateInterval definitionInterval,
SpacecraftState duringState)
Builds a map of
PhenomenaList, one list per CodingEventDetector instance. |
EventDetector.Action |
CombinedPhenomenaDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
EarthZoneDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle the event and choose what to do next.
|
EventDetector.Action |
GenericCodingEventDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
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 |
CombinedPhenomenaDetector.g(SpacecraftState s)
Compute the value of the switching function for a combination (AND or OR) of two phenomena.
|
double |
EarthZoneDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
GenericCodingEventDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
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)
Reset the state (including additional states) prior to continue propagation.
|
| Modifier and Type | Method and Description |
|---|---|
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)
Handle an event and choose what to do next.
|
double |
MultiGenericCodingEventDetector.g(Map<String,SpacecraftState> s)
Compute the value of the switching function.
|
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)
Reset the states (including additional states) prior to continue propagation.
|
| Constructor and Description |
|---|
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. |
| Modifier and Type | Method and Description |
|---|---|
EventDetector.Action |
SatToSatMutualVisibilityDetector.eventOccurred(Map<String,SpacecraftState> s,
boolean increasing,
boolean forward)
Handle an 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 |
StationToSatMutualVisibilityDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
SatToSatMutualVisibilityDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
SensorInhibitionDetector.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)
Handle an event and choose what to do next.
|
EventDetector.Action |
TargetInFieldOfViewDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
RFVisibilityDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
VisibilityFromStationDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle "visibility from station" event and choose what to do next.
|
EventDetector.Action |
MaskingDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle "masking" event and choose what to do next.
|
double |
SatToSatMutualVisibilityDetector.g(Map<String,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 |
StationToSatMutualVisibilityDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
SatToSatMutualVisibilityDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
SensorInhibitionDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
SensorVisibilityDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
TargetInFieldOfViewDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
RFVisibilityDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
VisibilityFromStationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
MaskingDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
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)
Reset the states (including additional states) prior to continue propagation.
|
void |
SecondarySpacecraft.updateSpacecraftState(AbsoluteDate date)
Updates the assembly frames at a given date from the orbit and attitude
information provided by the propagator.
|
| Modifier and Type | Method and Description |
|---|---|
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. |
| Modifier and Type | Method and Description |
|---|---|
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. |
| Modifier and Type | Method and Description |
|---|---|
void |
EmpiricalForce.addContribution(SpacecraftState state,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
ForceModel.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
EmpiricalForce.addDAccDParam(SpacecraftState state,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters.
|
void |
EmpiricalForce.addDAccDState(SpacecraftState state,
double[][] dAccdPos,
double[][] dAccdVel)
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 |
EmpiricalForce.computeAcceleration(SpacecraftState state)
Compute the acceleration due to the force.
|
Vector3D |
ForceModel.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
void |
ForceModelsData.updateAssembly(Assembly assembly)
Method to update the force models depending on the assembly (DragForce,
CustomSolarRadiationPressureEllipsoidCircular and CustomRediffusedRadiationPressure).
|
| Modifier and Type | Method and Description |
|---|---|
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 |
MSISE2000.getData(AbsoluteDate date,
Vector3D position,
Frame frame)
Get detailed atmospheric data.
|
AtmosphereData |
ExtendedAtmosphere.getData(AbsoluteDate date,
Vector3D position,
Frame frame)
Get detailed atmospheric data.
|
double |
HarrisPriester.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
DTM2000.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
US76.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density for altitude in interval [0, 1E6] m
Note: if altitude < 0 m or altitude > 1E6 m the density corresponding to the closest bound is returned.
|
double |
Atmosphere.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 |
JB2006.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
SimpleExponentialAtmosphere.getDensity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local density.
|
double |
HarrisPriester.getDensity(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 |
DTM2000InputParameters.getInstantFlux(AbsoluteDate date)
Get the value of the instantaneous solar flux.
|
double |
MSISE2000InputParameters.getInstantFlux(AbsoluteDate date)
Get the value of the instantaneous solar flux.
|
double |
DTM2000InputParameters.getMeanFlux(AbsoluteDate date)
Get the value of the mean solar flux.
|
double |
MSISE2000InputParameters.getMeanFlux(AbsoluteDate date)
Get the 81 day average of F10.7 flux.
|
double |
US76.getPress(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local pressure for altitude in interval [0, 1E6] m
Note: if altitude < 0 m or altitude > 1E6 m the pressure corresponding to the closest bound is returned.
|
double |
MSISE2000.getPressure(AbsoluteDate date,
Vector3D position,
Frame frame)
Returns pressure.
|
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 |
HarrisPriester.getSpeedOfSound(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local speed of sound.
|
double |
DTM2000.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 |
Atmosphere.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 |
JB2006.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 |
US76.getTemp(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local temperature for altitude in interval [0, 1E6] m
Note: if altitude < 0 m or altitude > 1E6 m the temperature corresponding to the closest bound is returned.
|
double |
DTM2000InputParameters.getThreeHourlyKP(AbsoluteDate date)
Get the value of the 3 hours geomagnetic index.
|
Vector3D |
HarrisPriester.getVelocity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the inertial velocity of atmosphere molecules.
|
Vector3D |
DTM2000.getVelocity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the inertial velocity of atmosphere molecules.
|
Vector3D |
US76.getVelocity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the spacecraft velocity relative to the atmosphere.
|
Vector3D |
Atmosphere.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 |
JB2006.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.
|
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.
|
| Constructor and Description |
|---|
DTM2000(DTM2000InputParameters parameters,
PVCoordinatesProvider sunIn,
BodyShape earthIn)
Simple constructor for independent computation.
|
| Modifier and Type | Method and Description |
|---|---|
static void |
SolarActivityDataFactory.addDefaultSolarActivityDataReaders()
Add the default READERS for solar activity
The default READERS supports ACSOL format with the default name
SolarActivityDataFactory.ACSOL_FILENAME |
double |
SolarActivityDataProvider.getAp(AbsoluteDate date)
Get Ap value at given user date
|
double |
ExtendedSolarActivityWrapper.getAp(AbsoluteDate date)
Get Ap value at given user date
|
SortedMap<AbsoluteDate,Double[]> |
SolarActivityDataProvider.getApKpValues(AbsoluteDate date1,
AbsoluteDate date2)
Get ap / kp values between the given dates
|
SortedMap<AbsoluteDate,Double[]> |
ExtendedSolarActivityWrapper.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 |
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)
|
double |
ExtendedSolarActivityWrapper.getInstantFluxValue(AbsoluteDate date)
Get instant flux values at the given dates (possibly interpolated)
|
SortedMap<AbsoluteDate,Double> |
SolarActivityDataProvider.getInstantFluxValues(AbsoluteDate date1,
AbsoluteDate date2)
Get raw instant flux values between the given dates
|
SortedMap<AbsoluteDate,Double> |
ExtendedSolarActivityWrapper.getInstantFluxValues(AbsoluteDate date1,
AbsoluteDate date2)
Get raw instant flux values between the given dates
|
double |
SolarActivityDataProvider.getKp(AbsoluteDate date)
Get Kp value at given user date
|
double |
ExtendedSolarActivityWrapper.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.
|
| Constructor and Description |
|---|
ACSOLFormatReader(String supportedNames)
Constructor.
|
NOAAFormatReader(String supportedNames)
Constructor.
|
| Modifier and Type | Method and Description |
|---|---|
double |
MarshallSolarActivityFutureEstimation.get24HoursKp(AbsoluteDate date)
The Kp index is derived from the Ap index.
|
double |
DTM2000SolarData.get24HoursKp(AbsoluteDate date)
Get the last 24H mean geomagnetic index.
|
double[] |
ClassicalMSISE2000SolarData.getApValues(AbsoluteDate date)
Get the array containing the 7 ap values
|
double[] |
ContinuousMSISE2000SolarData.getApValues(AbsoluteDate date)
Get the array containing the 7 ap values
|
abstract double[] |
AbstractMSISE2000SolarData.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 |
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.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.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 |
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.getThreeHourlyKP(AbsoluteDate date)
Get the value of the 3 hours geomagnetic index.
|
double |
DTM2000SolarData.getThreeHourlyKP(AbsoluteDate date)
Get the value of the 3 hours geomagnetic index.
|
void |
MarshallSolarActivityFutureEstimation.loadData(InputStream input,
String name)
Load data from a stream.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Constructor and Description |
|---|
DragForce(double kIn,
Atmosphere atmosphereIn,
Assembly assembly)
Creates a new instance.
|
DragForce(DragForce otherDragForce,
Assembly assembly)
Creates a new instance from the data in another one but with a different assembly.
|
| Modifier and Type | Method and Description |
|---|---|
void |
BalminoAttractionModel.addContribution(SpacecraftState state,
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 state,
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 |
DrozinerAttractionModel.addContribution(SpacecraftState state,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
BalminoAttractionModel.addDAccDParam(SpacecraftState state,
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 state,
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 state,
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 state,
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 state)
Compute the acceleration due to the force.
|
Vector3D |
ThirdBodyAttraction.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
CunninghamAttractionModel.computeAcceleration(SpacecraftState state)
Compute the acceleration due to the force.
|
Vector3D |
NewtonianAttraction.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
DrozinerAttractionModel.computeAcceleration(SpacecraftState state)
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,
boolean missingCoefficientsAllowed)
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,
boolean missingCoefficientsAllowed)
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,
boolean missingCoefficientsAllowed)
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.
|
static ForceModel |
EarthGravitationalModelFactory.getGravitationalModel(EarthGravitationalModelFactory.GravityFieldNames potentialFileName,
String filename,
int n,
int m,
boolean missingCoefficientsAllowed)
Create an default instance of a gravitational field of a celestial body using Balmino model and specific data.
|
| Modifier and Type | Method and Description |
|---|---|
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 |
EGMFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
SHMFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
abstract void |
PotentialCoefficientsReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
ICGEMFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
GRGSFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
| Modifier and Type | Method and Description |
|---|---|
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
|
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.
|
void |
OceanTides.updateCoefficientsCandS(AbsoluteDate date)
Update the C and the S coefficients for acceleration computation.
|
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 |
TerrestrialTides.updateCoefficientsCandS(AbsoluteDate date)
Update the C and the S coefficients for acceleration computation.
|
void |
OceanTides.updateCoefficientsCandSPD(AbsoluteDate date)
Update the C and the S coefficients for partial derivatives 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 |
TerrestrialTides.updateCoefficientsCandSPD(AbsoluteDate date)
Update the C and the S coefficients for partial derivatives computation.
|
| Constructor and Description |
|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
static OceanTidesCoefficientsProvider |
OceanTidesCoefficientsFactory.getCoefficientsProvider()
Get the ocean tides coefficients provider from the first supported file.
|
abstract void |
OceanTidesCoefficientsReader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
FES2004FormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
| Modifier and Type | Method and Description |
|---|---|
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)
Compute acceleration derivatives with respect to state parameters.
|
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.
|
| Constructor and Description |
|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
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
|
| Modifier and Type | Method and Description |
|---|---|
void |
ConstantThrustError.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the constant thrust error model to the perturbing acceleration.
|
void |
ContinuousThrustManeuver.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
ConstantThrustError.addDAccDParam(SpacecraftState state,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters.
|
void |
ContinuousThrustManeuver.addDAccDParam(SpacecraftState s,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters.
|
void |
ConstantThrustError.addDAccDState(SpacecraftState state,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters.
|
void |
ContinuousThrustManeuver.addDAccDState(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters.
|
SpacecraftState |
SmallManeuverAnalyticalModel.apply(SpacecraftState state1)
Compute the effect of the maneuver on a spacecraft state.
|
Vector3D |
ConstantThrustError.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
ContinuousThrustManeuver.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.
|
| Constructor and Description |
|---|
ContinuousThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
IDependentVectorVariable<SpacecraftState> inDirection,
MassProvider massProvider,
TankProperty tank)
Constructor for a variable direction in satellite frame.
|
ContinuousThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
IDependentVectorVariable<SpacecraftState> inDirection,
MassProvider massProvider,
TankProperty tank,
Frame frameIn)
Constructor for a variable direction in provided frame.
|
ContinuousThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
IDependentVectorVariable<SpacecraftState> inDirection,
MassProvider massProvider,
TankProperty tank,
LOFType lofTyp)
Constructor for a variable direction in local orbital frame.
|
ContinuousThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank)
Constructor for a constant direction in satellite frame.
|
ContinuousThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank,
Frame frameIn)
Constructor for a constant direction in provided frame.
|
ContinuousThrustManeuver(EventDetector startEventDetector,
EventDetector stopEventDetector,
PropulsiveProperty engine,
Vector3D inDirection,
MassProvider massProvider,
TankProperty tank,
LOFType lofTyp)
Constructor for a constant direction in local orbital frame.
|
ImpulseManeuver(EventDetector inTrigger,
Vector3D inDeltaVSat,
double isp,
MassProvider massModel,
String part)
Build a new instance.
|
ImpulseManeuver(EventDetector inTrigger,
Vector3D inDeltaVSat,
double isp,
MassProvider massModel,
String part,
LOFType inLofType)
Build a new instance with a LocalOrbitalFrame.
|
ImpulseManeuver(EventDetector inTrigger,
Vector3D inDeltaVSat,
Frame inFrame,
double isp,
MassProvider massModel,
String part)
Build a new instance.
|
SmallManeuverAnalyticalModel(SpacecraftState state0In,
Frame frame,
Vector3D dV,
double isp,
String partNameIn)
Build a maneuver defined in user-specified frame.
|
SmallManeuverAnalyticalModel(SpacecraftState state0In,
Vector3D dV,
double isp,
String partNameIn)
Build a maneuver defined in spacecraft frame.
|
| Modifier and Type | Method and Description |
|---|---|
void |
SolarRadiationPressureCircular.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
RediffusedRadiationPressure.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
SolarRadiationPressureEllipsoid.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
SolarRadiationPressureCircular.addDAccDParam(SpacecraftState s,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters.
|
void |
RediffusedRadiationPressure.addDAccDParam(SpacecraftState s,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters.
|
void |
SolarRadiationPressureEllipsoid.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 |
SolarRadiationPressureCircular.addDAccDState(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters.
|
void |
RediffusedRadiationPressure.addDAccDState(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters.
|
void |
SolarRadiationPressureEllipsoid.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 |
SolarRadiationPressureCircular.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
RediffusedRadiationPressure.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
SolarRadiationPressureEllipsoid.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
static double |
SolarRadiationPressureEllipsoid.getLightningRatio(PVCoordinatesProvider sun,
Vector3D satSunVector,
GeometricBodyShape earthModel,
Vector3D position,
Frame frame,
AbsoluteDate date)
Get the lightning ratio ([0-1]).
|
double |
SolarRadiationPressureCircular.getLightningRatio(Vector3D position,
Frame frame,
AbsoluteDate date)
Get the lightning ratio ([0-1]).
|
Vector3D |
SolarRadiationPressureEllipsoid.getSolarFlux(SpacecraftState s)
Compute solar flux.
|
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
|
| Constructor and Description |
|---|
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.
|
RediffusedRadiationPressure(RediffusedRadiationPressure otherInstance,
Assembly assembly)
Creates a new instance from the data in another one but with a different assembly.
|
| Modifier and Type | Method and Description |
|---|---|
void |
CoriolisRelativisticEffect.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 |
SchwarzschildRelativisticEffect.addContribution(SpacecraftState s,
TimeDerivativesEquations adder)
Compute the contribution of the force model to the perturbing
acceleration.
|
void |
CoriolisRelativisticEffect.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 |
SchwarzschildRelativisticEffect.addDAccDParam(SpacecraftState s,
Parameter param,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters.
|
void |
CoriolisRelativisticEffect.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 |
SchwarzschildRelativisticEffect.addDAccDState(SpacecraftState s,
double[][] dAccdPos,
double[][] dAccdVel)
Compute acceleration derivatives with respect to state parameters.
|
Vector3D |
CoriolisRelativisticEffect.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
LenseThirringRelativisticEffect.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
Vector3D |
SchwarzschildRelativisticEffect.computeAcceleration(SpacecraftState s)
Compute the acceleration due to the force.
|
| Modifier and Type | Method and Description |
|---|---|
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 |
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.
|
| Constructor and Description |
|---|
H0MinusNFrame(String name,
AbsoluteDate h0In,
double nIn,
double longitudeIn)
Constructor.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
void |
AbstractEOPHistory.checkEOPContinuity(double maxGap)
Check Earth orientation parameters continuity.
|
void |
EOP05C04FilesLoader.fillHistory(EOP1980History history)
Load celestial body.
|
void |
EOP08C04FilesLoader.fillHistory(EOP1980History history)
Load celestial body.
|
void |
RapidDataAndPredictionColumnsLoader.fillHistory(EOP1980History history)
Load celestial body.
|
void |
EOP1980HistoryLoader.fillHistory(EOP1980History history)
Load celestial body.
|
void |
RapidDataAndPredictionXMLLoader.fillHistory(EOP1980History history)
Load celestial body.
|
void |
NoEOP1980HistoryLoader.fillHistory(EOP1980History history)
History with zero orientation.
|
void |
BulletinBFilesLoader.fillHistory(EOP1980History history)
Load celestial body.
|
static void |
EOP05C04FilesLoader.fillHistory(EOP1980History history,
InputStream istream)
Fills the history object directy with data from the
InputStream, bypassing the Orekit data loaders. |
static void |
EOP08C04FilesLoader.fillHistory(EOP1980History history,
InputStream istream)
Fills the history object directy with data from the
InputStream, bypassing the Orekit data loaders. |
void |
EOP05C04FilesLoader.fillHistory(EOP2000History history)
Load celestial body.
|
void |
EOP2000HistoryLoader.fillHistory(EOP2000History history)
Load celestial body.
|
void |
EOP08C04FilesLoader.fillHistory(EOP2000History history)
Load celestial body.
|
void |
RapidDataAndPredictionColumnsLoader.fillHistory(EOP2000History history)
Load celestial body.
|
void |
RapidDataAndPredictionXMLLoader.fillHistory(EOP2000History history)
Load celestial body.
|
void |
BulletinBFilesLoader.fillHistory(EOP2000History history)
Load celestial body.
|
static void |
EOP05C04FilesLoader.fillHistory(EOP2000History history,
InputStream istream)
Fills the history object directy with data from the
InputStream, bypassing the Orekit data loaders. |
static void |
EOP08C04FilesLoader.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 |
EOP05C04FilesLoader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
EOP08C04FilesLoader.loadData(InputStream input,
String name)
Load data from a stream.
|
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 |
NoEOP1980HistoryLoader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
BulletinBFilesLoader.loadData(InputStream input,
String name)
Load data from a stream.
|
| Constructor and Description |
|---|
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 dtIn,
double lodIn,
double xIn,
double yIn,
double dxIn,
double dyIn)
Constructor with an AbsoluteDate parameter.
|
EOPEntry(AbsoluteDate adate,
double dtIn,
double lodIn,
double xIn,
double yIn,
double dxIn,
double dyIn,
EOPEntry.DtType type)
Constructor with an AbsoluteDate parameter.
|
EOPEntry(DateComponents datec,
double dtIn,
double lodIn,
double xIn,
double yIn,
double dxIn,
double dyIn)
Constructor with DateComponents parameter.
|
EOPEntry(DateComponents datec,
double dtIn,
double lodIn,
double xIn,
double yIn,
double dxIn,
double dyIn,
EOPEntry.DtType type)
Constructor with DateComponents parameter.
|
EOPEntry(int mjd,
double dtIn,
double lodIn,
double xIn,
double yIn,
double dxIn,
double dyIn)
Simple constructor.
|
EOPEntry(int mjd,
double dtIn,
double lodIn,
double xIn,
double yIn,
double dxIn,
double dyIn,
EOPEntry.DtType type)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
PoleCorrection |
IERS2010LibrationCorrection.getPoleCorrection(AbsoluteDate date)
This method provides the diurnal lunisolar effect on polar motion in time domain.
|
PoleCorrection |
LibrationCorrectionModel.getPoleCorrection(AbsoluteDate t)
Compute the pole corrections at a given date.
|
PoleCorrection |
LibrationCorrectionPerThread.getPoleCorrection(AbsoluteDate date)
Compute the pole corrections at a given date.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
double |
GTODProvider.getGAST(AbsoluteDate date,
FramesConfiguration config)
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.getGMST(AbsoluteDate date,
FramesConfiguration config)
Get the Greenwich mean sidereal time, in radians.
|
static double |
GTODProvider.getRotationRate(AbsoluteDate date)
Get the rotation rate of the Earth.
|
static double |
GTODProvider.getRotationRate(AbsoluteDate date,
FramesConfiguration config)
Get the rotation rate of the Earth.
|
Transform |
TODProvider.getTransform(AbsoluteDate date)
Get the transform from Mean Of Date at specified date.
|
Transform |
MODProvider.getTransform(AbsoluteDate date)
Get the transfrom from parent frame.
|
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 |
TransformProvider.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
Transform |
HelmertTransformation.getTransform(AbsoluteDate date)
Compute the transform at some date.
|
Transform |
VEISProvider.getTransform(AbsoluteDate date)
Get the transform from GTOD at specified date.
|
Transform |
TIRFProvider.getTransform(AbsoluteDate date)
Get the transform from CIRF 2000 at specified date.
|
Transform |
GTODProvider.getTransform(AbsoluteDate date)
Get the transform from TOD at 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 |
InterpolatingTransformProvider.getTransform(AbsoluteDate date)
Get the
Transform corresponding to specified date. |
Transform |
CIRFProvider.getTransform(AbsoluteDate date)
Get the transform from GCRF to CIRF2000 at the specified date.
|
Transform |
TODProvider.getTransform(AbsoluteDate date,
boolean computeSpinDerivatives)
Get the transform from Mean Of Date at specified date.
|
Transform |
MODProvider.getTransform(AbsoluteDate date,
boolean computeSpinDerivatives)
Get the transfrom from parent frame.
|
Transform |
EODProvider.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 |
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 |
VEISProvider.getTransform(AbsoluteDate date,
boolean computeSpinDerivatives)
Get the transform from GTOD at specified date.
|
Transform |
TIRFProvider.getTransform(AbsoluteDate date,
boolean computeSpinDerivatives)
Get the transform from CIRF 2000 at specified date.
|
Transform |
GTODProvider.getTransform(AbsoluteDate date,
boolean computeSpinDerivatives)
Get the transform from TOD at 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 |
InterpolatingTransformProvider.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 |
CIRFProvider.getTransform(AbsoluteDate date,
boolean computeSpinDerivatives)
Get the transform from GCRF to CIRF2000 at the specified date.
|
Transform |
TODProvider.getTransform(AbsoluteDate date,
FramesConfiguration config)
Get the transform from Mean Of Date 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 |
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 |
VEISProvider.getTransform(AbsoluteDate date,
FramesConfiguration config)
Get the
Transform corresponding to specified date. |
Transform |
TIRFProvider.getTransform(AbsoluteDate date,
FramesConfiguration config)
Get the transform from CIRF 2000 at specified date.
|
Transform |
GTODProvider.getTransform(AbsoluteDate date,
FramesConfiguration config)
Get the transform from TOD at 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 |
InterpolatingTransformProvider.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 |
TODProvider.getTransform(AbsoluteDate date,
FramesConfiguration config,
boolean computeSpinDerivatives)
Get the transform from Mean Of Date at specified date.
|
Transform |
EODProvider.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 |
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 |
VEISProvider.getTransform(AbsoluteDate date,
FramesConfiguration config,
boolean computeSpinDerivatives)
Get the
Transform corresponding to specified date. |
Transform |
TIRFProvider.getTransform(AbsoluteDate date,
FramesConfiguration config,
boolean computeSpinDerivatives)
Get the transform from CIRF 2000 at specified date.
|
Transform |
GTODProvider.getTransform(AbsoluteDate date,
FramesConfiguration config,
boolean computeSpinDerivatives)
Get the transform from TOD at 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 |
InterpolatingTransformProvider.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 |
CIRFProvider.getTransform(AbsoluteDate date,
FramesConfiguration config,
boolean computeSpinDerivatives)
Get the transform from GCRF to CIRF2000 at the 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.
|
| Constructor and Description |
|---|
GTODProvider()
Simple constructor.
|
H0MinusNProvider(AbsoluteDate h0MinusN,
double longitude)
Simple constructor.
|
ITRFEquinoxProvider()
Simple constructor.
|
TODProvider(boolean applyEOPCorr)
Simple constructor.
|
VEISProvider()
Constructor for the singleton.
|
| Modifier and Type | Method and Description |
|---|---|
PVCoordinates |
GeometricStationAntenna.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the station antenna in the selected frame. |
PVCoordinates |
RFStationAntenna.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Constructor and Description |
|---|
PiecewiseFunction(ArrayList<IParamDiffFunction> flist,
ArrayList<AbsoluteDate> xlist)
Simple constructor with 2 lists (IParamDiffFunction and AbsoluteDate) where
dates list represent the connection points between functions.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
void |
InterpolationTableLoader.loadData(InputStream input,
String name)
Loads an bi-variate interpolation table from the given
InputStream. |
void |
COFFileFormatReader.loadData(InputStream input,
String name)
Load data from a stream.
|
abstract void |
GeoMagneticModelReader.loadData(InputStream input,
String name)
Load data from a stream.
|
GeoMagneticField |
GeoMagneticField.transformModel(double year)
Time transform the model coefficients from the base year of the model
using secular variation coefficients.
|
GeoMagneticField |
GeoMagneticField.transformModel(GeoMagneticField otherModel,
double year)
Time transform the model coefficients from the base year of the model
using a linear interpolation with a second model.
|
| Constructor and Description |
|---|
FixedDelayModel(String supportedName,
double heightIn)
Creates a new
FixedDelayModel instance, and loads the delay values from the given
resource via the DataProvidersManager. |
| Modifier and Type | Method and Description |
|---|---|
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. |
| Modifier and Type | Method and Description |
|---|---|
AbsoluteDate |
GalileoAlmanacParameters.getDate(int weekNumber,
double milliInWeek)
Returns GNSS date given a week number and second in the week.
|
abstract AbsoluteDate |
AlmanacParameter.getDate(int weekNumber,
double milliInWeek)
Returns GNSS date given a week number and second in the week.
|
static double[][] |
JacobianTransformationMatrix.getJacobianCartesianToSpheric(PVCoordinates pvCoordinates)
Get Jacobian for spheric coordinates transformation to cartesian
coordinates
|
static double[][] |
JacobianTransformationMatrix.getJacobianSphericToCartesian(PVCoordinates pvCoordinates)
Get Jacobian for cartesian coordinates transformation to spheric
coordinates.
|
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. |
PVCoordinates |
AlmanacPVCoordinates.getPVCoordinates(AbsoluteDate date,
Frame frame)
Geometric computation of the position to a date.
|
PVCoordinates |
PVCoordinatesProvider.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
|
FieldVector3D<DerivativeStructure> |
PVCoordinates.toDerivativeStructureVector(int order)
Transform the instance to a
FieldVector3D<DerivativeStructure>. |
| Modifier and Type | Method and Description |
|---|---|
List<GeodeticPoint> |
AbstractProjection.applyInverseTo(double[] x,
double[] y)
Inversion transformation of arrays of x and y projected coordinates.
|
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)
Inverse projection.
|
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.
|
GeodeticPoint |
Mercator.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 |
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(double lat,
double lon)
Returns Easting value and Northing value in meters from latitude and longitude 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.
|
Vector2D |
Mercator.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)
Deprecated.
Compute the bearing (azimuth) between two geodetic Points.
|
static double |
ProjectionEllipsoidUtils.computeBearing(GeodeticPoint gv1,
GeodeticPoint gv2,
EllipsoidBodyShape shape)
Compute the bearing (azimuth) between two geodetic Points.
|
double |
ProjectionEllipsoid.computeLoxodromicDistance(GeodeticPoint p1,
GeodeticPoint p2)
Deprecated.
Loxodromic distance between P1 and P2.This is the distance of constant bearing (or along a line in Mercator).
|
static double |
ProjectionEllipsoidUtils.computeLoxodromicDistance(GeodeticPoint p1,
GeodeticPoint p2,
EllipsoidBodyShape shape)
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)
Deprecated.
Compute the point coordinates from an origin point, an azimuth and a distance along the rhumb
line (Loxodrome).
|
static GeodeticPoint |
ProjectionEllipsoidUtils.computePointAlongLoxodrome(GeodeticPoint p1,
double distance,
double azimuth,
EllipsoidBodyShape shape)
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)
Deprecated.
Discretize a rhumb line into N segments, between two points.
|
static List<GeodeticPoint> |
ProjectionEllipsoidUtils.discretizeRhumbLine(GeodeticPoint from,
GeodeticPoint to,
double maxLength,
EllipsoidBodyShape shape)
Discretize a rhumb line into N segments, between two points.
|
| Modifier and Type | Method and Description |
|---|---|
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.addAttitude(Attitude newAttitude,
AttitudeEquation.AttitudeType type)
Add attitude to the additional states map.
|
SpacecraftState |
SpacecraftState.addAttitudeToAdditionalStates(AttitudeEquation.AttitudeType attitudeType)
Add attitude to the additional states map.
|
void |
MultiPropagator.addInitialState(SpacecraftState initialState,
String satId)
Add a new spacecraft state to be propagated.
|
SpacecraftState |
SpacecraftState.addMassProvider(MassProvider massProvider)
Add the values of mass parts from MassProvider to additional states map.
|
double[] |
SpacecraftState.getAdditionalState(String name)
Get one additional state.
|
Attitude |
SpacecraftState.getAttitude()
Get the default attitude : the attitude for forces computation.
|
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()
Get the attitude for events computation.
|
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()
Get the attitude for forces computation.
|
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 |
AbstractPropagator.getInitialState()
Get the propagator initial state.
|
SpacecraftState |
Propagator.getInitialState()
Get the propagator initial state.
|
Map<String,SpacecraftState> |
MultiPropagator.getInitialStates()
Get the propagator initial states.
|
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 |
AbstractPropagator.setOrbitFrame(Frame frame)
Set propagation frame.
|
void |
Propagator.setOrbitFrame(Frame frame)
Set propagation frame.
|
SpacecraftState |
SpacecraftState.shiftedBy(double dt)
Get a time-shifted state.
|
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.
|
SpacecraftState |
SpacecraftState.updateOrbit(Orbit newOrbit)
Update the orbit.
|
| Constructor and Description |
|---|
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.
|
SpacecraftState(double[] y,
OrbitType orbitType,
PositionAngle angleType,
AbsoluteDate date,
double mu,
Frame frame,
Map<String,AdditionalStateInfo> addStatesInfo,
Attitude attForces,
Attitude attEvents)
Build a spacecraft from an array (a state vector) and an additional states informations map.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
SpacecraftState |
J2DifferentialEffect.apply(SpacecraftState state1)
Apply the effect to a
spacecraft state. |
SpacecraftState |
AdapterPropagator.DifferentialEffect.apply(SpacecraftState original)
Apply the effect to a
spacecraft state. |
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.
|
| Constructor and Description |
|---|
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 applyBeforeIn,
PotentialCoefficientsProvider gravityField)
Simple constructor.
|
J2DifferentialEffect(SpacecraftState original,
AdapterPropagator.DifferentialEffect directEffect,
boolean applyBeforeIn,
double referenceRadius,
double mu,
double j2)
Simple constructor.
|
J2DifferentialEffect(SpacecraftState original,
AdapterPropagator.DifferentialEffect directEffect,
boolean applyBeforeIn,
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.
|
| Modifier and Type | Method and Description |
|---|---|
SymmetricMatrix |
OrbitCovariance.getCovarianceMatrix(Orbit refOrbit,
OrbitType covTypeOut,
Frame frameOut)
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.
|
| Constructor and Description |
|---|
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
|
| Modifier and Type | Method and Description |
|---|---|
protected abstract double[] |
AbstractTLEFitter.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 double[] |
LevenbergMarquardtOrbitConverter.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 |
TLESeries.getPVCoordinates(AbsoluteDate date)
Get the extrapolated position and velocity from an initial date.
|
PVCoordinates |
TLEPropagator.getPVCoordinates(AbsoluteDate date)
Get the extrapolated position and velocity from an initial TLE.
|
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. |
| Constructor and Description |
|---|
TLE(String line1In,
String line2In)
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.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
boolean |
EventState.evaluateStep(PatriusStepInterpolator interpolator)
Evaluate the impact of the proposed step on the event detector.
|
boolean |
EventState.evaluateStep(SpacecraftState state)
Evaluate the impact of the proposed step on the event handler.
|
EventDetector.Action |
ExtremaThreeBodiesAngleDetector.eventOccurred(Map<String,SpacecraftState> s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
ThreeBodiesAngleDetector.eventOccurred(Map<String,SpacecraftState> s,
boolean increasing,
boolean forward)
Handle an 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 |
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 |
ExtremaLatitudeDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an extrema latitude 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.
|
EventDetector.Action |
AlignmentDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an alignment 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 |
GroundMaskElevationDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an azimuth-elevation 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 |
AltitudeDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an altitude 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 |
NodeDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle a node crossing 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 |
ApparentElevationDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an apparent elevation event and choose what to do next.
|
EventDetector.Action |
ThreeBodiesAngleDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an angle 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.
|
EventDetector.Action |
ExtremaLongitudeDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an extrema distance 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 |
IntervalOccurrenceDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an 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 |
LongitudeDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle a longitude reaching 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 |
DihedralFieldOfViewDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an fov 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 |
NadirSolarIncidenceDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle a solar incidence event and choose 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 |
BetaAngleDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handles a beta angle event and chooses 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 |
ExtremaElevationDetector.eventOccurred(SpacecraftState s,
boolean increasing,
boolean forward)
Handle an extrema distance 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.
|
double |
ExtremaThreeBodiesAngleDetector.g(Map<String,SpacecraftState> s)
Compute the value of the switching function.
|
double |
ThreeBodiesAngleDetector.g(Map<String,SpacecraftState> s)
Compute the value of the switching function.
|
double |
DateDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
SolarTimeAngleDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ExtremaThreeBodiesAngleDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ExtremaLatitudeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ExtremaDistanceDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AlignmentDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LocalTimeAngleDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
GroundMaskElevationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
EventShifter.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AltitudeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
abstract double |
AbstractDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
NullMassPartDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
NodeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
CircularFieldOfViewDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ApparentElevationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ThreeBodiesAngleDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AnomalyDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ExtremaLongitudeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ApsideDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
IntervalOccurrenceDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ElevationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LongitudeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
LatitudeDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
DihedralFieldOfViewDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
NullMassDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
DistanceDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
NadirSolarIncidenceDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
EclipseDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
BetaAngleDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
EventDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
ExtremaElevationDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
double |
AOLDetector.g(SpacecraftState s)
Compute the value of the switching function.
|
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 |
NullMassPartDetector.resetState(SpacecraftState oldState)
Reset the state (including additional states) prior to continue propagation.
|
SpacecraftState |
IntervalOccurrenceDetector.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.
|
Map<String,SpacecraftState> |
ExtremaThreeBodiesAngleDetector.resetStates(Map<String,SpacecraftState> oldStates)
Reset the states (including additional states) prior to continue propagation.
|
Map<String,SpacecraftState> |
ThreeBodiesAngleDetector.resetStates(Map<String,SpacecraftState> oldStates)
Reset the states (including additional states) prior to continue propagation.
|
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.
|
| Constructor and Description |
|---|
DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget,
Vector3D centerIn,
Vector3D axis1,
double halfAperture1In,
Vector3D axis2,
double halfAperture2In,
double maxCheck)
Build a new instance.
|
DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget,
Vector3D centerIn,
Vector3D axis1,
double halfAperture1In,
Vector3D axis2,
double halfAperture2In,
double maxCheck,
double epsilon)
Build a new instance.
|
DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget,
Vector3D centerIn,
Vector3D axis1,
double halfAperture1In,
Vector3D axis2,
double halfAperture2In,
double maxCheck,
EventDetector.Action entry,
EventDetector.Action exit)
Build a new instance.
|
DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget,
Vector3D centerIn,
Vector3D axis1,
double halfAperture1In,
Vector3D axis2,
double halfAperture2In,
double maxCheck,
EventDetector.Action entry,
EventDetector.Action exit,
boolean removeEntry,
boolean removeExit)
Build a new instance.
|
DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget,
Vector3D centerIn,
Vector3D axis1,
double halfAperture1In,
Vector3D axis2,
double halfAperture2In,
double maxCheck,
EventDetector.Action entry,
EventDetector.Action exit,
boolean removeEntry,
boolean removeExit,
double epsilon)
Build a new instance.
|
DihedralFieldOfViewDetector(PVCoordinatesProvider pvTarget,
Vector3D centerIn,
Vector3D axis1,
double halfAperture1In,
Vector3D axis2,
double halfAperture2In,
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
The default implementation behavior is to
stop propagation when the local time
is reached. |
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,
CelestialBody sunModel,
double maxCheck,
double threshold,
EventDetector.Action action,
boolean remove)
Constructor for a SolarTimeDetector instance with complimentary parameters.
|
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.
|
| Modifier and Type | Method and Description |
|---|---|
abstract EventDetector.Action |
MultiAbstractDetector.eventOccurred(Map<String,SpacecraftState> s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
OneSatEventDetectorWrapper.eventOccurred(Map<String,SpacecraftState> s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
EventDetector.Action |
MultiEventDetector.eventOccurred(Map<String,SpacecraftState> s,
boolean increasing,
boolean forward)
Handle an event and choose what to do next.
|
abstract double |
MultiAbstractDetector.g(Map<String,SpacecraftState> s)
Compute the value of the switching function.
|
double |
OneSatEventDetectorWrapper.g(Map<String,SpacecraftState> s)
Compute the value of the switching function.
|
double |
MultiEventDetector.g(Map<String,SpacecraftState> s)
Compute the value of the switching function.
|
double |
OneSatEventDetectorWrapper.g(SpacecraftState s)
Compute the value of the switching function.
|
Map<String,SpacecraftState> |
MultiAbstractDetector.resetStates(Map<String,SpacecraftState> oldStates)
Reset the states (including additional states) prior to continue propagation.
|
Map<String,SpacecraftState> |
OneSatEventDetectorWrapper.resetStates(Map<String,SpacecraftState> oldStates)
Reset the states (including additional states) prior to continue propagation.
|
Map<String,SpacecraftState> |
MultiEventDetector.resetStates(Map<String,SpacecraftState> oldStates)
Reset the states (including additional states) prior to continue propagation.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Constructor and Description |
|---|
PartialDerivativesEquations(String nameIn,
NumericalPropagator propagatorIn)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
void |
MultiNumericalPropagator.addInitialState(SpacecraftState initialState,
String satId)
Add a new spacecraft state to be propagated.
|
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,
MultiAttitudeProvider attProviderForces,
MultiAttitudeProvider attProviderEvents,
String id)
Extract a given SpacecraftState from the state vector.
|
Map<String,SpacecraftState> |
MultiStateVectorInfo.mapArrayToStates(double[] y,
AbsoluteDate currentDate,
OrbitType orbitType,
PositionAngle angleType,
Map<String,MultiAttitudeProvider> attProvidersForces,
Map<String,MultiAttitudeProvider> 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.
|
| Modifier and Type | Method and Description |
|---|---|
SpacecraftState |
IntegratedEphemeris.getInitialState()
Get the propagator initial 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. |
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.
|
| Constructor and Description |
|---|
IntegratedEphemeris(List<AbsoluteDate> startDatesIn,
List<AbsoluteDate> minDatesIn,
List<AbsoluteDate> maxDatesIn,
OrbitType orbitTypeIn,
PositionAngle angleTypeIn,
AttitudeProvider attitudeForcesProvider,
AttitudeProvider attitudeEventsProvider,
Map<String,AdditionalStateInfo> additionalStateInfos,
List<ContinuousOutputModel> modelsIn,
Frame referenceFrameIn,
double muIn)
Creates a new instance of IntegratedEphemeris.
|
| Modifier and Type | Method and Description |
|---|---|
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. |
| Constructor and Description |
|---|
MultiIntegratedEphemeris(List<AbsoluteDate> startDatesIn,
List<AbsoluteDate> minDatesIn,
List<AbsoluteDate> maxDatesIn,
OrbitType orbitTypeIn,
PositionAngle angleTypeIn,
MultiAttitudeProvider multiAttitudeProviderForcesIn,
MultiAttitudeProvider multiAttitudeProviderEventsIn,
MultiStateVectorInfo stateInfos,
List<ContinuousOutputModel> modelsIn,
Frame referenceFrameIn,
String satIdIn)
Creates a new instance of IntegratedEphemeris.
|
| Modifier and Type | Method and Description |
|---|---|
SpacecraftState |
BasicStepInterpolator.getInterpolatedState()
Get the interpolated state.
|
SpacecraftState |
PatriusStepInterpolator.getInterpolatedState()
Get the interpolated state.
|
SpacecraftState |
AdaptedStepHandler.getInterpolatedState()
Get the interpolated state.
|
void |
PatriusStepInterpolator.setInterpolatedDate(AbsoluteDate date)
Set the interpolated date.
|
| Modifier and Type | Method and Description |
|---|---|
Map<String,SpacecraftState> |
MultiPatriusStepInterpolator.getInterpolatedStates()
Get all the interpolated states.
|
Map<String,SpacecraftState> |
MultiAdaptedStepHandler.getInterpolatedStates()
Get all the interpolated states.
|
| Modifier and Type | Method and Description |
|---|---|
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) |
| Modifier and Type | Method and Description |
|---|---|
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)
Calculates the ionospheric signal delay for the signal path from the position
of the transmitter and the receiver and the current date.
|
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)
Provides the R12 value for the Bent model.
|
void |
R12Loader.loadData(InputStream input,
String name)
Load data from a stream.
|
void |
USKLoader.loadData(InputStream input,
String name)
Load data from a stream.
|
| Constructor and Description |
|---|
R12Loader(String supportedFileName)
Constructor.
|
USKLoader(String fileName)
Creates a USK data file reader and load the file.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
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 |
MeeusMoonStela.getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
static void |
MeeusMoonStela.updateTransform(AbsoluteDate date,
Frame frame)
Update cached transform from
FramesFactory.getMOD(boolean) to provided frame. |
| Constructor and Description |
|---|
MeeusMoonStela(double inEarthRadius)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
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.getPressure(AbsoluteDate date,
Vector3D position,
Frame frame)
Returns pressure.
|
double |
MSIS00Adapter.getSpeedOfSound(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the local speed of sound.
|
Vector3D |
MSIS00Adapter.getVelocity(AbsoluteDate date,
Vector3D position,
Frame frame)
Get the spacecraft velocity relative to the atmosphere.
|
| Modifier and Type | Method and Description |
|---|---|
void |
StelaAeroModel.addDDragAccDParam(SpacecraftState s,
Parameter param,
double density,
Vector3D relativeVelocity,
double[] dAccdParam)
Compute acceleration derivatives with respect to additional parameters (the ballistic coefficient).
|
void |
StelaAeroModel.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).
|
double[][] |
StelaAtmosphericDrag.computePartialDerivatives(StelaEquinoctialOrbit orbit)
Compute the partial derivatives for a given spacecraft state.
|
double[] |
StelaAtmosphericDrag.computePerturbation(StelaEquinoctialOrbit orbit,
OrbitNatureConverter converter)
Compute the dE/dt force derivatives for a given spacecraft state.
|
double[] |
StelaAtmosphericDrag.computeShortPeriods(StelaEquinoctialOrbit orbit)
Compute the short periodic variations for a given spacecraft state.
|
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.
|
| Constructor and Description |
|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
double[][] |
StelaTesseralAttraction.computePartialDerivatives(StelaEquinoctialOrbit orbit)
Compute the partial derivatives for a given spacecraft state.
|
double[][] |
SolidTidesAcc.computePartialDerivatives(StelaEquinoctialOrbit orbit)
Compute the partial derivatives for a given spacecraft state.
|
double[][] |
StelaThirdBodyAttraction.computePartialDerivatives(StelaEquinoctialOrbit orbit) |
double[] |
StelaTesseralAttraction.computePerturbation(StelaEquinoctialOrbit orbit)
Compute the dE/dt force derivatives for a given spacecraft state.
|
double[] |
SolidTidesAcc.computePerturbation(StelaEquinoctialOrbit orbit)
Compute the dE/dt force derivatives for a given spacecraft state.
|
double[] |
StelaThirdBodyAttraction.computePerturbation(StelaEquinoctialOrbit orbit) |
double[] |
StelaTesseralAttraction.computeShortPeriods(StelaEquinoctialOrbit orbit)
Compute the short periodic variations for a given spacecraft state.
|
double[] |
SolidTidesAcc.computeShortPeriods(StelaEquinoctialOrbit orbit)
Compute the short periodic variations for a given spacecraft state.
|
double[] |
StelaThirdBodyAttraction.computeShortPeriods(StelaEquinoctialOrbit orbit) |
| Constructor and Description |
|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
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)
Compute the partial derivatives for a given spacecraft state.
|
double[] |
NonInertialContribution.computePerturbation(StelaEquinoctialOrbit orbit,
OrbitNatureConverter converter)
Compute the dE/dt force derivatives for a given spacecraft state.
|
double[] |
NonInertialContribution.computeShortPeriods(StelaEquinoctialOrbit orbit)
Compute the short periodic variations for a given spacecraft state.
|
| Modifier and Type | Method and Description |
|---|---|
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[][] |
SRPSquaring.computePartialDerivatives(StelaEquinoctialOrbit orbit)
Compute the partial derivatives for a given spacecraft state.
|
double[][] |
SRPPotential.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.
|
| Constructor and Description |
|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
PVCoordinates |
StelaEquinoctialOrbit.getPVCoordinates(AbsoluteDate otherDate,
Frame otherFrame)
Get the
PVCoordinates of the body in the selected frame. |
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. |
| Modifier and Type | Method and Description |
|---|---|
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 |
StelaAdditionalEquations.addInitialAdditionalState(SpacecraftState state) |
SpacecraftState |
StelaPartialDerivativesEquations.addInitialAdditionalState(SpacecraftState state) |
void |
StelaAdditionalEquations.computeDerivatives(StelaEquinoctialOrbit o,
double[] p,
double[] pDot)
Compute the derivatives related to the additional state parameters.
|
void |
StelaPartialDerivativesEquations.computeDerivatives(StelaEquinoctialOrbit orbit,
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)
Extrapolate a spacecraftState up to a specific target date.
|
protected 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.
|
protected SpacecraftState |
StelaGTOPropagator.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.
|
void |
StelaGTOPropagator.setInitialState(SpacecraftState initialState,
double massIn,
boolean isOsculatingIn)
Set the initial state.
|
void |
StelaBasicInterpolator.setInterpolatedDate(AbsoluteDate interpolatedDate)
Set the interpolated date.
|
void |
StelaAbstractPropagator.setOrbitFrame(Frame frame)
Set propagation frame.
|
| Constructor and Description |
|---|
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.
|
| Modifier and Type | Method and Description |
|---|---|
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).
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
void |
UTCTAIHistoryFilesLoader.loadData(InputStream input,
String name)
Load UTC-TAI offsets entries read from some file.
|
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).
|
T |
TimeShiftable.shiftedBy(double dt)
Get a time-shifted instance.
|
| Modifier and Type | Method and Description |
|---|---|
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.
|
| Constructor and Description |
|---|
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 dateIn,
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 dateIn,
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.
|
| Modifier and Type | Class and Description |
|---|---|
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. |
| Modifier and Type | Method and Description |
|---|---|
PatriusException |
PatriusExceptionWrapper.getException()
Get the wrapped exception.
|
| Modifier and Type | Method and Description |
|---|---|
static TimeStampedCacheException |
TimeStampedCacheException.unwrap(PatriusException oe)
Recover a PropagationException, possibly embedded in a
PatriusException. |
static PropagationException |
PropagationException.unwrap(PatriusException oe)
Recover a PropagationException, possibly embedded in a
PatriusException. |
| Constructor and Description |
|---|
PatriusException(PatriusException exception)
Copy constructor.
|
PatriusExceptionWrapper(PatriusException wrappedExceptionIn)
Simple constructor.
|
PropagationException(PatriusException exception)
Simple constructor.
|
TimeStampedCacheException(PatriusException exception)
Simple constructor.
|
| Modifier and Type | Method and Description |
|---|---|
L |
LegsSequence.getLeg(AbsoluteDate date)
Return the leg whose time interval contains the input date.
|
L |
AbstractLegsSequence.getLeg(AbsoluteDate date)
{inheritDoc}
|
AbsoluteDateInterval |
Leg.getTimeInterval()
Return the time interval of validity of the leg
|
AbsoluteDateInterval |
AbstractLegsSequence.getTimeInterval()
Return the time interval of validity of the leg
|
boolean |
AbstractLegsSequence.hasOverlaping()
Returns a boolean indicating if the sequence contains overlapping legs.
|
boolean |
AbstractLegsSequence.isCompact()
Returns a boolean indicating if the sequence is compact or not.
|
| Modifier and Type | Method and Description |
|---|---|
Vector3D |
SolarRadiationWrench.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 |
GravitationalAttractionWrench.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 |
GenericWrenchModel.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)
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 |
DragWrench.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 |
MagneticWrench.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 |
WrenchModel.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 |
DragWrench.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 |
MagneticWrench.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 |
WrenchModel.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 |
DragWrench.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 |
MagneticWrench.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 |
WrenchModel.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.
|
| Constructor and Description |
|---|
Wrench(double[] data)
Constructor from an array.
|
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