public class NumericalPropagator extends AbstractIntegratedPropagator
orbits
using
numerical integration.
Numerical propagation is much more accurate than analytical propagation
like for example Keplerian
or Eckstein-Hechler
, but requires a few more steps to set up to be used properly.
Whereas analytical propagators are configured only thanks to their various
constructors and can be used immediately after construction, numerical propagators
configuration involve setting several parameters between construction time
and propagation time.
The configuration parameters that can be set are:
setInitialState(SpacecraftState)
)setMu(double)
)addForceModel(ForceModel)
,
removeForceModels()
)type
of orbital parameters to be used for propagation
(setOrbitType(OrbitType)
),
type
of position angle to be used in orbital parameters
to be used for propagation where it is relevant (setPositionAngleType(PositionAngle)
),
additional equations
(for example Jacobians
) should be propagated along with orbital state
(AbstractIntegratedPropagator.addAdditionalEquations(org.orekit.propagation.integration.AdditionalEquations)
),
AbstractIntegratedPropagator.addEventDetector(EventDetector)
,
AbstractIntegratedPropagator.clearEventsDetectors()
)AbstractIntegratedPropagator.setSlaveMode()
,
AbstractPropagator.setMasterMode(double, org.orekit.propagation.sampling.OrekitFixedStepHandler)
,
AbstractIntegratedPropagator.setMasterMode(org.orekit.propagation.sampling.OrekitStepHandler)
,
AbstractIntegratedPropagator.setEphemerisMode()
, AbstractIntegratedPropagator.getGeneratedEphemeris()
)From these configuration parameters, only the initial state is mandatory. The default
propagation settings are in equinoctial
parameters with
true
longitude argument. If the central attraction coefficient
is not explicitly specified, the one used to define the initial orbit will be used.
However, specifying only the initial state and perhaps the central attraction coefficient
would mean the propagator would use only Keplerian forces. In this case, the simpler KeplerianPropagator
class would
perhaps be more effective.
The underlying numerical integrator set up in the constructor may also have its own configuration parameters. Typical configuration parameters for adaptive stepsize integrators are the min, max and perhaps start step size as well as the absolute and/or relative errors thresholds.
The state that is seen by the integrator is a simple seven elements double array. The six first elements are either:
equinoctial orbit parameters
(a, ex,
ey, hx, hy, λM or λE
or λv) in meters and radians,Keplerian orbit parameters
(a, e, i, ω, Ω,
M or E or v) in meters and radians,circular orbit parameters
(a, ex, ey, i,
Ω, αM or αE or αv) in meters
and radians,Cartesian orbit parameters
(x, y, z, vx,
vy, vz) in meters and meters per seconds.
The last element is the mass in kilograms.
The following code snippet shows a typical setting for Low Earth Orbit propagation in equinoctial parameters and true longitude argument:
final double dP = 0.001; final double minStep = 0.001; final double maxStep = 500; final double initStep = 60; final double[][] tolerance = NumericalPropagator.tolerances(dP, orbit, OrbitType.EQUINOCTIAL); AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tolerance[0], tolerance[1]); integrator.setInitialStepSize(initStep); propagator = new NumericalPropagator(integrator);
By default, at the end of the propagation, the propagator resets the initial state to the final state,
thus allowing a new propagation to be started from there without recomputing the part already performed.
This behaviour can be chenged by calling AbstractIntegratedPropagator.setResetAtEnd(boolean)
.
Beware the same instance cannot be used simultaneously by different threads, the class is not thread-safe.
SpacecraftState
,
ForceModel
,
OrekitStepHandler
,
OrekitFixedStepHandler
,
IntegratedEphemeris
,
TimeDerivativesEquations
AbstractIntegratedPropagator.MainStateEquations
DEFAULT_LAW, DEFAULT_MASS, EPHEMERIS_GENERATION_MODE, MASTER_MODE, SLAVE_MODE
Constructor and Description |
---|
NumericalPropagator(ODEIntegrator integrator)
Create a new instance of NumericalPropagator, based on orbit definition mu.
|
Modifier and Type | Method and Description |
---|---|
void |
addForceModel(ForceModel model)
Add a force model.
|
protected void |
beforeIntegration(SpacecraftState initialState,
AbsoluteDate tEnd)
Method called just before integration.
|
protected StateMapper |
createMapper(AbsoluteDate referenceDate,
double mu,
OrbitType orbitType,
PositionAngle positionAngleType,
AttitudeProvider attitudeProvider,
Frame frame)
Create a mapper between raw double components and spacecraft state.
|
List<ForceModel> |
getAllForceModels()
Get all the force models, perturbing forces and Newtonian attraction included.
|
protected AbstractIntegratedPropagator.MainStateEquations |
getMainStateEquations(ODEIntegrator integrator)
Get the differential equations to integrate (for main state only).
|
OrbitType |
getOrbitType()
Get propagation parameter type.
|
PositionAngle |
getPositionAngleType()
Get propagation parameter type.
|
TimeStampedPVCoordinates |
getPVCoordinates(AbsoluteDate date,
Frame frame)
Get the
PVCoordinates of the body in the selected frame. |
void |
removeForceModels()
Remove all force models (except central attraction).
|
void |
resetInitialState(SpacecraftState state)
Reset the propagator initial state.
|
void |
setIgnoreCentralAttraction(boolean ignoreCentralAttraction) |
void |
setInitialState(SpacecraftState initialState)
Set the initial state.
|
void |
setMu(double mu)
Set the central attraction coefficient μ.
|
void |
setOrbitType(OrbitType orbitType)
Set propagation orbit type.
|
void |
setPositionAngleType(PositionAngle positionAngleType)
Set position angle type.
|
static double[][] |
tolerances(double dP,
AbsolutePVCoordinates absPva)
Estimate tolerance vectors for integrators when propagating in absolute position-velocity-acceleration.
|
static double[][] |
tolerances(double dP,
Orbit orbit,
OrbitType type)
Estimate tolerance vectors for integrators when propagating in orbits.
|
addAdditionalEquations, addEventDetector, afterIntegration, clearEventsDetectors, getBasicDimension, getCalls, getEventsDetectors, getGeneratedEphemeris, getInitialIntegrationState, getIntegrator, getManagedAdditionalStates, getMu, initMapper, isAdditionalStateManaged, isMeanOrbit, propagate, propagate, propagate, setAttitudeProvider, setEphemerisMode, setEphemerisMode, setMasterMode, setResetAtEnd, setSlaveMode, setUpEventDetector, setUpUserEventDetectors
addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFixedStepSize, getFrame, getInitialState, getMode, getStartDate, getStepHandler, setMasterMode, setStartDate, updateAdditionalStates
public NumericalPropagator(ODEIntegrator integrator)
addForceModel
is not
called after creation, the integrated orbit will follow a Keplerian
evolution only. The defaults are OrbitType.EQUINOCTIAL
for propagation
orbit type
and PositionAngle.TRUE
for position angle type
.integrator
- numerical integrator to use for propagation.public void setIgnoreCentralAttraction(boolean ignoreCentralAttraction)
public void setMu(double mu)
Setting the central attraction coefficient is
equivalent to add
a NewtonianAttraction
force model.
setMu
in class AbstractIntegratedPropagator
mu
- central attraction coefficient (m³/s²)addForceModel(ForceModel)
,
getAllForceModels()
public void addForceModel(ForceModel model)
If this method is not called at all, the integrated orbit will follow a Keplerian evolution only.
model
- ForceModel
to add (it can be either a perturbing force
model or an instance of NewtonianAttraction
)removeForceModels()
,
setMu(double)
public void removeForceModels()
Once all perturbing forces have been removed (and as long as no new force model is added), the integrated orbit will follow a Keplerian evolution only.
addForceModel(ForceModel)
public List<ForceModel> getAllForceModels()
addForceModel(ForceModel)
,
setMu(double)
public void setOrbitType(OrbitType orbitType)
setOrbitType
in class AbstractIntegratedPropagator
orbitType
- orbit type to use for propagation, null for
propagating using AbsolutePVCoordinates
rather than Orbit
public OrbitType getOrbitType()
getOrbitType
in class AbstractIntegratedPropagator
AbsolutePVCoordinates
rather than Orbit
public void setPositionAngleType(PositionAngle positionAngleType)
The position parameter type is meaningful only if propagation orbit type
support it. As an example, it is not meaningful for propagation
in Cartesian
parameters.
setPositionAngleType
in class AbstractIntegratedPropagator
positionAngleType
- angle type to use for propagationpublic PositionAngle getPositionAngleType()
getPositionAngleType
in class AbstractIntegratedPropagator
public void setInitialState(SpacecraftState initialState)
initialState
- initial statepublic void resetInitialState(SpacecraftState state)
resetInitialState
in interface Propagator
resetInitialState
in class AbstractPropagator
state
- new initial state to considerpublic TimeStampedPVCoordinates getPVCoordinates(AbsoluteDate date, Frame frame)
PVCoordinates
of the body in the selected frame.getPVCoordinates
in interface PVCoordinatesProvider
getPVCoordinates
in class AbstractPropagator
date
- current dateframe
- the frame where to define the positionprotected StateMapper createMapper(AbsoluteDate referenceDate, double mu, OrbitType orbitType, PositionAngle positionAngleType, AttitudeProvider attitudeProvider, Frame frame)
The position parameter type is meaningful only if propagation orbit type
support it. As an example, it is not meaningful for propagation
in Cartesian
parameters.
createMapper
in class AbstractIntegratedPropagator
referenceDate
- reference datemu
- central attraction coefficient (m³/s²)orbitType
- orbit type to use for mappingpositionAngleType
- angle type to use for propagationattitudeProvider
- attitude providerframe
- inertial frameprotected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(ODEIntegrator integrator)
getMainStateEquations
in class AbstractIntegratedPropagator
integrator
- numerical integrator to use for propagation.public static double[][] tolerances(double dP, AbsolutePVCoordinates absPva)
dP
- user specified position errorabsPva
- reference absolute position-velocity-accelerationtolerances(double, Orbit, OrbitType)
public static double[][] tolerances(double dP, Orbit orbit, OrbitType type)
The errors are estimated from partial derivatives properties of orbits, starting from a scalar position error specified by the user. Considering the energy conservation equation V = sqrt(mu (2/r - 1/a)), we get at constant energy (i.e. on a Keplerian trajectory):
V² r |dV| = mu |dr|
So we deduce a scalar velocity error consistent with the position error. From here, we apply orbits Jacobians matrices to get consistent errors on orbital parameters.
The tolerances are only orders of magnitude, and integrator tolerances are only local estimates, not global ones. So some care must be taken when using these tolerances. Setting 1mm as a position error does NOT mean the tolerances will guarantee a 1mm error position after several orbits integration.
dP
- user specified position errororbit
- reference orbittype
- propagation type for the meaning of the tolerance vectors elements
(it may be different from orbit.getType()
)protected void beforeIntegration(SpacecraftState initialState, AbsoluteDate tEnd)
The default implementation does nothing, it may be specialized in subclasses.
beforeIntegration
in class AbstractIntegratedPropagator
initialState
- initial statetEnd
- target date at which state should be propagatedCopyright © 2002-2019 CS Systèmes d'information. All rights reserved.