Package org.orekit.propagation.numerical

Class NumericalPropagator

java.lang.Object

org.orekit.propagation.AbstractPropagator

org.orekit.propagation.integration.AbstractIntegratedPropagator

org.orekit.propagation.numerical.NumericalPropagator

All Implemented Interfaces:

Propagator, PVCoordinatesProvider

public class NumericalPropagator
extends AbstractIntegratedPropagator

This class propagates 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:

• the initial spacecraft state (setInitialState(SpacecraftState))
• the central attraction coefficient (setMu(double))
• the various force models (addForceModel(ForceModel), removeForceModels())
• the type of orbital parameters to be used for propagation (setOrbitType(OrbitType)),
• the type of position angle to be used in orbital parameters to be used for propagation where it is relevant (setPositionAngleType(PositionAngle)),
• whether additional equations (for example Jacobians) should be propagated along with orbital state (AbstractIntegratedPropagator.addAdditionalEquations(org.orekit.propagation.integration.AdditionalEquations)),
• the discrete events that should be triggered during propagation (AbstractIntegratedPropagator.addEventDetector(EventDetector), AbstractIntegratedPropagator.clearEventsDetectors())
• the binding logic with the rest of the application (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:

• the equinoctial orbit parameters (a, e_x, e_y, h_x, h_y, λ_M or λ_E or λ_v) in meters and radians,
• the Keplerian orbit parameters (a, e, i, ω, Ω, M or E or v) in meters and radians,
• the circular orbit parameters (a, e_x, e_y, i, Ω, α_M or α_E or α_v) in meters and radians,
• the Cartesian orbit parameters (x, y, z, v_x, v_y, v_z) 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.

Author:

Mathieu Roméro, Luc Maisonobe, Guylaine Prat, Fabien Maussion, Véronique Pommier-Maurussane

See Also:

SpacecraftState, ForceModel, OrekitStepHandler, OrekitFixedStepHandler, IntegratedEphemeris, TimeDerivativesEquations

Nested Class Summary

Nested classes/interfaces inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator

AbstractIntegratedPropagator.MainStateEquations

Field Summary

Fields inherited from interface org.orekit.propagation.Propagator

DEFAULT_LAW, DEFAULT_MASS, EPHEMERIS_GENERATION_MODE, MASTER_MODE, SLAVE_MODE

Constructor Summary

Constructors

Constructor	Description
NumericalPropagator(org.hipparchus.ode.ODEIntegrator integrator)	Create a new instance of NumericalPropagator, based on orbit definition mu.

Method Summary

Modifier and Type	Method	Description
void	addForceModel(ForceModel model)	Add a force model.
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(org.hipparchus.ode.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	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, Orbit orbit, OrbitType type)	Estimate tolerance vectors for integrators.

Methods inherited from class org.orekit.propagation.integration.AbstractIntegratedPropagator

addAdditionalEquations, addEventDetector, afterIntegration, beforeIntegration, clearEventsDetectors, getBasicDimension, getCalls, getEventsDetectors, getGeneratedEphemeris, getInitialIntegrationState, getIntegrator, getManagedAdditionalStates, getMu, initMapper, isAdditionalStateManaged, isMeanOrbit, propagate, propagate, propagate, setAttitudeProvider, setEphemerisMode, setEphemerisMode, setMasterMode, setResetAtEnd, setSlaveMode, setUpEventDetector, setUpUserEventDetectors

Methods inherited from class org.orekit.propagation.AbstractPropagator

addAdditionalStateProvider, getAdditionalStateProviders, getAttitudeProvider, getFixedStepSize, getFrame, getInitialState, getMode, getStartDate, getStepHandler, setMasterMode, setStartDate, updateAdditionalStates

Methods inherited from class java.lang.Object

clone, equals, finalize, getClass, hashCode, notify, notifyAll, toString, wait, wait, wait

Constructor Detail

NumericalPropagator

public NumericalPropagator(org.hipparchus.ode.ODEIntegrator integrator)

Create a new instance of NumericalPropagator, based on orbit definition mu. After creation, the instance is empty, i.e. the attitude provider is set to an unspecified default law and there are no perturbing forces at all. This means that if 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.

Parameters:

integrator - numerical integrator to use for propagation.

Method Detail

setMu

public void setMu(double mu)

Set the central attraction coefficient μ.

Setting the central attraction coefficient is equivalent to add a NewtonianAttraction force model.

Overrides:

setMu in class AbstractIntegratedPropagator

Parameters:

mu - central attraction coefficient (m³/s²)

See Also:

addForceModel(ForceModel), getAllForceModels()

addForceModel

public void addForceModel(ForceModel model)

Add a force model.

If this method is not called at all, the integrated orbit will follow a Keplerian evolution only.

Parameters:

model - ForceModel to add (it can be either a perturbing force model or an instance of NewtonianAttraction)

See Also:

removeForceModels(), setMu(double)

removeForceModels

public void removeForceModels()

Remove all force models (except central attraction).

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.

See Also:

addForceModel(ForceModel)

getAllForceModels

public List<ForceModel> getAllForceModels()

Get all the force models, perturbing forces and Newtonian attraction included.

Returns:

list of perturbing force models, with Newtonian attraction being the last one

See Also:

addForceModel(ForceModel), setMu(double)

setOrbitType

public void setOrbitType(OrbitType orbitType)

Set propagation orbit type.

Overrides:

setOrbitType in class AbstractIntegratedPropagator

Parameters:

orbitType - orbit type to use for propagation

getOrbitType

public OrbitType getOrbitType()

Get propagation parameter type.

Overrides:

getOrbitType in class AbstractIntegratedPropagator

Returns:

orbit type used for propagation

setPositionAngleType

public void setPositionAngleType(PositionAngle positionAngleType)

Set position angle type.

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.

Overrides:

setPositionAngleType in class AbstractIntegratedPropagator

Parameters:

positionAngleType - angle type to use for propagation

getPositionAngleType

public PositionAngle getPositionAngleType()

Get propagation parameter type.

Overrides:

getPositionAngleType in class AbstractIntegratedPropagator

Returns:

angle type to use for propagation

setInitialState

public void setInitialState(SpacecraftState initialState)

Set the initial state.

Parameters:

initialState - initial state

resetInitialState

public void resetInitialState(SpacecraftState state)

Reset the propagator initial state.

Specified by:

resetInitialState in interface Propagator

Overrides:

resetInitialState in class AbstractPropagator

Parameters:

state - new initial state to consider

getPVCoordinates

public TimeStampedPVCoordinates getPVCoordinates(AbsoluteDate date,
                                                 Frame frame)

Get the PVCoordinates of the body in the selected frame.

Specified by:

getPVCoordinates in interface PVCoordinatesProvider

Overrides:

getPVCoordinates in class AbstractPropagator

Parameters:

date - current date

frame - the frame where to define the position

Returns:

time-stamped position/velocity of the body (m and m/s)

createMapper

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. /** Simple constructor.

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.

Specified by:

createMapper in class AbstractIntegratedPropagator

Parameters:

referenceDate - reference date

mu - central attraction coefficient (m³/s²)

orbitType - orbit type to use for mapping

positionAngleType - angle type to use for propagation

attitudeProvider - attitude provider

frame - inertial frame

Returns:

new mapper

getMainStateEquations

protected AbstractIntegratedPropagator.MainStateEquations getMainStateEquations(org.hipparchus.ode.ODEIntegrator integrator)

Get the differential equations to integrate (for main state only).

Specified by:

getMainStateEquations in class AbstractIntegratedPropagator

Parameters:

integrator - numerical integrator to use for propagation.

Returns:

differential equations for main state

tolerances

public static double[][] tolerances(double dP,
                                    Orbit orbit,
                                    OrbitType type)

Estimate tolerance vectors for integrators.

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.

Parameters:

dP - user specified position error

orbit - reference orbit

type - propagation type for the meaning of the tolerance vectors elements (it may be different from orbit.getType())

Returns:

a two rows array, row 0 being the absolute tolerance error and row 1 being the relative tolerance error