org.orekit.propagation.semianalytical.dsst.forces

Class AbstractGaussianContribution

java.lang.Object

org.orekit.propagation.semianalytical.dsst.forces.AbstractGaussianContribution

All Implemented Interfaces:

DSSTForceModel

Direct Known Subclasses:

DSSTAtmosphericDrag, DSSTSolarRadiationPressure

public abstract class AbstractGaussianContribution
extends Object
implements DSSTForceModel

Common handling of DSSTForceModel methods for Gaussian contributions to DSST propagation.

This abstract class allows to provide easily a subset of DSSTForceModel methods for specific Gaussian contributions.

This class implements the notion of numerical averaging of the DSST theory. Numerical averaging is mainly used for non-conservative disturbing forces such as atmospheric drag and solar radiation pressure.

Gaussian contributions can be expressed as: da_i/dt = δa_i/δv . q
where:

• a_i are the six equinoctial elements
• v is the velocity vector
• q is the perturbing acceleration due to the considered force

The averaging process and other considerations lead to integrate this contribution over the true longitude L possibly taking into account some limits.

To create a numerically averaged contribution, one needs only to provide a ForceModel and to implement in the derived class the method: getLLimits(SpacecraftState).

Author:

Pascal Parraud

Field Summary

Fields

Modifier and Type	Field and Description
protected double	a a.
protected double	A A = sqrt(μ * a).
protected double	B B = sqrt(1 - h² - k²).
protected double	C C = 1 + p² + q².
protected double	co2AB C / (2 * A * B) .
protected double	ecc Eccentricity.
protected Vector3D	f Equinoctial frame f vector.
protected Vector3D	g Equinoctial frame g vector.
protected double	h ey.
protected double	k ex.
protected double	lm Mean longitude.
protected double	mu μ .
protected double	n Kepler mean motion: n = sqrt(μ / a³).
protected double	ooA 1 / A .
protected double	ooAB 1 / (A * B) .
protected double	ooBpo 1 / (1 + B) .
protected double	ooMu 1 / μ .
protected double	p hy.
protected double	q hx.
protected double	ton2a 2 / (n² * a) .
protected Vector3D	w Equinoctial frame w vector.

Constructor Summary

Constructors

Modifier	Constructor and Description
protected	AbstractGaussianContribution(String coefficientsKeyPrefix, double threshold, ForceModel contribution) Build a new instance.

Method Summary

Methods

Modifier and Type	Method and Description
protected Vector3D	getAcceleration(SpacecraftState state) Compute the acceleration due to the non conservative perturbing force.
protected abstract double[]	getLLimits(SpacecraftState state) Compute the limits in L, the true longitude, for integration.
double[]	getMeanElementRate(SpacecraftState state) Computes the mean equinoctial elements rates dai / dt.
List<ShortPeriodTerms>	initialize(AuxiliaryElements aux, boolean meanOnly) Performs initialization prior to propagation for the current force model.
void	initializeStep(AuxiliaryElements aux) Performs initialization at each integration step for the current force model.
void	registerAttitudeProvider(AttitudeProvider provider) Register an attitude provider.
void	updateShortPeriodTerms(SpacecraftState... meanStates) Update the short period terms.

Methods inherited from class java.lang.Object

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

Methods inherited from interface org.orekit.propagation.semianalytical.dsst.forces.DSSTForceModel

getEventsDetectors

Field Detail

a

protected double a

a.

k

protected double k

e_x.

h

protected double h

e_y.

q

protected double q

h_x.

p

protected double p

h_y.

ecc

protected double ecc

Eccentricity.

n

protected double n

Kepler mean motion: n = sqrt(μ / a³).

lm

protected double lm

Mean longitude.

f

protected Vector3D f

Equinoctial frame f vector.

g

protected Vector3D g

Equinoctial frame g vector.

w

protected Vector3D w

Equinoctial frame w vector.

A

protected double A

A = sqrt(μ * a).

B

protected double B

B = sqrt(1 - h² - k²).

C

protected double C

C = 1 + p² + q².

ton2a

protected double ton2a

2 / (n² * a) .

ooA

protected double ooA

1 / A .

ooAB

protected double ooAB

1 / (A * B) .

co2AB

protected double co2AB

C / (2 * A * B) .

ooBpo

protected double ooBpo

1 / (1 + B) .

ooMu

protected double ooMu

1 / μ .

mu

protected double mu

μ .

Constructor Detail

AbstractGaussianContribution

protected AbstractGaussianContribution(String coefficientsKeyPrefix,
                            double threshold,
                            ForceModel contribution)

Build a new instance.

Parameters:

coefficientsKeyPrefix - prefix for coefficients keys

threshold - tolerance for the choice of the Gauss quadrature order

contribution - the ForceModel to be numerically averaged

Method Detail

initialize

public List<ShortPeriodTerms> initialize(AuxiliaryElements aux,
                                boolean meanOnly)

Performs initialization prior to propagation for the current force model.

This method aims at being called at the very beginning of a propagation.

Specified by:

initialize in interface DSSTForceModel

Parameters:

aux - auxiliary elements related to the current orbit

meanOnly - only mean elements are used during the propagation

Returns:

a list of objects that will hold short period terms (the objects are also retained by the force model, which will update them during propagation)

initializeStep

public void initializeStep(AuxiliaryElements aux)
                    throws OrekitException

Performs initialization at each integration step for the current force model.

This method aims at being called before mean elements rates computation.

Specified by:

initializeStep in interface DSSTForceModel

Parameters:

aux - auxiliary elements related to the current orbit

Throws:

OrekitException - if some specific error occurs

getMeanElementRate

public double[] getMeanElementRate(SpacecraftState state)
                            throws OrekitException

Computes the mean equinoctial elements rates da_i / dt.

Specified by:

getMeanElementRate in interface DSSTForceModel

Parameters:

state - current state information: date, kinematics, attitude

Returns:

the mean element rates dai/dt

Throws:

OrekitException - if some specific error occurs

getAcceleration

protected Vector3D getAcceleration(SpacecraftState state)
                            throws OrekitException

Compute the acceleration due to the non conservative perturbing force.

Parameters:

state - current state information: date, kinematics, attitude

Returns:

the perturbing acceleration

Throws:

OrekitException - if some specific error occurs

getLLimits

protected abstract double[] getLLimits(SpacecraftState state)
                                throws OrekitException

Compute the limits in L, the true longitude, for integration.

Parameters:

state - current state information: date, kinematics, attitude

Returns:

the integration limits in L

Throws:

OrekitException - if some specific error occurs

registerAttitudeProvider

public void registerAttitudeProvider(AttitudeProvider provider)

Register an attitude provider.

Register an attitude provider that can be used by the force model.

Specified by:

registerAttitudeProvider in interface DSSTForceModel

Parameters:

provider - the AttitudeProvider

updateShortPeriodTerms

public void updateShortPeriodTerms(SpacecraftState... meanStates)
                            throws OrekitException

Update the short period terms.

The short period terms that will be updated are the ones that were returned during the call to DSSTForceModel.initialize(AuxiliaryElements, boolean).

Specified by:

updateShortPeriodTerms in interface DSSTForceModel

Parameters:

meanStates - mean states information: date, kinematics, attitude

Throws:

OrekitException - if some specific error occurs