org.orekit.forces.maneuvers

Class SmallManeuverAnalyticalModel

java.lang.Object

org.orekit.forces.maneuvers.SmallManeuverAnalyticalModel

All Implemented Interfaces:

AdapterPropagator.DifferentialEffect

public class SmallManeuverAnalyticalModel
extends Object
implements AdapterPropagator.DifferentialEffect

Analytical model for small maneuvers.

The aim of this model is to compute quickly the effect at date t₁ of a small maneuver performed at an earlier date t₀. Both the direct effect of the maneuver and the Jacobian of this effect with respect to maneuver parameters are available.

These effect are computed analytically using two Jacobian matrices:

1. J₀: Jacobian of Keplerian or equinoctial elements with respect to cartesian parameters at date t₀
allows to compute maneuver effect as a change in orbital elements at maneuver date t₀,
2. J_1/0: Jacobian of Keplerian or equinoctial elements at date t₁ with respect to Keplerian or equinoctial elements at date t₀
allows to propagate the change in orbital elements to final date t₁.

The second Jacobian, J_1/0, is computed using a simple Keplerian model, i.e. it is the identity except for the mean motion row which also includes an off-diagonal element due to semi-major axis change.

The orbital elements change at date t₁ can be added to orbital elements extracted from state, and the final elements taking account the changes are then converted back to appropriate type, which may be different from Keplerian or equinoctial elements.

Note that this model takes only Keplerian effects into account. This means that using only this class to compute an inclination maneuver in Low Earth Orbit will not change ascending node drift rate despite inclination has changed (the same would be true for a semi-major axis change of course). In order to take this drift into account, an instance of J2DifferentialEffect must be used together with an instance of this class.

Author:

Luc Maisonobe

Constructor Summary

Constructors

Constructor and Description
SmallManeuverAnalyticalModel(SpacecraftState state0, Frame frame, Vector3D dV, double isp) Build a maneuver defined in user-specified frame.
SmallManeuverAnalyticalModel(SpacecraftState state0, Vector3D dV, double isp) Build a maneuver defined in spacecraft frame.

Method Summary

Methods

Modifier and Type	Method and Description
Orbit	apply(Orbit orbit1) Compute the effect of the maneuver on an orbit.
SpacecraftState	apply(SpacecraftState state1) Compute the effect of the maneuver on a spacecraft state.
AbsoluteDate	getDate() Get the date of the maneuver.
Vector3D	getInertialDV() Get the inertial velocity increment of the maneuver.
Frame	getInertialFrame() Get the inertial frame in which the velocity increment is defined.
void	getJacobian(Orbit orbit1, PositionAngle positionAngle, double[][] jacobian) Compute the Jacobian of the orbit with respect to maneuver parameters.
double	updateMass(double mass) Update a spacecraft mass due to maneuver.

Methods inherited from class java.lang.Object

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

Constructor Detail

SmallManeuverAnalyticalModel

public SmallManeuverAnalyticalModel(SpacecraftState state0,
                            Vector3D dV,
                            double isp)
                             throws OrekitException

Build a maneuver defined in spacecraft frame.

Parameters:

state0 - state at maneuver date, before the maneuver is performed

dV - velocity increment in spacecraft frame

isp - engine specific impulse (s)

Throws:

OrekitException - if spacecraft frame cannot be transformed

SmallManeuverAnalyticalModel

public SmallManeuverAnalyticalModel(SpacecraftState state0,
                            Frame frame,
                            Vector3D dV,
                            double isp)
                             throws OrekitException

Build a maneuver defined in user-specified frame.

Parameters:

state0 - state at maneuver date, before the maneuver is performed

frame - frame in which velocity increment is defined

dV - velocity increment in specified frame

isp - engine specific impulse (s)

Throws:

OrekitException - if velocity increment frame cannot be transformed

Method Detail

getDate

public AbsoluteDate getDate()

Get the date of the maneuver.

Returns:

date of the maneuver

getInertialDV

public Vector3D getInertialDV()

Get the inertial velocity increment of the maneuver.

Returns:

velocity increment in a state-dependent inertial frame

See Also:

getInertialFrame()

getInertialFrame

public Frame getInertialFrame()

Get the inertial frame in which the velocity increment is defined.

Returns:

inertial frame in which the velocity increment is defined

See Also:

getInertialDV()

apply

public Orbit apply(Orbit orbit1)

Compute the effect of the maneuver on an orbit.

Parameters:

orbit1 - original orbit at t₁, without maneuver

Returns:

orbit at t₁, taking the maneuver into account if t₁ > t₀

See Also:

apply(SpacecraftState), getJacobian(Orbit, PositionAngle, double[][])

apply

public SpacecraftState apply(SpacecraftState state1)

Compute the effect of the maneuver on a spacecraft state.

Specified by:

apply in interface AdapterPropagator.DifferentialEffect

Parameters:

state1 - original spacecraft state at t₁, without maneuver

Returns:

spacecraft state at t₁, taking the maneuver into account if t₁ > t₀

See Also:

apply(Orbit), getJacobian(Orbit, PositionAngle, double[][])

getJacobian

public void getJacobian(Orbit orbit1,
               PositionAngle positionAngle,
               double[][] jacobian)
                 throws OrekitException

Compute the Jacobian of the orbit with respect to maneuver parameters.

The Jacobian matrix is a 6x4 matrix. Element jacobian[i][j] corresponds to the partial derivative of orbital parameter i with respect to maneuver parameter j. The rows order is the same order as used in Orbit.getJacobianWrtCartesian method. Columns (0, 1, 2) correspond to the velocity increment coordinates (ΔV_x, ΔV_y, ΔV_z) in the inertial frame returned by getInertialFrame(), and column 3 corresponds to the maneuver date t₀.

Parameters:

orbit1 - original orbit at t₁, without maneuver

positionAngle - type of the position angle to use

jacobian - placeholder 6x4 (or larger) matrix to be filled with the Jacobian, if matrix is larger than 6x4, only the 6x4 upper left corner will be modified

Throws:

OrekitException - if time derivative of the initial Jacobian cannot be computed

See Also:

apply(Orbit)

updateMass

public double updateMass(double mass)

Update a spacecraft mass due to maneuver.

Parameters:

mass - masse before maneuver

Returns:

mass after maneuver