Path: FormationFlying/EccDynamics
% Compute the state-transition matrix, R, given the eccentricity and true anomaly.
xH = R*D
where D is the vector of integration constants found from initial conditions.
Since version 7.
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Form:
R = FFEccRMat( nu, e, dH );
[R,dH] = FFEccRMat( nu, e );
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Inputs
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nu (1,N) True anomaly [rad]
e (1) Eccentricity
dH (1) Integration constant found by setting H = 0 at nu0
(Optional. If not provided, it will be computed
using the first nu supplied)
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Outputs
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R (6*N,6) Transformation matrix for equation xH = R*D
where D is the integration constant vector.
dH (1) Integration constant found by setting H = 0 at nu0
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References: Inalhan, Tillerson, How, "Relative Dynamics and Control of
Spacecraft Formations in Eccentric Orbits", Journal of Guidance,
Control & Dynamics, Vol.25, No.1, Jan-Feb 2002.
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Copyright 2004 Princeton Satellite Systems, Inc.
All rights reserved.
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