Path: ACPro/ACPointMass
% Compute the state derivatives of a point mass aircraft model.
State: x = [V;gamma;psi;x;y;h;Tbar]
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V true airspeed
gamma air relative flight path angle
psi air relative flight heading angle
x East position
y North position
h altitude
Tbar normalized excess thrust
Control: u = [Lbar;phi;Tcbar]
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Lbar normalized excess lift
phi bank angle
Tcbar normalized excess thrust command
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Form:
xDot = AircraftPointMassRHS( x, u )
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Inputs
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x (7,1) State vector
u (3,1) Control vector
data Data structure with fields:
a Body-frame disturbance acceleration
(forward,x-track,normal)
W Wind speeds (East,North,up)
g Gravitational acceleration
tau Engine thrust response time
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Outputs
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xDot (7,1) State time derivative
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