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   Compute the state derivatives of a point mass aircraft model. 

   State:     x = [V;gama;psi;x;y;h;Tbar]
   --------------------------------------
     V     true airspeed
     gama  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]
   -------------------------------
     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
   ------
   x        (7,1)   State vector
   u        (3,1)   Control vector
   data             Data structure with fields:
                       a     Body-frame disturbance accel. (forward,x-track,normal)
                       W     Wind speeds (East,North,up)
                       g     Gravitatioanl acceleration
                       tau   Engine thrust response time

   -------
   Outputs
   -------
   xDot     (7,1)   State time derivative

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