Static Stability of Aircraft designed in X Plane

[Sicknote]

HI Everyone.

First of all, I'm not an aeronautical or mechanical engineer. I'm an electronics engineer but I'm currently doing a group project where I'm required to carry out a stability analysis on an aircraft that we've designed using the simulator software, X Plane.

I apologise in advance if any of these questions are stupid...

I know where the aircraft centre of gravity is but
How do I determine the Aerodynamic Centre of my aircraft?

I know that during steady flight, I must prove that all resultant moments acting around the centre of gravity should = 0, and
I've got Cm vs Alpha graphs for the tail airfoil (NACA 0009) and I'm in the process of trying to get the same graphs off X plane for the wing airfoils (Boeing Subsonic) but these will give me 2 completely separate moment coefficients for 2 different airfoils.

How do I get the Moment Coefficient about the Centre of Gravity from these?

Any help would be very gratefully received.

Thanks

 

[jhae2.718]

You might look at the USAF Stability and Control Data Compendium (DATACOM) which lists methods for determining S&C characteristics. http://www.pdas.com/datcomrefs.html has a PDF copy of the report, and also FORTRAN source code for Digital DATCOM.

Hans Multhopp created a method for determining the aerodynamic center. Roskam's Aircraft Design books discuss the method in Part 6, and there is a NACA technical report by Multhopp, but I don't remember the report number. Multhopp's method is accurate to about 10%. (It also is extremely tedious and a pain to do.)

For static stability you'll want to show that $C_{D_u}> 0$, $C_{y_\beta} < 0$, $C_{L_\alpha} > 0$, $C_{l_\beta} < 0$, $C_{l_p} < 0$, $C_{m_\alpha} < 0$, $C_{m_q} < 0$, $C_{n_\beta} > 0$, and $C_{n_r} < 0$. (These are at steady state, with body-axis velocities $U_1,V_1,W_1$ and angular rates $P_1,Q_1,R_1$.)