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Static Stability of Aircraft designed in X Plane

  1. Mar 19, 2013 #1
    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 seperate 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.

  2. jcsd
  3. Apr 4, 2013 #2


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    Gold Member

    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 [itex]C_{D_u}> 0[/itex], [itex]C_{y_\beta} < 0[/itex], [itex]C_{L_\alpha} > 0[/itex], [itex]C_{l_\beta} < 0[/itex], [itex]C_{l_p} < 0[/itex], [itex]C_{m_\alpha} < 0[/itex], [itex]C_{m_q} < 0[/itex], [itex]C_{n_\beta} > 0[/itex], and [itex]C_{n_r} < 0[/itex]. (These are at steady state, with body-axis velocities [itex]U_1,V_1,W_1[/itex] and angular rates [itex]P_1,Q_1,R_1[/itex].)
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