 ---------------------------------------------------------------
 Vortex Lattice Output -- Total Forces

 Configuration: Franck Aguerre Crobe XLM (http://lecrobe.free.fr)          
     # Surfaces =   4
     # Strips   =  52
     # Vortices = 392

  Sref =  46100.       Cref =  67.400       Bref =  684.00    
  Xref =  36.500       Yref =  0.0000       Zref = -14.000    

 Standard axis orientation,  X fwd, Z down         

 Run case:  -unnamed-                              

  Alpha =   0.00000     pb/2V =   0.00000     p'b/2V =   0.00000
  Beta  =   0.00000     qc/2V =   0.00000
  Mach  =     0.000     rb/2V =   0.00000     r'b/2V =   0.00000

  CXtot =  -0.00131     Cltot =   0.00000     Cl'tot =   0.00000
  CYtot =   0.00000     Cmtot =   0.01007
  CZtot =  -0.20500     Cntot =   0.00000     Cn'tot =   0.00000

  CLtot =   0.20500
  CDtot =   0.00131
  CDvis =   0.00000     CDind =   0.00131
  CLff  =   0.20499     CDff  =   0.00135    | Trefftz
  CYff  =   0.00000         e =    0.9786    | Plane  

   aileron         =   0.00000
   elevator        =   0.00000

 ---------------------------------------------------------------

 Stability-axis derivatives...

                             alpha                beta
                  ----------------    ----------------
 z' force CL |    CLa =   5.384886    CLb =   0.000000
 y  force CY |    CYa =   0.000000    CYb =  -0.103391
 x' mom.  Cl'|    Cla =   0.000000    Clb =  -0.024586
 y  mom.  Cm |    Cma =  -0.160171    Cmb =   0.000000
 z' mom.  Cn'|    Cna =   0.000000    Cnb =   0.039640

                     roll rate  p'      pitch rate  q'        yaw rate  r'
                  ----------------    ----------------    ----------------
 z' force CL |    CLp =   0.000000    CLq =   6.955002    CLr =   0.000000
 y  force CY |    CYp =   0.017341    CYq =   0.000001    CYr =   0.085258
 x' mom.  Cl'|    Clp =  -0.543702    Clq =   0.000000    Clr =   0.060817
 y  mom.  Cm |    Cmp =   0.000000    Cmq = -13.192294    Cmr =   0.000000
 z' mom.  Cn'|    Cnp =  -0.013391    Cnq =   0.000000    Cnr =  -0.033263

                  aileron      d1     elevator     d2 
                  ----------------    ----------------
 z' force CL |   CLd1 =   0.518822   CLd2 =  -0.778233
 y  force CY |   CYd1 =   0.000000   CYd2 =   0.000000
 x' mom.  Cl'|   Cld1 =   0.000000   Cld2 =   0.000000
 y  mom.  Cm |   Cmd1 =   0.146256   Cmd2 =  -0.219385
 z' mom.  Cn'|   Cnd1 =   0.000000   Cnd2 =   0.000000
 Trefftz drag| CDffd1 =   0.006805 CDffd2 =  -0.010208
 span eff.   |    ed1 =   0.008479    ed2 =  -0.012717



 Neutral point  Xnp =  38.504787

 Clb Cnr / Clr Cnb  =   0.339237    (  > 1 if spirally stable )
