baseline_out_2.txt

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

 Configuration: WING_2022                                                   
     # Surfaces = 10
     # Strips   = 147
     # Vortices = 2456

  Sref = 0.39034       Cref = 0.28100       Bref = 1.4730    
  Xref = 0.0000       Yref = 0.0000       Zref = -0.70000E-01

 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 = -0.01979     Cltot = -0.00000     Cl'tot = -0.00000
  CYtot = 0.00000     Cmtot = 0.19341
  CZtot = -0.47811     Cntot = -0.00000     Cn'tot = -0.00000

  CLtot = 0.47811
  CDtot = 0.01979
  CDvis = 0.00000     CDind = 0.0197947
  CLff  = 0.47797     CDff  = 0.0221800    | Trefftz
  CYff  = 0.00000         e = 0.5898    | Plane  

   flaperon        = 0.00000
   elevator        = 0.00000
   rudder          = 0.00000

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

 Stability-axis derivatives...

                             alpha                beta
                  ----------------    ----------------
 z' force CL |    CLa = 5.583371    CLb = -0.000000
 y  force CY |    CYa = 0.000000    CYb = -0.245984
 x' mom.  Cl'|    Cla = -0.000000    Clb = -0.064577
 y  mom.  Cm |    Cma = -0.902036    Cmb = 0.000000
 z' mom.  Cn'|    Cna = -0.000000    Cnb = 0.067156

                     roll rate  p'      pitch rate  q'        yaw rate  r'
                  ----------------    ----------------    ----------------
 z' force CL |    CLp = 0.000000    CLq = 10.239977    CLr = -0.000000
 y  force CY |    CYp = 0.028715    CYq = -0.000000    CYr = 0.184588
 x' mom.  Cl'|    Clp = -0.527452    Clq = -0.000000    Clr = 0.166628
 y  mom.  Cm |    Cmp = -0.000000    Cmq = -11.679779    Cmr = 0.000000
 z' mom.  Cn'|    Cnp = -0.002406    Cnq = 0.000000    Cnr = -0.089882

                  flaperon     d01     elevator     d02     rudder       d03 
                  ----------------    ----------------    ----------------
 z' force CL |   CLd01 = 0.000000   CLd02 = 0.010368   CLd03 = -0.000000
 y  force CY |   CYd01 = 0.000033   CYd02 = 0.000000   CYd03 = -0.002300
 x' mom.  Cl'|   Cld01 = -0.005966   Cld02 = 0.000000   Cld03 = -0.000435
 y  mom.  Cm |   Cmd01 = -0.000000   Cmd02 = -0.027749   Cmd03 = 0.000000
 z' mom.  Cn'|   Cnd01 = -0.000549   Cnd02 = -0.000000   Cnd03 = 0.001186
 Trefftz drag| CDffd01 = 0.000000 CDffd02 = -0.000627 CDffd03 = -0.000000
 span eff.   |    ed01 = -0.000000    ed02 = 0.042223    ed03 = 0.000000



 Neutral point  Xnp =   0.045398

 Clb Cnr / Clr Cnb  =   0.518703    (  > 1 if spirally stable )