baseline_out_3.txt

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

 Configuration: WING_2022                                                   
     # Surfaces = 8
     # Strips   = 103
     # Vortices = 2060

  Sref = 0.39000       Cref = 0.23200       Bref = 1.6700    
  Xref = 0.0000       Yref = 0.0000       Zref = 0.0000

 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 = -0.00304     Cltot = 0.00000     Cl'tot = 0.00000
  CYtot = 0.00000     Cmtot = 0.26191
  CZtot = 0.08174     Cntot = -0.00000     Cn'tot = -0.00000

  CLtot = -0.08174
  CDtot = 0.00304
  CDvis = 0.00000     CDind = 0.0030421
  CLff  = -0.08174     CDff  = 0.0030431    | Trefftz
  CYff  = 0.00000         e = 0.0977    | Plane  

   aileron        = 0.00000
   elevator        = 0.00000
   rudder          = 0.00000

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

 Stability-axis derivatives...

                             alpha                beta
                  ----------------    ----------------
 z' force CL |    CLa = 5.743073    CLb = -0.000000
 y  force CY |    CYa = -0.000000    CYb = -0.312294
 x' mom.  Cl'|    Cla = -0.000000    Clb = -0.013705
 y  mom.  Cm |    Cma = -0.907315    Cmb = 0.000000
 z' mom.  Cn'|    Cna = -0.000000    Cnb = 0.057990

                     roll rate  p'      pitch rate  q'        yaw rate  r'
                  ----------------    ----------------    ----------------
 z' force CL |    CLp = -0.000000    CLq = 11.672971    CLr = -0.000000
 y  force CY |    CYp = -0.034163    CYq = -0.000000    CYr = 0.191161
 x' mom.  Cl'|    Clp = -0.541604    Clq = 0.000000    Clr = 0.010390
 y  mom.  Cm |    Cmp = 0.000000    Cmq = -16.872805    Cmr = -0.000000
 z' mom.  Cn'|    Cnp = 0.013280    Cnq = 0.000000    Cnr = -0.068533

                  aileron     d01     elevator     d02     rudder       d03 
                  ----------------    ----------------    ----------------
 z' force CL |   CLd01 = 0.000000   CLd02 = 0.010320   CLd03 = -0.000000
 y  force CY |   CYd01 = 0.000000   CYd02 = 0.000000   CYd03 = -0.002307
 x' mom.  Cl'|   Cld01 = -0.000000   Cld02 = -0.000000   Cld03 = -0.000291
 y  mom.  Cm |   Cmd01 = 0.000000   Cmd02 = -0.033505   Cmd03 = 0.000000
 z' mom.  Cn'|   Cnd01 = -0.000000   Cnd02 = -0.000000   Cnd03 = 0.001048
 Trefftz drag| CDffd01 = 0.000000 CDffd02 = -0.000770 CDffd03 = -0.000000
 span eff.   |    ed01 = 0.000000    ed02 = 0.000039    ed03 = 0.000000



 Neutral point  Xnp =   0.036652

 Clb Cnr / Clr Cnb  =   1.558854    (  > 1 if spirally stable )