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

 Configuration: Fennec                                                      
     # Surfaces =   4
     # Strips   =  72
     # Vortices = 768

  Sref = 0.17565E+06   Cref =  120.30       Bref =  1460.0    
  Xref =  67.000       Yref = -5.5000       Zref = -18.000    

 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 =  -0.00154     Cltot =  -0.00091     Cl'tot =  -0.00091
  CYtot =   0.00000     Cmtot =   0.00000
  CZtot =  -0.24187     Cntot =   0.00001     Cn'tot =   0.00001

  CLtot =   0.24187
  CDtot =   0.00154
  CDvis =   0.00000     CDind =   0.00154
  CLff  =   0.24177     CDff  =   0.00155    | Trefftz
  CYff  =   0.00000         e =    0.9895    | Plane  

   aileron         =   0.00000
   elevator        =  -0.00861
   rudder          =   0.00000

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

 Stability-axis derivatives...

                             alpha                beta
                  ----------------    ----------------
 z' force CL |    CLa =   5.490069    CLb =   0.000000
 y  force CY |    CYa =   0.000000    CYb =  -0.112870
 x' mom.  Cl'|    Cla =  -0.020682    Clb =  -0.066824
 y  mom.  Cm |    Cma =  -0.363511    Cmb =   0.000000
 z' mom.  Cn'|    Cna =   0.000261    Cnb =   0.033785

                     roll rate  p'      pitch rate  q'        yaw rate  r'
                  ----------------    ----------------    ----------------
 z' force CL |    CLp =   0.041375    CLq =   7.140500    CLr =  -0.003575
 y  force CY |    CYp =  -0.059150    CYq =   0.000000    CYr =   0.083033
 x' mom.  Cl'|    Clp =  -0.583685    Clq =  -0.026899    Clr =   0.070563
 y  mom.  Cm |    Cmp =  -0.002739    Cmq = -12.582802    Cmr =   0.000015
 z' mom.  Cn'|    Cnp =  -0.018473    Cnq =   0.000323    Cnr =  -0.025506

                  aileron      d1     elevator     d2     rudder       d3 
                  ----------------    ----------------    ----------------
 z' force CL |   CLd1 =   0.000000   CLd2 =  -1.069027   CLd3 =   0.000000
 y  force CY |   CYd1 =  -0.000831   CYd2 =   0.000000   CYd3 =   0.036726
 x' mom.  Cl'|   Cld1 =   0.134450   Cld2 =   0.004027   Cld3 =   0.004405
 y  mom.  Cm |   Cmd1 =   0.000000   Cmd2 =  -0.229952   Cmd3 =   0.000000
 z' mom.  Cn'|   Cnd1 =   0.008429   Cnd2 =   0.000109   Cnd3 =  -0.012265
 Trefftz drag| CDffd1 =   0.000000 CDffd2 =  -0.013714 CDffd3 =   0.000000
 span eff.   |    ed1 =   0.000000    ed2 =   0.012320    ed3 =   0.000000



 Neutral point  Xnp =  74.965370

 Clb Cnr / Clr Cnb  =   0.714970    (  > 1 if spirally stable )
