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

 Configuration: cruise                                                      
     # Surfaces =   7
     # Strips   = 170
     # Vortices =1700

  Sref =  124.86       Cref =  4.2350       Bref =  34.320    
  Xref =  15.310       Yref =  0.0000       Zref =-0.48024    

 Standard axis orientation,  X fwd, Z down         

 Run case: case_06_02                              

  Alpha =   0.00000     pb/2V =  -0.00000     p'b/2V =  -0.00000
  Beta  =   0.00000     qc/2V =   0.00000
  Mach  =     0.850     rb/2V =  -0.00000     r'b/2V =  -0.00000

  CXtot =  -0.00057     Cltot =  -0.00000     Cl'tot =  -0.00000
  CYtot =  -0.00000     Cmtot =  -0.07665
  CZtot =  -0.11228     Cntot =  -0.00000     Cn'tot =  -0.00000

  CLtot =   0.11228
  CDtot =   0.00057
  CDvis =   0.00000     CDind = 0.0005704
  CLff  =   0.11222     CDff  = 0.0004812    | Trefftz
  CYff  =   0.00000         e =    0.8831    | Plane  

   slat            =   0.00000
   flap            =   0.00000
   aileron         =   0.00000
   elevator        =   0.00000

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

 Geometry-axis derivatives...

                    axial   vel. u     sideslip vel. v      normal  vel. w
                  ----------------    ----------------    ----------------
 x force CX  |    CXu =  -0.001141    CXv =  -0.000000    CXw =   0.050210
 y force CY  |    CYu =  -0.000000    CYv =  -0.848325    CYw =  -0.000001
 z force CZ  |    CZu =  -0.224561    CZv =  -0.000001    CZw =  -7.801182
 x mom.  Cl  |    Clu =  -0.000000    Clv =  -0.213462    Clw =  -0.000007
 y mom.  Cm  |    Cmu =  -0.153290    Cmv =   0.000003    Cmw = -10.691133
 z mom.  Cn  |    Cnu =  -0.000000    Cnv =   0.299343    Cnw =  -0.000000

                      roll rate  p       pitch rate  q         yaw rate  r
                  ----------------    ----------------    ----------------
 x force CX  |    CXp =  -0.000000    CXq =  -0.047075    CXr =   0.000000
 y force CY  |    CYp =  -0.269713    CYq =  -0.000012    CYr =   1.038168
 z force CZ  |    CZp =  -0.000013    CZq = -35.386948    CZr =   0.000000
 x mom.  Cl  |    Clp =  -0.624460    Clq =  -0.000040    Clr =   0.207818
 y mom.  Cm  |    Cmp =  -0.000023    Cmq =-100.329689    Cmr =  -0.000002
 z mom.  Cn  |    Cnp =   0.094013    Cnq =   0.000007    Cnr =  -0.554393

                  slat         d1     flap         d2     aileron      d3     elevator     d4 
                  ----------------    ----------------    ----------------    ----------------
 x force CX  |   CXd1 =  -0.000129   CXd2 =  -0.000059   CXd3 =   0.000000   CXd4 =  -0.000028
 y force CY  |   CYd1 =   0.000000   CYd2 =   0.000000   CYd3 =   0.000459   CYd4 =  -0.000000
 z force CZ  |   CZd1 =   0.000780   CZd2 =  -0.038435   CZd3 =   0.000000   CZd4 =  -0.013454
 x mom.  Cl  |   Cld1 =   0.000000   Cld2 =  -0.000000   Cld3 =   0.002578   Cld4 =  -0.000000
 y mom.  Cm  |   Cmd1 =  -0.000047   Cmd2 =  -0.038311   Cmd3 =   0.000000   Cmd4 =  -0.069543
 z mom.  Cn  |   Cnd1 =   0.000000   Cnd2 =  -0.000000   Cnd3 =  -0.000070   Cnd4 =   0.000000


