XFOIL         Version 6.94
  
 Calculated polar for: GIII BL167 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.0976   0.00608   0.00199  -0.0136  0.8230  0.9233
   0.500   0.1529   0.00622   0.00208  -0.0128  0.8039  0.9531
   1.000   0.2174   0.00629   0.00210  -0.0147  0.7865  0.9656
   1.500   0.2872   0.00636   0.00211  -0.0179  0.7686  0.9751
   2.000   0.3585   0.00642   0.00214  -0.0215  0.7489  0.9851
   2.500   0.4297   0.00645   0.00213  -0.0250  0.7109  0.9936
   3.000   0.4905   0.00683   0.00198  -0.0264  0.5668  1.0000
   3.500   0.5304   0.00813   0.00232  -0.0243  0.3432  1.0000
   4.000   0.5640   0.01024   0.00308  -0.0216  0.0700  1.0000
   4.500   0.6078   0.01100   0.00362  -0.0197  0.0279  1.0000
   5.000   0.6539   0.01158   0.00421  -0.0181  0.0176  1.0000
   5.500   0.7005   0.01222   0.00485  -0.0167  0.0099  1.0000
   6.000   0.7478   0.01296   0.00569  -0.0153  0.0043  1.0000
   6.500   0.7944   0.01396   0.00683  -0.0138  0.0034  1.0000
   7.000   0.8399   0.01528   0.00839  -0.0121  0.0034  1.0000
   7.500   0.8819   0.01728   0.01071  -0.0099  0.0035  1.0000
   8.000   0.9194   0.02028   0.01411  -0.0072  0.0036  1.0000
   8.500   0.9453   0.02765   0.02257  -0.0022  0.0044  1.0000
   9.000   0.9272   0.04370   0.04016   0.0058  0.0060  1.0000