XFOIL         Version 6.94
  
 Calculated polar for: FX 67-K-150/17 AIRFOIL                          
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.4618   0.01582   0.01058  -0.0835  0.6956  0.9274
   1.000   0.5177   0.01587   0.01051  -0.0837  0.6932  0.9404
   1.500   0.5504   0.01706   0.01180  -0.0817  0.6828  0.9579
   2.000   0.6155   0.01718   0.01187  -0.0845  0.6792  0.9687
   2.500   0.6965   0.01701   0.01165  -0.0902  0.6768  0.9742
   3.000   0.7757   0.01679   0.01139  -0.0955  0.6746  0.9799
   3.500   0.8451   0.01709   0.01171  -0.0995  0.6705  0.9880
   5.500   1.0810   0.01368   0.00847  -0.1033  0.6340  1.0000
   6.000   1.1300   0.01247   0.00722  -0.1018  0.6051  1.0000
   6.500   1.1738   0.01251   0.00737  -0.1002  0.5851  1.0000
   7.000   1.1865   0.01287   0.00760  -0.0927  0.5294  1.0000
   7.500   1.1433   0.01545   0.00962  -0.0777  0.4365  1.0000
   8.000   1.0938   0.02006   0.01366  -0.0651  0.3410  1.0000
   8.500   1.0487   0.02547   0.01843  -0.0549  0.2327  1.0000
   9.000   1.0055   0.03169   0.02374  -0.0465  0.0854  1.0000
   9.500   0.9925   0.03667   0.02827  -0.0416  0.0056  1.0000
  10.000   1.0118   0.03951   0.03131  -0.0394  0.0045  1.0000
  10.500   1.0259   0.04291   0.03494  -0.0371  0.0040  1.0000
  11.000   1.0361   0.04686   0.03908  -0.0349  0.0040  1.0000
  11.500   1.0441   0.05123   0.04362  -0.0329  0.0039  1.0000
  12.000   1.0498   0.05596   0.04848  -0.0309  0.0039  1.0000
  12.500   1.0635   0.05998   0.05261  -0.0292  0.0040  1.0000
  13.000   1.0902   0.06290   0.05575  -0.0277  0.0044  1.0000