XFOIL         Version 6.94
  
 Calculated polar for: Davis 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5323   0.00697   0.00229  -0.1026  0.7049  0.9857
   1.000   0.6153   0.00711   0.00229  -0.1087  0.6807  1.0000
   1.500   0.6585   0.00724   0.00231  -0.1062  0.6553  1.0000
   2.000   0.7014   0.00740   0.00240  -0.1035  0.6253  1.0000
   2.500   0.7432   0.00765   0.00251  -0.1006  0.5908  1.0000
   3.000   0.7754   0.00829   0.00274  -0.0957  0.5216  1.0000
   4.000   0.7911   0.01271   0.00455  -0.0782  0.0576  1.0000
   4.500   0.8292   0.01355   0.00517  -0.0749  0.0043  1.0000
   5.000   0.8724   0.01407   0.00581  -0.0724  0.0039  1.0000
   5.500   0.9146   0.01464   0.00650  -0.0698  0.0040  1.0000
   6.000   0.9545   0.01533   0.00734  -0.0667  0.0046  1.0000
   6.500   0.9903   0.01612   0.00826  -0.0628  0.0053  1.0000
   7.000   1.0217   0.01719   0.00953  -0.0581  0.0062  1.0000
   7.500   1.0406   0.01900   0.01159  -0.0516  0.0070  1.0000
   8.000   1.0655   0.02048   0.01321  -0.0466  0.0082  1.0000
   8.500   1.0781   0.02282   0.01586  -0.0399  0.0106  1.0000
   9.000   1.0773   0.02725   0.02046  -0.0321  0.0133  1.0000