XFOIL         Version 6.94
  
 Calculated polar for: OAF128 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1380   0.01082   0.00394  -0.0288  0.5669  0.5320
   0.500   0.1982   0.01083   0.00397  -0.0291  0.5394  0.5681
   1.000   0.2582   0.01082   0.00405  -0.0294  0.5147  0.6110
   1.500   0.3176   0.01081   0.00418  -0.0296  0.4905  0.6683
   2.000   0.3754   0.01076   0.00437  -0.0293  0.4674  0.7439
   2.500   0.4285   0.01068   0.00454  -0.0277  0.4438  0.8394
   3.000   0.4753   0.01028   0.00426  -0.0245  0.4147  1.0000
   3.500   0.5377   0.01058   0.00444  -0.0256  0.3856  1.0000
   4.000   0.5991   0.01096   0.00466  -0.0264  0.3565  1.0000
   4.500   0.6598   0.01138   0.00496  -0.0272  0.3266  1.0000
   5.000   0.7201   0.01184   0.00532  -0.0278  0.2953  1.0000
   5.500   0.7798   0.01240   0.00575  -0.0284  0.2635  1.0000
   6.000   0.8386   0.01310   0.00628  -0.0290  0.2284  1.0000
   6.500   0.8968   0.01392   0.00698  -0.0296  0.1874  1.0000
   7.000   0.9535   0.01509   0.00788  -0.0302  0.1442  1.0000
   7.500   1.0090   0.01640   0.00904  -0.0305  0.1212  1.0000
   8.000   1.0628   0.01782   0.01039  -0.0306  0.1079  1.0000
   8.500   1.1153   0.01926   0.01180  -0.0306  0.0982  1.0000
   9.500   1.2169   0.02217   0.01485  -0.0298  0.0847  1.0000
  10.000   1.2631   0.02406   0.01671  -0.0290  0.0793  1.0000
  10.500   1.3100   0.02556   0.01839  -0.0282  0.0746  1.0000
  11.000   1.3517   0.02765   0.02054  -0.0270  0.0705  1.0000
  11.500   1.3920   0.02951   0.02256  -0.0257  0.0666  1.0000
  12.000   1.4264   0.03197   0.02511  -0.0239  0.0634  1.0000
  12.500   1.4549   0.03427   0.02764  -0.0217  0.0602  1.0000
  13.000   1.4774   0.03716   0.03052  -0.0190  0.0577  1.0000
  13.500   1.4803   0.04094   0.03470  -0.0166  0.0556  1.0000
  14.000   1.4893   0.04520   0.03915  -0.0162  0.0535  1.0000
  14.500   1.5016   0.04966   0.04364  -0.0156  0.0516  1.0000
  15.000   1.4912   0.05701   0.05145  -0.0180  0.0501  1.0000
  15.500   1.4820   0.06471   0.05944  -0.0210  0.0484  1.0000
  16.000   1.4816   0.07121   0.06604  -0.0231  0.0470  1.0000
  16.500   1.4678   0.07980   0.07484  -0.0261  0.0458  1.0000
  17.000   1.4275   0.09342   0.08894  -0.0336  0.0450  1.0000
  17.500   1.3814   0.10945   0.10540  -0.0431  0.0441  1.0000