XFOIL         Version 6.94
  
 Calculated polar for: OAF117 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   0.2548   0.01042   0.00370  -0.0550  0.6160  0.5186
   0.500   0.3138   0.01044   0.00374  -0.0552  0.5877  0.5658
   1.000   0.3726   0.01037   0.00385  -0.0553  0.5601  0.6320
   1.500   0.4300   0.01022   0.00397  -0.0550  0.5314  0.7236
   2.000   0.4752   0.00955   0.00378  -0.0513  0.5057  1.0000
   2.500   0.5358   0.00987   0.00391  -0.0520  0.4794  1.0000
   3.000   0.5957   0.01024   0.00411  -0.0524  0.4536  1.0000
   3.500   0.6554   0.01058   0.00435  -0.0529  0.4277  1.0000
   4.000   0.7147   0.01094   0.00465  -0.0532  0.4010  1.0000
   4.500   0.7736   0.01135   0.00497  -0.0535  0.3732  1.0000
   5.000   0.8319   0.01183   0.00534  -0.0538  0.3427  1.0000
   5.500   0.8899   0.01231   0.00575  -0.0541  0.3079  1.0000
   6.000   0.9473   0.01292   0.00625  -0.0543  0.2719  1.0000
   6.500   1.0039   0.01365   0.00689  -0.0545  0.2343  1.0000
   7.000   1.0595   0.01454   0.00763  -0.0546  0.1946  1.0000
   7.500   1.1138   0.01564   0.00858  -0.0547  0.1581  1.0000
   8.000   1.1666   0.01691   0.00972  -0.0546  0.1212  1.0000
   8.500   1.2167   0.01857   0.01120  -0.0543  0.0885  1.0000
   9.000   1.2643   0.02040   0.01297  -0.0536  0.0704  1.0000
   9.500   1.3080   0.02249   0.01503  -0.0526  0.0603  1.0000
  10.500   1.3874   0.02673   0.01945  -0.0493  0.0490  1.0000
  11.000   1.4193   0.02924   0.02210  -0.0470  0.0453  1.0000
  11.500   1.4427   0.03219   0.02505  -0.0439  0.0425  1.0000
  12.000   1.4588   0.03505   0.02821  -0.0401  0.0405  1.0000
  12.500   1.4694   0.03867   0.03198  -0.0372  0.0388  1.0000
  13.000   1.4806   0.04300   0.03637  -0.0351  0.0373  1.0000
  13.500   1.4852   0.04817   0.04191  -0.0347  0.0359  1.0000
  14.000   1.4886   0.05382   0.04782  -0.0349  0.0346  1.0000
  14.500   1.4943   0.05930   0.05340  -0.0349  0.0335  1.0000
  15.000   1.4927   0.06594   0.06025  -0.0350  0.0327  1.0000
  15.500   1.4739   0.07529   0.07004  -0.0382  0.0322  1.0000
  16.000   1.4492   0.08591   0.08107  -0.0425  0.0317  1.0000
  16.500   1.4189   0.09828   0.09383  -0.0484  0.0313  1.0000
  17.000   1.3805   0.11330   0.10927  -0.0569  0.0311  1.0000
  17.500   1.3245   0.13371   0.13015  -0.0701  0.0312  1.0000