XFOIL         Version 6.94
  
 Calculated polar for: ONERA OA206 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   0.0019   0.00718   0.00291   0.0247  0.8177  0.9736
   1.000   0.1046   0.00735   0.00289   0.0163  0.7620  0.9987
   1.500   0.1591   0.00747   0.00261   0.0168  0.6648  1.0000
   2.000   0.2118   0.00783   0.00248   0.0173  0.5464  1.0000
   2.500   0.2642   0.00844   0.00250   0.0175  0.4224  1.0000
   3.000   0.3155   0.00896   0.00268   0.0180  0.3521  1.0000
   3.500   0.3686   0.00941   0.00295   0.0182  0.3075  1.0000
   4.000   0.4243   0.00991   0.00323   0.0179  0.2519  1.0000
   4.500   0.4818   0.01053   0.00367   0.0174  0.2020  1.0000
   5.000   0.5391   0.01126   0.00419   0.0167  0.1617  1.0000
   5.500   0.5964   0.01207   0.00487   0.0160  0.1270  1.0000
   6.000   0.6535   0.01297   0.00566   0.0154  0.0994  1.0000
   6.500   0.7094   0.01412   0.00670   0.0148  0.0734  1.0000
   7.000   0.7648   0.01531   0.00790   0.0145  0.0585  1.0000
   7.500   0.8182   0.01695   0.00952   0.0143  0.0502  1.0000
   8.000   0.8713   0.01843   0.01111   0.0142  0.0433  1.0000
   8.500   0.9218   0.02045   0.01321   0.0143  0.0399  1.0000
   9.000   0.9715   0.02254   0.01536   0.0145  0.0366  1.0000
   9.500   1.0194   0.02515   0.01829   0.0150  0.0345  1.0000
  10.000   1.0637   0.02824   0.02163   0.0155  0.0323  1.0000
  10.500   1.1031   0.03180   0.02572   0.0162  0.0303  1.0000
  11.000   1.1380   0.03583   0.03005   0.0168  0.0294  1.0000
  11.500   1.1534   0.04191   0.03679   0.0177  0.0286  1.0000
  12.000   1.1357   0.04969   0.04535   0.0185  0.0283  1.0000
  12.500   1.0898   0.06108   0.05730   0.0124  0.0283  1.0000
  13.000   1.0446   0.07769   0.07427  -0.0015  0.0286  1.0000