XFOIL         Version 6.96
  
 Calculated polar for: A18 (smoothed)
                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -3.000   0.1669   0.00682   0.00181  -0.1212   0.9461   0.0786
  -2.500   0.2228   0.00654   0.00159  -0.1210   0.9307   0.1088
  -2.000   0.2785   0.00632   0.00142  -0.1208   0.9121   0.1428
  -1.500   0.3343   0.00615   0.00129  -0.1205   0.8901   0.1815
  -1.000   0.3903   0.00601   0.00121  -0.1204   0.8649   0.2324
  -0.500   0.4462   0.00589   0.00119  -0.1203   0.8365   0.3035
   0.000   0.5020   0.00582   0.00121  -0.1202   0.8054   0.3852
   0.500   0.5577   0.00570   0.00128  -0.1202   0.7722   0.5063
   3.000   0.8249   0.00621   0.00197  -0.1171   0.5370   1.0000
   3.500   0.8780   0.00676   0.00225  -0.1166   0.4684   1.0000
   4.000   0.9289   0.00761   0.00265  -0.1159   0.3631   1.0000
   4.500   0.9743   0.00927   0.00346  -0.1145   0.1875   1.0000
   5.000   1.0242   0.01022   0.00408  -0.1136   0.1160   1.0000
   5.500   1.0729   0.01128   0.00480  -0.1125   0.0544   1.0000
   6.000   1.1216   0.01229   0.00566  -0.1113   0.0259   1.0000
   6.500   1.1706   0.01322   0.00660  -0.1101   0.0179   1.0000
   7.000   1.2193   0.01408   0.00753  -0.1088   0.0148   1.0000
   7.500   1.2640   0.01540   0.00898  -0.1069   0.0125   1.0000
   8.000   1.3098   0.01646   0.01015  -0.1051   0.0112   1.0000
   8.500   1.3537   0.01761   0.01140  -0.1032   0.0101   1.0000
   9.000   1.3851   0.02017   0.01417  -0.0992   0.0088   1.0000
   9.500   1.4261   0.02139   0.01553  -0.0968   0.0083   1.0000
  10.000   1.4623   0.02301   0.01733  -0.0938   0.0076   1.0000