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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -4.500  -0.0713   0.01181   0.00600  -0.1074   0.9837   0.0304
  -4.000  -0.0036   0.01079   0.00489  -0.1101   0.9775   0.0406
  -3.500   0.0698   0.00980   0.00388  -0.1140   0.9737   0.0605
  -3.000   0.1457   0.00914   0.00328  -0.1184   0.9710   0.0942
  -2.500   0.2110   0.00865   0.00289  -0.1204   0.9620   0.1333
  -2.000   0.2765   0.00819   0.00254  -0.1223   0.9529   0.1776
  -1.500   0.3371   0.00776   0.00228  -0.1231   0.9396   0.2338
  -1.000   0.3941   0.00738   0.00212  -0.1231   0.9217   0.3128
  -0.500   0.4504   0.00703   0.00201  -0.1229   0.9015   0.4155
   0.000   0.5049   0.00648   0.00202  -0.1224   0.8760   0.6209
   4.000   0.9268   0.00823   0.00308  -0.1148   0.5136   1.0000
   4.500   0.9760   0.00908   0.00357  -0.1136   0.4161   1.0000
   5.000   1.0173   0.01105   0.00454  -0.1115   0.2172   1.0000
   5.500   1.0626   0.01255   0.00550  -0.1100   0.1145   1.0000
   6.000   1.1066   0.01415   0.00671  -0.1081   0.0439   1.0000
   6.500   1.1504   0.01574   0.00829  -0.1059   0.0265   1.0000
   7.000   1.1956   0.01701   0.00970  -0.1040   0.0219   1.0000
   7.500   1.2290   0.01969   0.01256  -0.1000   0.0185   1.0000
   8.000   1.2722   0.02099   0.01401  -0.0979   0.0167   1.0000
   8.500   1.3100   0.02300   0.01622  -0.0948   0.0151   1.0000
   9.000   1.3447   0.02543   0.01880  -0.0915   0.0139   1.0000
  10.000   1.4040   0.03267   0.02677  -0.0837   0.0116   1.0000