XFOIL         Version 6.94
  
 Calculated polar for: A18 (smoothed)                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.5057   0.00653   0.00224  -0.1230  0.8840  0.7111
   0.500   0.5551   0.00596   0.00205  -0.1206  0.8555  1.0000
   1.000   0.6104   0.00611   0.00204  -0.1200  0.8241  1.0000
   1.500   0.6649   0.00634   0.00207  -0.1193  0.7850  1.0000
   2.000   0.7179   0.00666   0.00215  -0.1182  0.7308  1.0000
   2.500   0.7718   0.00702   0.00232  -0.1175  0.6873  1.0000
   3.000   0.8255   0.00741   0.00258  -0.1168  0.6446  1.0000
   3.500   0.8785   0.00786   0.00286  -0.1160  0.5954  1.0000
   4.500   0.9792   0.00924   0.00370  -0.1136  0.4377  1.0000
   5.000   1.0203   0.01118   0.00464  -0.1114  0.2380  1.0000
   5.500   1.0644   0.01284   0.00572  -0.1097  0.1233  1.0000
   6.000   1.1076   0.01456   0.00699  -0.1076  0.0449  1.0000
   6.500   1.1503   0.01632   0.00877  -0.1051  0.0276  1.0000
   7.000   1.1936   0.01785   0.01050  -0.1026  0.0231  1.0000
   7.500   1.2316   0.01991   0.01267  -0.0995  0.0191  1.0000
   8.000   1.2651   0.02284   0.01586  -0.0955  0.0175  1.0000
   8.500   1.3035   0.02528   0.01856  -0.0924  0.0162  1.0000
   9.000   1.3406   0.02775   0.02128  -0.0894  0.0145  1.0000
   9.500   1.3732   0.03022   0.02393  -0.0861  0.0130  1.0000
  10.000   1.3897   0.03769   0.03201  -0.0808  0.0121  1.0000
  11.000   1.3822   0.04810   0.04356  -0.0654  0.0114  1.0000
  11.500   1.3547   0.05476   0.05074  -0.0578  0.0113  1.0000
  12.000   1.3171   0.06347   0.05992  -0.0534  0.0113  1.0000
  12.500   1.2729   0.07481   0.07170  -0.0539  0.0114  1.0000
  13.000   1.2239   0.08960   0.08686  -0.0599  0.0116  1.0000