XFOIL         Version 6.94
  
 Calculated polar for: AIRFOIL MH 18  11.14%                           
  
 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   0.3253   0.00959   0.00430  -0.0534  0.7196  1.0000
   0.500   0.3733   0.00957   0.00411  -0.0521  0.7116  1.0000
   1.000   0.4236   0.00957   0.00408  -0.0514  0.7012  1.0000
   1.500   0.4737   0.00959   0.00404  -0.0504  0.6916  1.0000
   2.000   0.5239   0.00958   0.00396  -0.0494  0.6812  1.0000
   2.500   0.5750   0.00962   0.00402  -0.0486  0.6689  1.0000
   3.000   0.6260   0.00965   0.00406  -0.0477  0.6563  1.0000
   3.500   0.6770   0.00968   0.00413  -0.0468  0.6424  1.0000
   4.000   0.7281   0.00971   0.00419  -0.0459  0.6264  1.0000
   4.500   0.7790   0.00972   0.00429  -0.0449  0.6066  1.0000
   5.000   0.8297   0.00975   0.00437  -0.0438  0.5834  1.0000
   5.500   0.8798   0.00986   0.00455  -0.0428  0.5510  1.0000
   6.000   0.9280   0.01015   0.00486  -0.0415  0.5050  1.0000
   6.500   0.9709   0.01095   0.00540  -0.0395  0.4313  1.0000
   7.000   1.0091   0.01213   0.00629  -0.0371  0.3547  1.0000
   7.500   1.0432   0.01346   0.00738  -0.0342  0.2830  1.0000
   8.000   1.0684   0.01520   0.00881  -0.0300  0.1972  1.0000
   8.500   1.0676   0.01816   0.01100  -0.0224  0.0775  1.0000
   9.000   1.0556   0.02089   0.01343  -0.0128  0.0346  1.0000
   9.500   1.0591   0.02347   0.01616  -0.0067  0.0256  1.0000