XFOIL         Version 6.94
  
 Calculated polar for: AIRFOIL MH 43  8.5%                             
  
 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   0.1635   0.00595   0.00164  -0.0277  0.7193  0.9848
   0.500   0.2364   0.00600   0.00153  -0.0314  0.6975  1.0000
   1.000   0.2868   0.00610   0.00149  -0.0302  0.6762  1.0000
   1.500   0.3384   0.00621   0.00151  -0.0292  0.6534  1.0000
   2.000   0.3904   0.00635   0.00156  -0.0283  0.6289  1.0000
   2.500   0.4431   0.00651   0.00166  -0.0276  0.6023  1.0000
   3.000   0.4960   0.00671   0.00179  -0.0268  0.5711  1.0000
   3.500   0.5492   0.00695   0.00198  -0.0262  0.5355  1.0000
   4.000   0.6014   0.00734   0.00218  -0.0255  0.4725  1.0000
   4.500   0.6509   0.00822   0.00255  -0.0246  0.3512  1.0000
   5.000   0.7006   0.00920   0.00315  -0.0240  0.2489  1.0000
   5.500   0.7477   0.01066   0.00401  -0.0232  0.1228  1.0000
   6.000   0.7924   0.01261   0.00543  -0.0219  0.0232  1.0000
   6.500   0.8419   0.01367   0.00665  -0.0209  0.0195  1.0000
   7.000   0.8891   0.01500   0.00815  -0.0197  0.0177  1.0000
   7.500   0.9320   0.01690   0.01023  -0.0179  0.0164  1.0000
   8.000   0.9681   0.01993   0.01353  -0.0153  0.0155  1.0000
   8.500   1.0062   0.02346   0.01733  -0.0130  0.0154  1.0000
   9.000   1.0444   0.02755   0.02176  -0.0110  0.0154  1.0000
   9.500   1.0750   0.03288   0.02760  -0.0086  0.0153  1.0000
  10.000   1.1014   0.03710   0.03230  -0.0058  0.0156  1.0000
  11.500   0.8912   0.05303   0.05034   0.0036  0.0207  1.0000
  12.000   0.8418   0.06700   0.06459  -0.0028  0.0206  1.0000
  12.500   0.7886   0.08277   0.08061  -0.0125  0.0201  1.0000
  13.000   0.7139   0.10512   0.10317  -0.0281  0.0203  1.0000