XFOIL         Version 6.94
  
 Calculated polar for: AIRFOIL MH 20  9.02%                            
  
 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   0.2016   0.00813   0.00290  -0.0292  0.7248  1.0000
   0.500   0.2524   0.00814   0.00280  -0.0282  0.7112  1.0000
   1.000   0.3034   0.00818   0.00275  -0.0273  0.6974  1.0000
   1.500   0.3546   0.00824   0.00275  -0.0263  0.6829  1.0000
   2.000   0.4061   0.00831   0.00277  -0.0253  0.6674  1.0000
   2.500   0.4577   0.00838   0.00284  -0.0243  0.6503  1.0000
   3.000   0.5094   0.00845   0.00291  -0.0234  0.6307  1.0000
   3.500   0.5610   0.00852   0.00300  -0.0223  0.6080  1.0000
   4.000   0.6121   0.00851   0.00299  -0.0211  0.5640  1.0000
   4.500   0.6616   0.00880   0.00305  -0.0198  0.4811  1.0000
   5.000   0.7098   0.00958   0.00350  -0.0186  0.3905  1.0000
   5.500   0.7523   0.01129   0.00442  -0.0173  0.2204  1.0000
   6.000   0.7864   0.01426   0.00632  -0.0151  0.0285  1.0000
   6.500   0.8284   0.01579   0.00804  -0.0130  0.0223  1.0000
   7.000   0.8675   0.01747   0.00994  -0.0105  0.0210  1.0000
   7.500   0.9010   0.01964   0.01229  -0.0072  0.0202  1.0000
   8.000   0.9330   0.02229   0.01515  -0.0036  0.0201  1.0000
   8.500   0.9689   0.02568   0.01879  -0.0005  0.0207  1.0000
   9.000   1.0060   0.02937   0.02282   0.0021  0.0203  1.0000
   9.500   1.0358   0.03465   0.02862   0.0052  0.0210  1.0000
  10.000   0.9855   0.02378   0.01860   0.0130  0.0233  1.0000