XFOIL         Version 6.94
  
 Calculated polar for: AH21 AIRFOIL                                    
  
 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   0.1366   0.00983   0.00545  -0.0432  0.9566  1.0000
   0.500   0.2250   0.00979   0.00534  -0.0500  0.9498  1.0000
   1.500   0.3811   0.00874   0.00433  -0.0581  0.9208  1.0000
   2.000   0.4704   0.00784   0.00354  -0.0640  0.9028  1.0000
   2.500   0.5444   0.00719   0.00301  -0.0666  0.8622  1.0000
   3.000   0.6164   0.00710   0.00264  -0.0687  0.7416  1.0000
   3.500   0.6372   0.00884   0.00297  -0.0609  0.4507  1.0000
   4.000   0.6651   0.01086   0.00375  -0.0561  0.1952  1.0000
   5.000   0.7499   0.01347   0.00550  -0.0517  0.0424  1.0000
   5.500   0.7960   0.01445   0.00648  -0.0500  0.0299  1.0000
   6.000   0.8429   0.01529   0.00733  -0.0486  0.0219  1.0000
   6.500   0.8884   0.01635   0.00853  -0.0468  0.0189  1.0000
   7.000   0.9330   0.01755   0.00992  -0.0449  0.0171  1.0000
   7.500   0.9795   0.01843   0.01100  -0.0435  0.0133  1.0000
   8.000   1.0246   0.01953   0.01233  -0.0418  0.0099  1.0000
   8.500   1.0670   0.02112   0.01424  -0.0396  0.0054  1.0000
   9.000   1.1035   0.02366   0.01716  -0.0365  0.0042  1.0000
   9.500   1.1321   0.02722   0.02132  -0.0324  0.0038  1.0000
  10.000   1.1453   0.03257   0.02750  -0.0266  0.0037  1.0000
  10.500   1.1208   0.04089   0.03685  -0.0168  0.0038  1.0000
  11.000   1.0643   0.05023   0.04694  -0.0068  0.0039  1.0000
  11.500   1.0075   0.06105   0.05822  -0.0050  0.0040  1.0000