XFOIL         Version 6.94
  
 Calculated polar for: E64  (8.45%)                                    
  
 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   0.4554   0.00764   0.00331  -0.1110  0.8933  1.0000
   0.500   0.5186   0.00748   0.00302  -0.1121  0.8759  1.0000
   1.000   0.5788   0.00740   0.00282  -0.1125  0.8544  1.0000
   1.500   0.6348   0.00741   0.00274  -0.1121  0.8277  1.0000
   2.000   0.6900   0.00749   0.00274  -0.1115  0.7970  1.0000
   2.500   0.7440   0.00765   0.00278  -0.1107  0.7617  1.0000
   3.000   0.7962   0.00790   0.00294  -0.1095  0.7208  1.0000
   3.500   0.8464   0.00825   0.00315  -0.1080  0.6712  1.0000
   4.000   0.8945   0.00874   0.00348  -0.1061  0.6124  1.0000
   4.500   0.9375   0.00950   0.00389  -0.1033  0.5237  1.0000
   5.000   0.9759   0.01065   0.00452  -0.1000  0.4060  1.0000
   5.500   1.0150   0.01200   0.00534  -0.0972  0.2932  1.0000
   6.000   1.0549   0.01344   0.00634  -0.0947  0.1942  1.0000
   6.500   1.0911   0.01538   0.00767  -0.0919  0.0849  1.0000
   7.000   1.1241   0.01782   0.00977  -0.0880  0.0256  1.0000
   7.500   1.1621   0.01956   0.01170  -0.0849  0.0215  1.0000
   8.000   1.1880   0.02247   0.01482  -0.0800  0.0196  1.0000
   8.500   1.2194   0.02521   0.01781  -0.0762  0.0190  1.0000
   9.000   1.2541   0.02778   0.02064  -0.0730  0.0177  1.0000
   9.500   1.2901   0.03170   0.02493  -0.0702  0.0173  1.0000
  10.000   1.3219   0.03687   0.03063  -0.0671  0.0172  1.0000
  10.500   1.3342   0.04333   0.03780  -0.0617  0.0176  1.0000
  11.000   1.3213   0.05048   0.04558  -0.0543  0.0183  1.0000
  11.500   1.2897   0.05934   0.05499  -0.0479  0.0191  1.0000