XFOIL         Version 6.94
  
 Calculated polar for: AG10                                            
  
 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   0.1411   0.00581   0.00173  -0.0263  1.0000  1.0000
   0.500   0.1960   0.00591   0.00176  -0.0260  0.9983  1.0000
   1.000   0.2844   0.00591   0.00168  -0.0323  0.9247  1.0000
   1.500   0.3322   0.00636   0.00163  -0.0288  0.7682  1.0000
   2.000   0.3812   0.00727   0.00172  -0.0267  0.5861  1.0000
   2.500   0.4344   0.00811   0.00195  -0.0260  0.4551  1.0000
   3.000   0.4889   0.00880   0.00228  -0.0257  0.3704  1.0000
   3.500   0.5437   0.00942   0.00266  -0.0254  0.3011  1.0000
   4.000   0.5984   0.01009   0.00313  -0.0252  0.2380  1.0000
   4.500   0.6529   0.01083   0.00368  -0.0249  0.1785  1.0000
   5.000   0.7069   0.01173   0.00440  -0.0247  0.1239  1.0000
   5.500   0.7603   0.01286   0.00536  -0.0244  0.0723  1.0000
   6.000   0.8125   0.01454   0.00698  -0.0238  0.0389  1.0000
   7.000   0.9125   0.01923   0.01205  -0.0218  0.0229  1.0000
   7.500   0.9587   0.02300   0.01610  -0.0204  0.0205  1.0000
   8.000   1.0011   0.02825   0.02211  -0.0188  0.0194  1.0000
   8.500   1.0371   0.03423   0.02902  -0.0172  0.0185  1.0000
   9.000   1.0510   0.04437   0.04021  -0.0159  0.0190  1.0000
   9.500   1.0484   0.05536   0.05178  -0.0157  0.0203  1.0000