XFOIL         Version 6.94
  
 Calculated polar for: AG03 (flat aft bottom)                          
  
 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   0.2154   0.00506   0.00131  -0.0376  0.8461  1.0000
   0.500   0.2666   0.00529   0.00120  -0.0360  0.7844  1.0000
   1.000   0.3188   0.00559   0.00117  -0.0348  0.7207  1.0000
   1.500   0.3718   0.00592   0.00120  -0.0339  0.6559  1.0000
   2.000   0.4251   0.00630   0.00129  -0.0332  0.5886  1.0000
   2.500   0.4784   0.00674   0.00145  -0.0326  0.5185  1.0000
   3.000   0.5318   0.00725   0.00166  -0.0321  0.4455  1.0000
   3.500   0.5851   0.00781   0.00197  -0.0317  0.3744  1.0000
   4.000   0.6384   0.00841   0.00233  -0.0313  0.3098  1.0000
   4.500   0.6916   0.00904   0.00278  -0.0310  0.2528  1.0000
   5.000   0.7446   0.00972   0.00330  -0.0306  0.2000  1.0000
   5.500   0.7971   0.01049   0.00394  -0.0302  0.1515  1.0000
   6.000   0.8489   0.01140   0.00470  -0.0298  0.1080  1.0000
   6.500   0.8999   0.01247   0.00570  -0.0292  0.0755  1.0000
   7.000   0.9499   0.01369   0.00693  -0.0285  0.0546  1.0000
   7.500   1.0002   0.01480   0.00816  -0.0277  0.0409  1.0000
   9.000   1.1353   0.02059   0.01445  -0.0239  0.0144  1.0000
   9.500   1.1788   0.02246   0.01658  -0.0226  0.0111  1.0000
  10.500   1.2364   0.03025   0.02514  -0.0175  0.0075  1.0000
  11.000   1.2587   0.03406   0.02951  -0.0149  0.0063  1.0000
  11.500   1.2652   0.03839   0.03421  -0.0117  0.0056  1.0000
  12.000   1.2453   0.04524   0.04145  -0.0100  0.0054  1.0000
  13.500   1.1563   0.08685   0.08421  -0.0367  0.0054  1.0000
  14.000   1.0840   0.11306   0.11084  -0.0514  0.0063  1.0000
  14.500   1.0271   0.13642   0.13431  -0.0631  0.0073  1.0000