XFOIL         Version 6.94
  
 Calculated polar for: AG09                                            
  
 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   0.1543   0.00585   0.00175  -0.0294  0.9970  1.0000
   0.500   0.2413   0.00584   0.00162  -0.0357  0.9561  1.0000
   1.000   0.3025   0.00593   0.00152  -0.0357  0.8763  1.0000
   1.500   0.3502   0.00630   0.00150  -0.0327  0.7757  1.0000
   2.000   0.4006   0.00684   0.00157  -0.0308  0.6648  1.0000
   2.500   0.4527   0.00751   0.00175  -0.0297  0.5495  1.0000
   3.000   0.5057   0.00823   0.00205  -0.0291  0.4446  1.0000
   3.500   0.5593   0.00895   0.00243  -0.0286  0.3600  1.0000
   4.000   0.6131   0.00965   0.00292  -0.0283  0.2898  1.0000
   4.500   0.6669   0.01039   0.00348  -0.0279  0.2285  1.0000
   5.000   0.7203   0.01119   0.00414  -0.0276  0.1704  1.0000
   5.500   0.7731   0.01219   0.00495  -0.0273  0.1053  1.0000
   6.000   0.8235   0.01414   0.00660  -0.0266  0.0396  1.0000
   6.500   0.8730   0.01635   0.00895  -0.0256  0.0244  1.0000
   7.000   0.9195   0.01949   0.01234  -0.0240  0.0203  1.0000
   7.500   0.9662   0.02231   0.01538  -0.0227  0.0171  1.0000
   8.500   1.0397   0.03532   0.02989  -0.0188  0.0164  1.0000
   9.000   1.0376   0.04964   0.04587  -0.0163  0.0209  1.0000