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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.1875   0.00471   0.00130  -0.0359  0.9516  1.0000
   0.500   0.2421   0.00485   0.00117  -0.0345  0.8707  1.0000
   1.000   0.2915   0.00522   0.00111  -0.0322  0.7747  1.0000
   1.500   0.3436   0.00570   0.00114  -0.0309  0.6722  1.0000
   2.000   0.3970   0.00626   0.00124  -0.0301  0.5648  1.0000
   2.500   0.4510   0.00686   0.00142  -0.0297  0.4636  1.0000
   3.000   0.5054   0.00745   0.00170  -0.0294  0.3798  1.0000
   3.500   0.5600   0.00804   0.00204  -0.0291  0.3097  1.0000
   4.000   0.6145   0.00863   0.00243  -0.0289  0.2494  1.0000
   4.500   0.6688   0.00928   0.00292  -0.0287  0.1947  1.0000
   5.000   0.7227   0.01000   0.00347  -0.0284  0.1398  1.0000
   5.500   0.7757   0.01096   0.00421  -0.0281  0.0795  1.0000
   6.000   0.8268   0.01270   0.00570  -0.0275  0.0273  1.0000
   7.500   0.9729   0.01958   0.01330  -0.0240  0.0128  1.0000
   8.000   1.0174   0.02325   0.01744  -0.0223  0.0120  1.0000
   8.500   1.0596   0.02662   0.02118  -0.0210  0.0109  1.0000
   9.000   1.0910   0.03294   0.02828  -0.0189  0.0109  1.0000
   9.500   1.0685   0.04958   0.04653  -0.0165  0.0132  1.0000