XFOIL         Version 6.94
  
 Calculated polar for: AG16                                            
  
 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   0.2381   0.00504   0.00130  -0.0521  0.8325  1.0000
   0.500   0.2913   0.00530   0.00124  -0.0510  0.7713  1.0000
   1.000   0.3454   0.00562   0.00124  -0.0501  0.7095  1.0000
   1.500   0.3998   0.00597   0.00132  -0.0495  0.6477  1.0000
   2.000   0.4546   0.00635   0.00143  -0.0490  0.5859  1.0000
   2.500   0.5092   0.00677   0.00160  -0.0486  0.5236  1.0000
   3.000   0.5639   0.00724   0.00186  -0.0482  0.4596  1.0000
   3.500   0.6183   0.00777   0.00216  -0.0479  0.3931  1.0000
   4.000   0.6725   0.00838   0.00256  -0.0476  0.3262  1.0000
   4.500   0.7262   0.00908   0.00303  -0.0473  0.2607  1.0000
   5.000   0.7796   0.00985   0.00362  -0.0469  0.1988  1.0000
   5.500   0.8323   0.01075   0.00431  -0.0465  0.1400  1.0000
   6.000   0.8842   0.01183   0.00521  -0.0461  0.0850  1.0000
   6.500   0.9349   0.01315   0.00636  -0.0454  0.0428  1.0000
   7.000   0.9845   0.01473   0.00794  -0.0445  0.0210  1.0000
   7.500   1.0305   0.01712   0.01053  -0.0430  0.0120  1.0000
   8.000   1.0735   0.02002   0.01381  -0.0410  0.0098  1.0000
   8.500   1.1139   0.02344   0.01762  -0.0388  0.0089  1.0000
   9.000   1.1523   0.02683   0.02137  -0.0368  0.0080  1.0000
   9.500   1.1714   0.03384   0.02904  -0.0338  0.0070  1.0000
  10.000   1.1716   0.04247   0.03848  -0.0302  0.0070  1.0000
  10.500   1.1617   0.04874   0.04530  -0.0269  0.0068  1.0000
  11.500   1.0135   0.05747   0.05467  -0.0234  0.0071  1.0000
  12.000   0.9746   0.07235   0.06981  -0.0302  0.0072  1.0000
  12.500   0.9409   0.08727   0.08493  -0.0369  0.0073  1.0000
  13.000   0.8937   0.10555   0.10345  -0.0453  0.0077  1.0000