XFOIL         Version 6.94
  
 Calculated polar for: AG14                                            
  
 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   0.2532   0.00469   0.00121  -0.0528  0.9208  1.0000
   0.500   0.3047   0.00491   0.00109  -0.0508  0.8352  1.0000
   1.000   0.3558   0.00533   0.00107  -0.0490  0.7417  1.0000
   1.500   0.4083   0.00583   0.00112  -0.0479  0.6469  1.0000
   2.000   0.4618   0.00636   0.00126  -0.0472  0.5568  1.0000
   2.500   0.5157   0.00690   0.00145  -0.0466  0.4736  1.0000
   3.000   0.5699   0.00745   0.00174  -0.0462  0.3978  1.0000
   3.500   0.6240   0.00804   0.00206  -0.0459  0.3277  1.0000
   4.000   0.6779   0.00867   0.00247  -0.0456  0.2617  1.0000
   4.500   0.7316   0.00939   0.00300  -0.0452  0.2014  1.0000
   5.000   0.7849   0.01018   0.00362  -0.0449  0.1463  1.0000
   5.500   0.8378   0.01107   0.00440  -0.0444  0.1018  1.0000
   6.000   0.8901   0.01208   0.00533  -0.0439  0.0683  1.0000
   6.500   0.9420   0.01315   0.00639  -0.0433  0.0417  1.0000
   7.500   1.0394   0.01712   0.01071  -0.0410  0.0127  1.0000
   8.000   1.0834   0.02014   0.01404  -0.0393  0.0099  1.0000
   9.000   1.1576   0.02920   0.02435  -0.0349  0.0075  1.0000
   9.500   1.1815   0.03553   0.03143  -0.0324  0.0069  1.0000
  10.000   1.1824   0.04423   0.04095  -0.0300  0.0067  1.0000
  10.500   1.1500   0.05421   0.05157  -0.0292  0.0067  1.0000
  11.000   1.1114   0.07064   0.06842  -0.0440  0.0070  1.0000
  11.500   1.0646   0.09507   0.09308  -0.0620  0.0075  1.0000
  12.000   1.0039   0.12090   0.11888  -0.0747  0.0085  1.0000
  12.500   0.9713   0.13907   0.13700  -0.0823  0.0091  1.0000