XFOIL         Version 6.94
  
 Calculated polar for: AG10                                            
  
 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   0.1474   0.00479   0.00146  -0.0274  0.9965  1.0000
   0.500   0.2272   0.00481   0.00133  -0.0320  0.9267  1.0000
   1.000   0.2740   0.00526   0.00126  -0.0287  0.7839  1.0000
   1.500   0.3249   0.00604   0.00129  -0.0270  0.6131  1.0000
   2.000   0.3790   0.00678   0.00144  -0.0265  0.4758  1.0000
   2.500   0.4342   0.00737   0.00165  -0.0263  0.3878  1.0000
   3.000   0.4896   0.00791   0.00192  -0.0262  0.3203  1.0000
   3.500   0.5450   0.00844   0.00223  -0.0260  0.2603  1.0000
   4.000   0.6002   0.00902   0.00262  -0.0259  0.2030  1.0000
   4.500   0.6550   0.00970   0.00310  -0.0257  0.1495  1.0000
   5.000   0.7095   0.01047   0.00370  -0.0255  0.1011  1.0000
   5.500   0.7634   0.01145   0.00449  -0.0253  0.0568  1.0000
   6.000   0.8166   0.01279   0.00580  -0.0248  0.0296  1.0000
   7.000   0.9192   0.01649   0.00987  -0.0232  0.0173  1.0000
   7.500   0.9677   0.01886   0.01244  -0.0222  0.0151  1.0000
   8.500   1.0546   0.02695   0.02162  -0.0192  0.0133  1.0000
   9.000   1.0853   0.03370   0.02924  -0.0174  0.0131  1.0000
   9.500   1.0905   0.04434   0.04090  -0.0159  0.0135  1.0000
  10.500   0.9903   0.08878   0.08661  -0.0458  0.0157  1.0000