XFOIL         Version 6.94
  
 Calculated polar for: AG46c -03f                                      
  
 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   0.0732   0.00572   0.00184  -0.0292  0.8868  1.0000
   0.500   0.1233   0.00607   0.00162  -0.0270  0.7987  1.0000
   1.000   0.1761   0.00642   0.00157  -0.0259  0.7534  1.0000
   1.500   0.2299   0.00679   0.00161  -0.0251  0.7189  1.0000
   2.000   0.2846   0.00707   0.00172  -0.0245  0.6851  1.0000
   2.500   0.3395   0.00720   0.00173  -0.0240  0.6472  1.0000
   3.000   0.3943   0.00736   0.00176  -0.0234  0.5996  1.0000
   3.500   0.4487   0.00766   0.00184  -0.0228  0.5416  1.0000
   4.000   0.5029   0.00808   0.00205  -0.0223  0.4745  1.0000
   4.500   0.5568   0.00865   0.00235  -0.0219  0.3999  1.0000
   5.000   0.6105   0.00934   0.00276  -0.0216  0.3207  1.0000
   5.500   0.6639   0.01018   0.00335  -0.0214  0.2420  1.0000
   6.000   0.7168   0.01114   0.00407  -0.0212  0.1772  1.0000
   6.500   0.7697   0.01214   0.00498  -0.0210  0.1320  1.0000
   7.000   0.8218   0.01331   0.00606  -0.0206  0.0980  1.0000
   7.500   0.8735   0.01454   0.00736  -0.0202  0.0704  1.0000
   8.000   0.9230   0.01629   0.00913  -0.0196  0.0482  1.0000
   9.000   1.0151   0.02165   0.01497  -0.0174  0.0259  1.0000
   9.500   1.0624   0.02366   0.01739  -0.0164  0.0209  1.0000
  10.000   1.0976   0.02887   0.02309  -0.0148  0.0177  1.0000
  10.500   1.1187   0.03684   0.03203  -0.0130  0.0166  1.0000
  11.000   1.1356   0.04298   0.03890  -0.0118  0.0151  1.0000
  11.500   1.1239   0.05126   0.04785  -0.0117  0.0144  1.0000
  12.000   1.0848   0.06400   0.06106  -0.0209  0.0148  1.0000
  12.500   1.0236   0.08736   0.08472  -0.0399  0.0165  1.0000