XFOIL         Version 6.94
  
 Calculated polar for: AG45ct -02f                                     
  
 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   0.1899   0.00652   0.00168  -0.0320  0.7486  1.0000
   0.500   0.2432   0.00686   0.00169  -0.0311  0.7118  1.0000
   1.000   0.2974   0.00710   0.00174  -0.0304  0.6749  1.0000
   1.500   0.3517   0.00728   0.00175  -0.0297  0.6341  1.0000
   2.000   0.4058   0.00752   0.00178  -0.0290  0.5876  1.0000
   2.500   0.4598   0.00784   0.00191  -0.0284  0.5359  1.0000
   3.000   0.5137   0.00824   0.00210  -0.0278  0.4819  1.0000
   3.500   0.5675   0.00871   0.00239  -0.0272  0.4277  1.0000
   4.000   0.6212   0.00924   0.00275  -0.0268  0.3748  1.0000
   4.500   0.6747   0.00982   0.00322  -0.0263  0.3224  1.0000
   5.000   0.7280   0.01048   0.00375  -0.0259  0.2712  1.0000
   5.500   0.7807   0.01125   0.00441  -0.0255  0.2187  1.0000
   6.000   0.8328   0.01217   0.00519  -0.0251  0.1686  1.0000
   6.500   0.8841   0.01323   0.00616  -0.0246  0.1249  1.0000
   7.000   0.9344   0.01450   0.00738  -0.0240  0.0859  1.0000
   7.500   0.9834   0.01603   0.00892  -0.0232  0.0538  1.0000
   8.000   1.0290   0.01824   0.01122  -0.0220  0.0348  1.0000
   8.500   1.0699   0.02132   0.01456  -0.0203  0.0257  1.0000
   9.000   1.1134   0.02361   0.01713  -0.0189  0.0204  1.0000
  10.000   1.1707   0.03421   0.02887  -0.0141  0.0160  1.0000
  10.500   1.1921   0.03888   0.03409  -0.0120  0.0142  1.0000
  11.000   1.1890   0.04570   0.04157  -0.0097  0.0136  1.0000
  11.500   1.1623   0.05412   0.05050  -0.0107  0.0135  1.0000
  12.000   1.1254   0.06756   0.06436  -0.0202  0.0137  1.0000
  12.500   1.0779   0.08634   0.08351  -0.0344  0.0143  1.0000
  13.000   1.0101   0.11255   0.10995  -0.0511  0.0159  1.0000