XFOIL         Version 6.94
  
 Calculated polar for: AG44ct -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   0.2125   0.00654   0.00173  -0.0375  0.7589  1.0000
   0.500   0.2653   0.00686   0.00173  -0.0365  0.7218  1.0000
   1.000   0.3189   0.00714   0.00180  -0.0357  0.6864  1.0000
   1.500   0.3726   0.00733   0.00182  -0.0349  0.6478  1.0000
   2.000   0.4264   0.00754   0.00187  -0.0341  0.6019  1.0000
   2.500   0.4796   0.00786   0.00197  -0.0333  0.5485  1.0000
   3.000   0.5328   0.00828   0.00217  -0.0325  0.4900  1.0000
   3.500   0.5857   0.00878   0.00245  -0.0319  0.4318  1.0000
   4.000   0.6387   0.00935   0.00284  -0.0313  0.3777  1.0000
   4.500   0.6915   0.00996   0.00330  -0.0307  0.3281  1.0000
   5.000   0.7443   0.01061   0.00383  -0.0302  0.2818  1.0000
   5.500   0.7967   0.01131   0.00447  -0.0297  0.2384  1.0000
   6.000   0.8487   0.01209   0.00518  -0.0292  0.1966  1.0000
   6.500   0.9000   0.01297   0.00603  -0.0286  0.1557  1.0000
   7.000   0.9505   0.01400   0.00700  -0.0280  0.1149  1.0000
   7.500   0.9989   0.01539   0.00827  -0.0272  0.0737  1.0000
   8.000   1.0446   0.01732   0.01023  -0.0259  0.0425  1.0000
   9.500   1.1621   0.02573   0.01934  -0.0201  0.0185  1.0000
  11.000   1.2152   0.04185   0.03719  -0.0113  0.0139  1.0000
  11.500   1.1981   0.04851   0.04437  -0.0086  0.0136  1.0000
  12.000   1.1697   0.05832   0.05466  -0.0124  0.0136  1.0000
  12.500   1.1339   0.07194   0.06870  -0.0215  0.0137  1.0000
  13.000   1.0903   0.08955   0.08665  -0.0340  0.0141  1.0000
  13.500   1.0339   0.11203   0.10932  -0.0482  0.0151  1.0000