XFOIL         Version 6.94
  
 Calculated polar for: AG455ct -02f rot.                               
  
 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   0.1876   0.00632   0.00160  -0.0317  0.7633  1.0000
   0.500   0.2408   0.00671   0.00160  -0.0307  0.7171  1.0000
   1.000   0.2951   0.00701   0.00167  -0.0300  0.6767  1.0000
   1.500   0.3498   0.00721   0.00169  -0.0294  0.6320  1.0000
   2.000   0.4041   0.00749   0.00174  -0.0287  0.5815  1.0000
   2.500   0.4583   0.00784   0.00187  -0.0281  0.5263  1.0000
   3.000   0.5126   0.00826   0.00208  -0.0276  0.4684  1.0000
   3.500   0.5667   0.00875   0.00238  -0.0272  0.4096  1.0000
   4.000   0.6206   0.00931   0.00275  -0.0268  0.3504  1.0000
   4.500   0.6743   0.00996   0.00325  -0.0264  0.2930  1.0000
   5.000   0.7276   0.01071   0.00382  -0.0261  0.2376  1.0000
   5.500   0.7806   0.01153   0.00455  -0.0257  0.1865  1.0000
   6.000   0.8329   0.01250   0.00540  -0.0253  0.1425  1.0000
   6.500   0.8847   0.01356   0.00642  -0.0249  0.1051  1.0000
   7.000   0.9350   0.01493   0.00778  -0.0242  0.0734  1.0000
   7.500   0.9836   0.01667   0.00956  -0.0234  0.0488  1.0000
   9.000   1.1165   0.02412   0.01778  -0.0194  0.0203  1.0000
  10.000   1.1736   0.03515   0.03009  -0.0148  0.0158  1.0000
  10.500   1.1918   0.04040   0.03591  -0.0130  0.0141  1.0000
  11.000   1.1838   0.04763   0.04376  -0.0111  0.0136  1.0000
  11.500   1.1529   0.05743   0.05406  -0.0149  0.0136  1.0000
  12.000   1.1092   0.07407   0.07116  -0.0283  0.0140  1.0000
  12.500   1.0480   0.09818   0.09558  -0.0458  0.0152  1.0000
  13.500   0.9131   0.16005   0.15731  -0.0749  0.0250  1.0000