XFOIL         Version 6.94
  
 Calculated polar for: RAE6-9CK AIRFOIL                                
  
 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   0.2207   0.00792   0.00293  -0.0680  0.8693  0.6335
   0.500   0.2772   0.00783   0.00288  -0.0680  0.8546  0.6405
   1.000   0.3334   0.00776   0.00285  -0.0678  0.8369  0.6482
   1.500   0.3879   0.00771   0.00279  -0.0671  0.8109  0.6568
   2.000   0.4399   0.00761   0.00268  -0.0657  0.7697  0.6653
   2.500   0.4926   0.00770   0.00274  -0.0648  0.7312  0.6746
   3.000   0.5421   0.00788   0.00283  -0.0631  0.6693  0.6839
   3.500   0.5764   0.00888   0.00313  -0.0586  0.4801  0.6947
   4.000   0.5999   0.01117   0.00407  -0.0533  0.1790  0.7056
   4.500   0.6384   0.01264   0.00497  -0.0504  0.0665  0.7176
   5.000   0.6830   0.01352   0.00586  -0.0484  0.0474  0.7300
   6.500   0.8098   0.01672   0.00928  -0.0414  0.0276  0.7779
   7.000   0.8518   0.01789   0.01062  -0.0390  0.0245  0.7971
   8.000   0.9340   0.02100   0.01420  -0.0341  0.0194  0.8488
   8.500   0.9708   0.02258   0.01608  -0.0308  0.0179  0.8914
   9.000   1.0126   0.02555   0.01958  -0.0288  0.0166  1.0000
   9.500   1.0484   0.02948   0.02404  -0.0263  0.0155  1.0000
  10.000   1.0750   0.03313   0.02806  -0.0228  0.0149  1.0000
  10.500   1.0775   0.03879   0.03432  -0.0164  0.0147  1.0000
  11.000   1.0636   0.04350   0.03941  -0.0085  0.0145  1.0000