XFOIL         Version 6.94
  
 Calculated polar for: KC-135 BL351.6 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   0.2201   0.00673   0.00177  -0.0386  0.7871  0.7055
   0.500   0.2768   0.00620   0.00191  -0.0375  0.7680  0.9422
   1.000   0.3701   0.00638   0.00197  -0.0453  0.7514  0.9807
   1.500   0.4534   0.00645   0.00196  -0.0515  0.7318  0.9951
   2.000   0.5164   0.00648   0.00193  -0.0535  0.7119  1.0000
   2.500   0.5629   0.00656   0.00197  -0.0517  0.6902  1.0000
   3.000   0.6098   0.00667   0.00206  -0.0499  0.6641  1.0000
   3.500   0.6564   0.00684   0.00220  -0.0480  0.6290  1.0000
   4.000   0.6997   0.00720   0.00237  -0.0455  0.5557  1.0000
   4.500   0.7311   0.00852   0.00286  -0.0411  0.3703  1.0000
   5.000   0.7684   0.00970   0.00356  -0.0380  0.2636  1.0000
   5.500   0.8100   0.01059   0.00420  -0.0357  0.1997  1.0000
   6.000   0.8421   0.01244   0.00528  -0.0322  0.0553  1.0000
   6.500   0.8791   0.01406   0.00682  -0.0289  0.0279  1.0000
   7.000   0.9153   0.01575   0.00864  -0.0257  0.0212  1.0000
   7.500   0.9522   0.01738   0.01043  -0.0226  0.0186  1.0000
   8.000   0.9832   0.01989   0.01305  -0.0190  0.0165  1.0000
   8.500   1.0197   0.02251   0.01592  -0.0161  0.0152  1.0000
   9.000   1.0581   0.02564   0.01936  -0.0137  0.0146  1.0000
   9.500   1.0918   0.03002   0.02424  -0.0108  0.0142  1.0000
  10.000   1.1154   0.03406   0.02875  -0.0069  0.0133  1.0000
  10.500   1.1119   0.04172   0.03723  -0.0003  0.0142  1.0000
  11.000   1.0767   0.05182   0.04783   0.0072  0.0159  1.0000