XFOIL         Version 6.94
  
 Calculated polar for: NACA 16-012                                     
  
 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   0.0000   0.01317   0.00902   0.0000  0.9312  0.9312
   1.000   0.1598   0.01301   0.00893  -0.0117  0.9227  0.9338
   1.500   0.2634   0.01100   0.00695  -0.0205  0.9024  0.9325
   2.000   0.3326   0.01046   0.00645  -0.0230  0.8776  0.9347
   2.500   0.3782   0.01017   0.00618  -0.0206  0.8432  0.9425
   3.000   0.4387   0.01035   0.00587  -0.0211  0.6927  0.9444
   3.500   0.4435   0.01274   0.00650  -0.0109  0.2993  0.9539
   4.000   0.4893   0.01459   0.00730  -0.0102  0.0712  0.9571
   4.500   0.5317   0.01525   0.00792  -0.0079  0.0561  0.9640
   5.000   0.5916   0.01598   0.00870  -0.0096  0.0466  0.9665
   5.500   0.6338   0.01681   0.00963  -0.0073  0.0394  0.9729
   6.000   0.6917   0.01791   0.01080  -0.0087  0.0308  0.9752
   6.500   0.7423   0.01923   0.01225  -0.0084  0.0244  0.9791
   7.000   0.7912   0.02122   0.01436  -0.0079  0.0199  0.9825
   7.500   0.8473   0.02264   0.01609  -0.0086  0.0170  0.9853
   8.000   0.8942   0.02462   0.01829  -0.0077  0.0151  0.9892
   8.500   0.9364   0.02961   0.02404  -0.0061  0.0135  0.9914
   9.000   0.9543   0.03714   0.03258  -0.0005  0.0130  0.9959
   9.500   0.9330   0.04827   0.04472   0.0092  0.0135  1.0000
  10.000   0.8802   0.05335   0.05012   0.0250  0.0138  1.0000