XFOIL         Version 6.94
  
 Calculated polar for: NACA M12                                        
  
 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   0.1944   0.01122   0.00280  -0.0315  0.6998  0.0432
   0.500   0.2461   0.01092   0.00241  -0.0306  0.6786  0.0463
   1.500   0.3496   0.01067   0.00222  -0.0297  0.6068  0.1619
   2.000   0.4041   0.01075   0.00217  -0.0298  0.5889  0.1686
   3.500   0.5598   0.01083   0.00236  -0.0287  0.4954  0.3146
   4.000   0.6139   0.01103   0.00253  -0.0288  0.4791  0.2984
   4.500   0.6661   0.01139   0.00270  -0.0284  0.4462  0.2937
   5.000   0.7165   0.01189   0.00304  -0.0278  0.3990  0.2892
   5.500   0.7685   0.01225   0.00336  -0.0275  0.3688  0.2848
   6.000   0.8135   0.01312   0.00382  -0.0261  0.2883  0.2807
   6.500   0.8386   0.01596   0.00525  -0.0223  0.0928  0.2774
   7.000   0.8720   0.01824   0.00735  -0.0190  0.0494  0.2739
   7.500   0.9125   0.01955   0.00884  -0.0169  0.0487  0.2701
   8.000   0.9534   0.02069   0.01010  -0.0150  0.0468  0.2662
   8.500   0.9928   0.02183   0.01135  -0.0131  0.0438  0.2623
   9.000   1.0266   0.02339   0.01307  -0.0104  0.0425  0.2586
   9.500   1.0478   0.02526   0.01541  -0.0061  0.0286  0.2561
  10.000   1.0880   0.02589   0.01593  -0.0050  0.0196  0.2520
  10.500   1.1129   0.02772   0.01783  -0.0020  0.0172  0.2503
  11.000   1.1282   0.03040   0.02065   0.0014  0.0159  0.2493
  11.500   1.1402   0.03380   0.02426   0.0046  0.0151  0.2485
  12.000   1.1449   0.03805   0.02879   0.0074  0.0147  0.2479
  12.500   1.1491   0.04260   0.03365   0.0100  0.0146  0.2474
  13.000   1.1471   0.04829   0.03967   0.0116  0.0147  0.2467