XFOIL         Version 6.94
  
 Calculated polar for: NACA M18                                        
  
 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   0.2292   0.01991   0.01398  -0.0354  0.5791  0.0985
   0.500   0.2756   0.01617   0.01018  -0.0350  0.5633  0.1338
   1.000   0.3241   0.01483   0.00885  -0.0345  0.5342  0.1811
   2.000   0.4480   0.01142   0.00395  -0.0314  0.5054  0.0433
   2.500   0.5021   0.01116   0.00360  -0.0313  0.4994  0.0394
   3.000   0.5575   0.01094   0.00324  -0.0315  0.4966  0.0375
   3.500   0.6128   0.01080   0.00299  -0.0319  0.4890  0.0376
   4.000   0.6594   0.00981   0.00300  -0.0310  0.4736  0.4623
   4.500   0.7698   0.00931   0.00356  -0.0434  0.4378  0.9829
   5.000   0.8747   0.01017   0.00432  -0.0552  0.4032  1.0000
   6.000   0.9771   0.01036   0.00458  -0.0552  0.3983  1.0000
   6.500   1.0279   0.01046   0.00474  -0.0552  0.3952  1.0000
   7.000   1.0767   0.01062   0.00478  -0.0548  0.3710  1.0000
   7.500   1.1189   0.01135   0.00514  -0.0535  0.3142  1.0000
   8.000   1.1641   0.01186   0.00560  -0.0526  0.2898  1.0000
   8.500   1.1952   0.01328   0.00650  -0.0497  0.2165  1.0000
   9.000   1.2033   0.01604   0.00843  -0.0434  0.1219  1.0000
   9.500   1.2057   0.01853   0.01047  -0.0361  0.0545  1.0000
  10.000   1.2101   0.02005   0.01190  -0.0286  0.0389  1.0000
  11.000   1.1767   0.02907   0.02169  -0.0159  0.0131  1.0000
  11.500   1.1610   0.03546   0.02839  -0.0136  0.0131  1.0000
  12.000   1.1451   0.04207   0.03525  -0.0122  0.0133  1.0000
  12.500   1.1448   0.04716   0.04049  -0.0108  0.0138  1.0000
  13.000   1.1547   0.05133   0.04478  -0.0096  0.0147  1.0000
  13.500   1.1584   0.05590   0.04988  -0.0044  0.0196  1.0000