XFOIL         Version 6.94
  
 Calculated polar for: NACA M20 AIRFOIL                                
  
 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   0.3773   0.01596   0.00813  -0.0522  0.5797  0.0340
   0.500   0.4340   0.01469   0.00637  -0.0521  0.5694  0.0366
   1.000   0.4894   0.01509   0.00689  -0.0528  0.5579  0.0526
   1.500   0.5457   0.01410   0.00560  -0.0526  0.5482  0.0539
   2.000   0.6012   0.01328   0.00477  -0.0525  0.5382  0.0552
   2.500   0.6556   0.01294   0.00440  -0.0524  0.5285  0.0611
   3.000   0.7108   0.01272   0.00434  -0.0526  0.5193  0.0781
   4.000   0.8505   0.01137   0.00470  -0.0596  0.5004  1.0000
   4.500   0.9003   0.01090   0.00397  -0.0584  0.4443  1.0000
   5.000   0.9511   0.01121   0.00412  -0.0579  0.4052  1.0000
   5.500   0.9999   0.01190   0.00451  -0.0572  0.3426  1.0000
   6.000   1.0238   0.01620   0.00721  -0.0550  0.0562  1.0000
   6.500   1.0652   0.01763   0.00853  -0.0534  0.0075  1.0000
   7.000   1.1091   0.01865   0.00969  -0.0520  0.0075  1.0000
   7.500   1.1486   0.02001   0.01128  -0.0502  0.0080  1.0000
   8.000   1.1825   0.02172   0.01325  -0.0477  0.0088  1.0000
   8.500   1.1967   0.02453   0.01636  -0.0434  0.0097  1.0000
   9.000   1.1908   0.02935   0.02139  -0.0400  0.0103  1.0000
   9.500   1.1899   0.03473   0.02704  -0.0380  0.0117  1.0000
  10.000   1.1812   0.04062   0.03298  -0.0351  0.0127  1.0000