XFOIL         Version 6.94
  
 Calculated polar for: NACA M2                                         
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.0537   0.01309   0.00202  -0.0002  0.0426  0.0420
   1.000   0.1074   0.01316   0.00207  -0.0005  0.0433  0.0426
   1.500   0.1610   0.01337   0.00229  -0.0006  0.0442  0.0449
   2.000   0.2140   0.01372   0.00281  -0.0007  0.0451  0.0469
   3.000   0.3003   0.01234   0.00461   0.0034  0.0471  0.9004
   3.500   0.3822   0.01408   0.00658  -0.0028  0.0486  0.9887
   4.000   0.4432   0.01660   0.00927  -0.0045  0.0506  1.0000
   4.500   0.4911   0.01724   0.00991  -0.0033  0.0528  1.0000
   5.000   0.5371   0.01973   0.01258  -0.0020  0.0519  1.0000
   5.500   0.5834   0.02065   0.01362  -0.0011  0.0461  1.0000
   6.000   0.6303   0.02219   0.01531  -0.0001  0.0430  1.0000
   6.500   0.6772   0.02302   0.01616   0.0005  0.0359  1.0000
   7.000   0.7180   0.02513   0.01832   0.0013  0.0228  1.0000
   7.500   0.7475   0.03391   0.02813   0.0056  0.0203  1.0000
   9.500   0.7039   0.08484   0.08125  -0.0023  0.0283  1.0000
  10.000   0.6913   0.09833   0.09472  -0.0113  0.0283  1.0000
  10.500   0.6845   0.10984   0.10621  -0.0170  0.0282  1.0000
  11.000   0.6814   0.12044   0.11679  -0.0214  0.0280  1.0000
  11.500   0.6799   0.13029   0.12661  -0.0249  0.0278  1.0000
  12.000   0.6815   0.14004   0.13635  -0.0282  0.0275  1.0000
  12.500   0.6852   0.14974   0.14603  -0.0314  0.0271  1.0000