XFOIL         Version 6.94
  
 Calculated polar for: NACA 64-108 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   0.0809   0.00717   0.00262  -0.0195  0.8822  0.8891
   0.500   0.1287   0.00715   0.00260  -0.0171  0.8598  0.9153
   1.000   0.1791   0.00713   0.00259  -0.0153  0.8376  0.9433
   1.500   0.2431   0.00712   0.00260  -0.0167  0.8137  0.9672
   2.000   0.3154   0.00703   0.00246  -0.0195  0.7610  0.9867
   2.500   0.3773   0.00722   0.00222  -0.0205  0.6346  1.0000
   3.000   0.4080   0.00904   0.00254  -0.0167  0.2697  1.0000
   3.500   0.4490   0.01181   0.00428  -0.0150  0.0353  1.0000
   4.000   0.4982   0.01329   0.00582  -0.0140  0.0329  1.0000
   4.500   0.5475   0.01519   0.00786  -0.0128  0.0332  1.0000
   5.000   0.5993   0.01756   0.01040  -0.0118  0.0338  1.0000
   5.500   0.6499   0.01903   0.01204  -0.0116  0.0261  1.0000
   6.000   0.7010   0.02099   0.01428  -0.0110  0.0232  1.0000
   6.500   0.7443   0.02679   0.02071  -0.0098  0.0198  1.0000
   7.000   0.7917   0.02777   0.02228  -0.0086  0.0125  1.0000
   7.500   0.8236   0.03546   0.03097  -0.0059  0.0080  1.0000
   8.000   0.8345   0.04729   0.04373  -0.0027  0.0073  1.0000
   8.500   0.8278   0.05888   0.05591  -0.0014  0.0072  1.0000
   9.000   0.8010   0.06909   0.06643  -0.0019  0.0073  1.0000
   9.500   0.7640   0.08266   0.08011  -0.0118  0.0088  1.0000