XFOIL         Version 6.94
  
 Calculated polar for: GOE 308 (MVA H.40) 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5002   0.01054   0.00280  -0.0773  0.6343  0.1291
   1.000   0.5457   0.01075   0.00284  -0.0750  0.5892  0.1548
   1.500   0.5920   0.01095   0.00291  -0.0730  0.5511  0.1742
   2.000   0.6381   0.01111   0.00302  -0.0711  0.5221  0.2094
   2.500   0.8078   0.01022   0.00356  -0.0965  0.4854  1.0000
   3.000   0.8513   0.01059   0.00376  -0.0940  0.4668  1.0000
   3.500   0.8881   0.01102   0.00389  -0.0902  0.4267  1.0000
   4.000   0.9239   0.01141   0.00414  -0.0862  0.3710  1.0000
   4.500   0.9613   0.01189   0.00442  -0.0826  0.3243  1.0000
   5.000   0.9916   0.01283   0.00489  -0.0779  0.2392  1.0000
   5.500   1.0147   0.01439   0.00586  -0.0721  0.1429  1.0000
   6.000   1.0263   0.01669   0.00749  -0.0643  0.0077  1.0000
   7.000   1.0927   0.01829   0.00941  -0.0555  0.0082  1.0000
   7.500   1.1194   0.01935   0.01071  -0.0499  0.0091  1.0000
   8.000   1.1366   0.02068   0.01230  -0.0428  0.0101  1.0000
   8.500   1.1457   0.02257   0.01445  -0.0350  0.0114  1.0000
   9.000   1.1460   0.02514   0.01723  -0.0269  0.0125  1.0000
   9.500   1.1379   0.02879   0.02106  -0.0193  0.0143  1.0000
  10.000   1.1434   0.03265   0.02497  -0.0135  0.0172  1.0000