XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 546 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   0.2324   0.01449   0.00843  -0.0382  0.6670  0.7588
   0.500   0.2855   0.01419   0.00802  -0.0388  0.6529  0.7627
   1.000   0.3417   0.01401   0.00766  -0.0400  0.6405  0.7659
   1.500   0.3933   0.01373   0.00740  -0.0397  0.6270  0.7679
   2.000   0.4473   0.01369   0.00731  -0.0399  0.6151  0.7701
   2.500   0.5005   0.01364   0.00727  -0.0399  0.6024  0.7731
   3.000   0.5543   0.01363   0.00727  -0.0401  0.5898  0.7756
   3.500   0.6098   0.01366   0.00723  -0.0407  0.5765  0.7783
   4.000   0.6618   0.01359   0.00726  -0.0407  0.5626  0.7812
   4.500   0.7162   0.01366   0.00733  -0.0411  0.5491  0.7840
   5.000   0.7690   0.01369   0.00737  -0.0411  0.5342  0.7871
   5.500   0.8174   0.01366   0.00749  -0.0401  0.5181  0.7898
   6.000   0.8663   0.01376   0.00769  -0.0392  0.5015  0.7927
   6.500   0.9138   0.01387   0.00789  -0.0380  0.4829  0.7960
   7.000   0.9587   0.01399   0.00812  -0.0364  0.4612  0.7997
   7.500   1.0003   0.01416   0.00838  -0.0343  0.4347  0.8035
   8.000   1.0355   0.01446   0.00868  -0.0309  0.4035  0.8076
   8.500   1.0585   0.01485   0.00912  -0.0252  0.3630  0.8125
   9.000   1.0683   0.01568   0.00985  -0.0175  0.3133  0.8179
   9.500   1.0722   0.01710   0.01107  -0.0098  0.2592  0.8241
  10.000   1.0719   0.01885   0.01268  -0.0023  0.2144  0.8304
  10.500   1.0723   0.02089   0.01466   0.0042  0.1757  0.8382
  11.500   1.0700   0.02624   0.01995   0.0146  0.1137  0.8580
  12.000   1.0674   0.02952   0.02325   0.0187  0.0904  0.8719
  12.500   1.0627   0.03325   0.02705   0.0222  0.0724  0.8923
  13.000   1.0716   0.03768   0.03169   0.0216  0.0558  0.9570