XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 59 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   0.6386   0.01398   0.00801  -0.1695  0.9305  0.3222
   1.000   0.7927   0.01057   0.00606  -0.1761  0.9096  1.0000
   2.000   0.9491   0.00894   0.00280  -0.1810  0.5881  1.0000
   2.500   0.9759   0.01104   0.00369  -0.1753  0.3686  1.0000
   3.000   1.0129   0.01299   0.00480  -0.1722  0.1875  1.0000
   3.500   1.0512   0.01516   0.00602  -0.1694  0.0291  1.0000
   4.000   1.1009   0.01580   0.00675  -0.1682  0.0244  1.0000
   4.500   1.1499   0.01652   0.00759  -0.1668  0.0223  1.0000
   5.000   1.1996   0.01708   0.00831  -0.1656  0.0117  1.0000
   5.500   1.2458   0.01812   0.00929  -0.1636  0.0079  1.0000
   6.000   1.2909   0.01934   0.01079  -0.1612  0.0074  1.0000
   6.500   1.3312   0.02124   0.01320  -0.1577  0.0068  1.0000
   7.000   1.3618   0.02462   0.01705  -0.1526  0.0048  1.0000
   7.500   1.3971   0.02983   0.02288  -0.1481  0.0051  1.0000
   8.000   1.4341   0.03783   0.03179  -0.1439  0.0057  1.0000
   8.500   1.4584   0.04794   0.04292  -0.1371  0.0078  1.0000
   9.000   1.4396   0.06034   0.05633  -0.1262  0.0097  1.0000
   9.500   1.4032   0.06859   0.06511  -0.1154  0.0104  1.0000
  10.000   1.3616   0.07717   0.07413  -0.1076  0.0107  1.0000
  10.500   1.3195   0.08727   0.08463  -0.1044  0.0108  1.0000
  11.000   1.2766   0.09994   0.09764  -0.1072  0.0107  1.0000
  11.500   1.2406   0.11667   0.11470  -0.1178  0.0099  1.0000
  12.000   1.2149   0.14185   0.14000  -0.1357  0.0089  1.0000
  12.500   1.2033   0.15918   0.15729  -0.1462  0.0089  1.0000
  13.000   1.1975   0.17434   0.17238  -0.1550  0.0097  1.0000
  13.500   1.1982   0.18621   0.18422  -0.1618  0.0109  1.0000