XFOIL         Version 6.94
  
 Calculated polar for: EPPLER E853 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   0.4649   0.00959   0.00349  -0.1175  0.7742  0.6549
   0.500   0.5217   0.00969   0.00358  -0.1178  0.7638  0.6720
   1.000   0.5756   0.00976   0.00377  -0.1176  0.7525  0.6909
   1.500   0.6320   0.00988   0.00395  -0.1178  0.7427  0.7126
   2.000   0.6850   0.00998   0.00422  -0.1173  0.7312  0.7380
   2.500   0.7390   0.01005   0.00446  -0.1168  0.7188  0.7678
   3.000   0.7894   0.01001   0.00462  -0.1154  0.7029  0.8076
   3.500   0.8345   0.00986   0.00472  -0.1126  0.6824  0.8643
   4.000   0.8838   0.00963   0.00467  -0.1106  0.6599  1.0000
   4.500   0.9370   0.00978   0.00489  -0.1102  0.6365  1.0000
   5.000   0.9851   0.00990   0.00511  -0.1085  0.6024  1.0000
   5.500   1.0283   0.01017   0.00537  -0.1057  0.5458  1.0000
   6.000   1.0510   0.01142   0.00597  -0.0992  0.3947  1.0000
   6.500   1.0549   0.01418   0.00763  -0.0905  0.1987  1.0000
   7.000   1.0660   0.01647   0.00927  -0.0833  0.0928  1.0000
   7.500   1.0835   0.01830   0.01080  -0.0772  0.0453  1.0000
   8.000   1.0958   0.02066   0.01296  -0.0704  0.0164  1.0000
   8.500   1.1077   0.02324   0.01569  -0.0639  0.0103  1.0000
   9.000   1.1187   0.02611   0.01868  -0.0582  0.0088  1.0000
   9.500   1.1355   0.02919   0.02198  -0.0536  0.0081  1.0000
  10.000   1.1600   0.03268   0.02572  -0.0502  0.0076  1.0000
  10.500   1.1932   0.03706   0.03052  -0.0479  0.0076  1.0000
  11.000   1.2154   0.04366   0.03789  -0.0447  0.0080  1.0000
  11.500   1.2120   0.05079   0.04565  -0.0404  0.0085  1.0000
  12.000   1.1977   0.05827   0.05356  -0.0372  0.0088  1.0000