XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 377 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.6830   0.01181   0.00519  -0.1240  0.5517  0.0128
   1.500   0.8491   0.01067   0.00353  -0.1230  0.5230  0.0294
   2.500   0.9586   0.01065   0.00345  -0.1219  0.5044  0.0788
   3.000   1.0132   0.01063   0.00349  -0.1214  0.4954  0.0951
   3.500   1.0669   0.01075   0.00358  -0.1208  0.4854  0.1124
   4.000   1.1210   0.01071   0.00368  -0.1203  0.4754  0.1381
   4.500   1.1743   0.01088   0.00386  -0.1198  0.4641  0.1699
   5.000   1.2278   0.01093   0.00406  -0.1192  0.4522  0.2116
   5.500   1.2807   0.01109   0.00447  -0.1187  0.4402  0.2786
   7.500   1.4717   0.01439   0.00718  -0.1155  0.1619  1.0000
   8.000   1.5091   0.01666   0.00900  -0.1136  0.0831  1.0000
   8.500   1.5343   0.02009   0.01189  -0.1104  0.0027  1.0000
   9.000   1.5715   0.02164   0.01362  -0.1080  0.0020  1.0000
   9.500   1.6030   0.02346   0.01570  -0.1050  0.0019  1.0000
  10.000   1.6241   0.02572   0.01823  -0.1009  0.0018  1.0000
  10.500   1.6231   0.02851   0.02128  -0.0942  0.0019  1.0000
  11.000   1.6170   0.03284   0.02588  -0.0901  0.0019  1.0000
  11.500   1.6093   0.03863   0.03196  -0.0881  0.0019  1.0000
  12.000   1.5978   0.04561   0.03920  -0.0872  0.0020  1.0000
  12.500   1.5828   0.05359   0.04744  -0.0871  0.0021  1.0000
  13.000   1.5665   0.06224   0.05633  -0.0877  0.0022  1.0000
  13.500   1.5517   0.07118   0.06552  -0.0887  0.0023  1.0000
  14.000   1.5410   0.08011   0.07468  -0.0899  0.0024  1.0000
  14.500   1.5351   0.08861   0.08340  -0.0908  0.0026  1.0000
  15.000   1.5384   0.09572   0.09069  -0.0903  0.0028  1.0000