XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 550 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   0.3082   0.01150   0.00480  -0.0585  0.5274  0.6321
   0.500   0.3653   0.01142   0.00474  -0.0590  0.5218  0.6349
   1.000   0.4227   0.01146   0.00477  -0.0596  0.5165  0.6379
   1.500   0.4801   0.01157   0.00484  -0.0602  0.5114  0.6411
   2.000   0.5378   0.01183   0.00506  -0.0609  0.5056  0.6449
   2.500   0.5947   0.01190   0.00518  -0.0614  0.5013  0.6485
   3.000   0.6516   0.01198   0.00527  -0.0620  0.4959  0.6516
   3.500   0.7080   0.01203   0.00537  -0.0624  0.4905  0.6553
   4.000   0.7650   0.01239   0.00573  -0.0630  0.4838  0.6591
   4.500   0.8196   0.01242   0.00590  -0.0630  0.4787  0.6634
   5.000   0.8747   0.01252   0.00607  -0.0632  0.4720  0.6682
   5.500   0.9301   0.01266   0.00623  -0.0635  0.4651  0.6728
   6.000   0.9836   0.01284   0.00656  -0.0634  0.4577  0.6776
   6.500   1.0361   0.01287   0.00674  -0.0630  0.4494  0.6830
   7.000   1.0896   0.01311   0.00697  -0.0629  0.4398  0.6888
   7.500   1.1390   0.01310   0.00721  -0.0621  0.4306  0.6948
   8.000   1.1878   0.01322   0.00740  -0.0610  0.4184  0.7016
   8.500   1.2345   0.01330   0.00769  -0.0596  0.4042  0.7098
   9.000   1.2777   0.01347   0.00802  -0.0577  0.3862  0.7184
   9.500   1.3152   0.01382   0.00844  -0.0547  0.3607  0.7283
  10.000   1.3360   0.01446   0.00908  -0.0489  0.3257  0.7392
  10.500   1.3387   0.01575   0.01026  -0.0405  0.2847  0.7530
  11.000   1.3326   0.01756   0.01204  -0.0319  0.2474  0.7706
  11.500   1.3226   0.01991   0.01445  -0.0242  0.2167  0.7954
  12.000   1.3054   0.02307   0.01778  -0.0172  0.1905  0.8451