XFOIL         Version 6.94
  
 Calculated polar for: EPPLER 863 STRUT 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   0.0024   0.02477   0.01648  -0.0004  0.4510  0.4515
   0.500   0.0535   0.02500   0.01672  -0.0004  0.4477  0.4548
   1.000   0.1040   0.02462   0.01638  -0.0005  0.4442  0.4588
   1.500   0.1571   0.02461   0.01647  -0.0010  0.4404  0.4624
   2.000   0.2119   0.02471   0.01664  -0.0015  0.4367  0.4661
   2.500   0.2691   0.02473   0.01663  -0.0024  0.4331  0.4704
   3.000   0.3268   0.02489   0.01670  -0.0035  0.4298  0.4745
   3.500   0.3750   0.02570   0.01752  -0.0034  0.4260  0.4780
   4.000   0.4200   0.02604   0.01794  -0.0025  0.4224  0.4807
   4.500   0.4638   0.02598   0.01801  -0.0017  0.4182  0.4863
   5.000   0.5114   0.02623   0.01835  -0.0013  0.4145  0.4905
   5.500   0.5669   0.02628   0.01839  -0.0019  0.4112  0.4949
   6.000   0.6290   0.02642   0.01844  -0.0036  0.4078  0.4998
   6.500   0.6615   0.02751   0.01964  -0.0014  0.4040  0.5043
   7.000   0.6727   0.02849   0.02076   0.0040  0.3993  0.5073
   7.500   0.6798   0.02921   0.02162   0.0097  0.3949  0.5128
   8.000   0.7275   0.02943   0.02190   0.0098  0.3915  0.5186
   8.500   0.8005   0.02906   0.02145   0.0070  0.3885  0.5251
   9.000   0.7722   0.03270   0.02533   0.0140  0.3830  0.5295
   9.500   0.6797   0.04134   0.03424   0.0210  0.3725  0.5319
  10.000   0.7701   0.03896   0.03179   0.0185  0.3713  0.5401
  10.500   0.8655   0.03657   0.02939   0.0156  0.3697  0.5505
  11.000   0.6836   0.05459   0.04776   0.0220  0.3513  0.5499
  11.500   0.7739   0.05148   0.04463   0.0203  0.3510  0.5604