XFOIL         Version 6.94
  
 Calculated polar for: S6061 9%                                        
  
 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   0.2004   0.00732   0.00272  -0.0381  0.8132  0.9731
   0.500   0.2881   0.00729   0.00252  -0.0446  0.7901  0.9917
   1.000   0.3527   0.00724   0.00234  -0.0467  0.7624  1.0000
   1.500   0.3961   0.00729   0.00228  -0.0443  0.7345  1.0000
   2.000   0.4418   0.00742   0.00228  -0.0423  0.7055  1.0000
   2.500   0.4892   0.00759   0.00237  -0.0405  0.6732  1.0000
   3.000   0.5383   0.00781   0.00251  -0.0391  0.6385  1.0000
   3.500   0.5884   0.00808   0.00270  -0.0378  0.5997  1.0000
   4.000   0.6389   0.00842   0.00298  -0.0367  0.5540  1.0000
   4.500   0.6869   0.00897   0.00325  -0.0352  0.4718  1.0000
   5.000   0.7342   0.00978   0.00375  -0.0339  0.3787  1.0000
   5.500   0.7800   0.01087   0.00444  -0.0326  0.2783  1.0000
   6.000   0.8242   0.01227   0.00535  -0.0312  0.1689  1.0000
   6.500   0.8601   0.01490   0.00721  -0.0288  0.0364  1.0000
   7.500   0.9400   0.01890   0.01160  -0.0240  0.0253  1.0000
   8.000   0.9772   0.02145   0.01436  -0.0214  0.0236  1.0000
   8.500   1.0167   0.02438   0.01758  -0.0190  0.0226  1.0000
   9.000   1.0559   0.02829   0.02193  -0.0167  0.0216  1.0000
   9.500   1.0899   0.03162   0.02562  -0.0143  0.0192  1.0000
  10.000   1.1127   0.03693   0.03149  -0.0109  0.0184  1.0000
  10.500   1.1088   0.04488   0.04028  -0.0054  0.0191  1.0000
  11.000   1.0657   0.05369   0.04979   0.0015  0.0206  1.0000
  11.500   1.0223   0.06278   0.05932   0.0025  0.0217  1.0000
  12.000   0.9799   0.07382   0.07070  -0.0022  0.0220  1.0000
  12.500   0.9359   0.08866   0.08580  -0.0122  0.0225  1.0000
  13.000   0.8877   0.10995   0.10727  -0.0273  0.0231  1.0000