XFOIL         Version 6.94
  
 Calculated polar for: S2050 8.93%                                     
  
 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   0.2446   0.00607   0.00187  -0.0599  0.8602  0.7872
   0.500   0.2986   0.00622   0.00188  -0.0587  0.8053  0.8086
   1.000   0.3501   0.00644   0.00194  -0.0571  0.7528  0.8290
   1.500   0.4012   0.00666   0.00204  -0.0556  0.7009  0.8485
   2.000   0.4520   0.00688   0.00213  -0.0542  0.6450  0.8664
   2.500   0.5015   0.00717   0.00227  -0.0525  0.5823  0.8860
   3.500   0.5925   0.00780   0.00261  -0.0475  0.4495  0.9405
   4.000   0.6518   0.00828   0.00286  -0.0484  0.3712  1.0000
   5.000   0.7578   0.00998   0.00385  -0.0484  0.2057  1.0000
   5.500   0.8086   0.01098   0.00453  -0.0479  0.1327  1.0000
   6.000   0.8584   0.01205   0.00534  -0.0472  0.0789  1.0000
   6.500   0.9053   0.01358   0.00662  -0.0458  0.0324  1.0000
   7.000   0.9526   0.01502   0.00820  -0.0443  0.0246  1.0000
   7.500   0.9960   0.01686   0.01017  -0.0424  0.0209  1.0000
   8.000   1.0365   0.01918   0.01273  -0.0399  0.0197  1.0000
   8.500   1.0786   0.02138   0.01519  -0.0378  0.0187  1.0000
   9.000   1.1185   0.02411   0.01821  -0.0355  0.0177  1.0000
   9.500   1.1566   0.02667   0.02103  -0.0333  0.0163  1.0000
  10.000   1.1883   0.03018   0.02489  -0.0305  0.0155  1.0000
  10.500   1.2101   0.03465   0.02978  -0.0271  0.0149  1.0000
  11.000   1.2124   0.04075   0.03641  -0.0222  0.0145  1.0000
  11.500   1.1824   0.04765   0.04383  -0.0152  0.0142  1.0000
  12.000   1.1494   0.05506   0.05166  -0.0126  0.0142  1.0000
  12.500   1.1063   0.06598   0.06299  -0.0162  0.0141  1.0000
  13.000   1.0747   0.07833   0.07565  -0.0242  0.0142  1.0000
  13.500   1.0440   0.09386   0.09148  -0.0358  0.0145  1.0000