XFOIL         Version 6.94
  
 Calculated polar for: S2055                                           
  
 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   0.2060   0.00599   0.00223  -0.0428  0.8741  0.9655
   1.000   0.3743   0.00605   0.00208  -0.0543  0.8299  0.9867
   1.500   0.4491   0.00607   0.00202  -0.0584  0.8001  0.9956
   2.000   0.5100   0.00611   0.00196  -0.0596  0.7629  1.0000
   2.500   0.5518   0.00622   0.00197  -0.0567  0.7153  1.0000
   3.000   0.5900   0.00654   0.00201  -0.0530  0.6346  1.0000
   3.500   0.6193   0.00746   0.00224  -0.0478  0.4744  1.0000
   4.000   0.6540   0.00839   0.00269  -0.0439  0.3564  1.0000
   4.500   0.6949   0.00916   0.00316  -0.0413  0.2817  1.0000
   5.000   0.7391   0.00992   0.00367  -0.0393  0.2173  1.0000
   5.500   0.7850   0.01070   0.00429  -0.0378  0.1600  1.0000
   6.000   0.8270   0.01206   0.00518  -0.0357  0.0700  1.0000
   6.500   0.8703   0.01342   0.00647  -0.0336  0.0372  1.0000
   7.000   0.9113   0.01512   0.00831  -0.0310  0.0285  1.0000
   7.500   0.9552   0.01641   0.00976  -0.0291  0.0236  1.0000
   8.500   1.0334   0.02043   0.01427  -0.0240  0.0154  1.0000
   9.000   1.0638   0.02393   0.01808  -0.0205  0.0111  1.0000
   9.500   1.1009   0.02571   0.02013  -0.0180  0.0090  1.0000
  10.000   1.1301   0.02805   0.02276  -0.0148  0.0076  1.0000
  10.500   1.1306   0.03388   0.02918  -0.0085  0.0067  1.0000
  11.000   1.0960   0.04149   0.03747   0.0004  0.0065  1.0000
  11.500   1.0727   0.04750   0.04394   0.0047  0.0064  1.0000
  12.000   1.0378   0.05627   0.05312   0.0047  0.0064  1.0000
  12.500   1.0099   0.06675   0.06397  -0.0006  0.0062  1.0000