XFOIL         Version 6.94
  
 Calculated polar for: S2048                                           
  
 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   0.2680   0.00782   0.00342  -0.0653  0.9033  0.8936
   0.500   0.3252   0.00745   0.00308  -0.0645  0.8857  0.9374
   1.000   0.4057   0.00719   0.00280  -0.0692  0.8687  0.9797
   1.500   0.4734   0.00707   0.00262  -0.0718  0.8452  1.0000
   2.000   0.5249   0.00712   0.00259  -0.0707  0.8156  1.0000
   2.500   0.5773   0.00724   0.00264  -0.0696  0.7816  1.0000
   3.000   0.6286   0.00742   0.00273  -0.0683  0.7373  1.0000
   3.500   0.6766   0.00777   0.00285  -0.0662  0.6583  1.0000
   4.500   0.7579   0.00997   0.00378  -0.0602  0.3753  1.0000
   5.000   0.8029   0.01103   0.00448  -0.0586  0.2907  1.0000
   5.500   0.8496   0.01198   0.00523  -0.0572  0.2253  1.0000
   6.000   0.8951   0.01313   0.00608  -0.0557  0.1429  1.0000
   6.500   0.9351   0.01503   0.00758  -0.0534  0.0717  1.0000
   7.000   0.9780   0.01657   0.00908  -0.0514  0.0483  1.0000
   7.500   1.0204   0.01815   0.01077  -0.0493  0.0321  1.0000
   8.000   1.0530   0.02109   0.01380  -0.0457  0.0205  1.0000
   8.500   1.0916   0.02319   0.01608  -0.0432  0.0155  1.0000
   9.000   1.1243   0.02724   0.02053  -0.0398  0.0137  1.0000
   9.500   1.1555   0.03182   0.02568  -0.0363  0.0128  1.0000
  10.000   1.1726   0.03776   0.03236  -0.0316  0.0125  1.0000
  10.500   1.1837   0.04118   0.03610  -0.0269  0.0113  1.0000
  11.000   1.1661   0.04733   0.04285  -0.0200  0.0113  1.0000
  11.500   1.1414   0.05377   0.04974  -0.0156  0.0111  1.0000
  12.000   1.0622   0.06983   0.06665  -0.0176  0.0125  1.0000
  12.500   1.0162   0.08556   0.08273  -0.0278  0.0127  1.0000
  13.000   0.9643   0.10888   0.10626  -0.0442  0.0141  1.0000