XFOIL         Version 6.94
  
 Calculated polar for: SC(2)-0714 Supercritical 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   0.4324   0.01625   0.01184  -0.1016  0.8956  0.6976
   0.500   0.5018   0.01539   0.01103  -0.1029  0.8753  0.7036
   1.000   0.5868   0.01427   0.00987  -0.1094  0.8384  0.7110
   1.500   0.6473   0.01386   0.00920  -0.1087  0.7422  0.7131
   2.000   0.6725   0.01563   0.00950  -0.1018  0.4607  0.7154
   2.500   0.7030   0.01751   0.01026  -0.0973  0.2426  0.7187
   3.000   0.7496   0.01875   0.01091  -0.0962  0.1426  0.7227
   3.500   0.8064   0.01963   0.01154  -0.0972  0.1092  0.7272
   4.000   0.8659   0.02043   0.01222  -0.0988  0.0945  0.7311
   4.500   0.9112   0.02120   0.01299  -0.0966  0.0852  0.7336
   6.000   1.0549   0.02439   0.01630  -0.0922  0.0672  0.7432
   6.500   1.1084   0.02565   0.01752  -0.0924  0.0626  0.7469
   7.500   1.2101   0.02881   0.02091  -0.0913  0.0552  0.7529
   8.000   1.2561   0.03077   0.02289  -0.0898  0.0524  0.7557
   9.000   1.3418   0.03533   0.02802  -0.0858  0.0478  0.7630
   9.500   1.3815   0.03747   0.03035  -0.0838  0.0459  0.7670
  10.000   1.4194   0.04018   0.03317  -0.0819  0.0444  0.7705
  10.500   1.4384   0.04494   0.03839  -0.0772  0.0432  0.7730
  11.000   1.4341   0.04841   0.04239  -0.0687  0.0425  0.7761
  11.500   1.4209   0.05341   0.04795  -0.0609  0.0418  0.7794
  12.000   1.3948   0.05974   0.05483  -0.0538  0.0412  0.7826
  12.500   1.3534   0.06821   0.06383  -0.0489  0.0408  0.7858
  13.000   1.2882   0.08105   0.07726  -0.0487  0.0408  0.7880
  13.500   1.1519   0.11217   0.10913  -0.0673  0.0420  0.7877