XFOIL         Version 6.94
  
 Calculated polar for: S7012 8.75%                                     
  
 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   0.2476   0.00699   0.00229  -0.0513  0.7757  0.9088
   0.500   0.3012   0.00703   0.00214  -0.0500  0.7331  0.9566
   1.000   0.3754   0.00716   0.00201  -0.0538  0.6900  1.0000
   1.500   0.4284   0.00744   0.00207  -0.0536  0.6517  1.0000
   2.000   0.4829   0.00775   0.00218  -0.0534  0.6157  1.0000
   2.500   0.5374   0.00807   0.00234  -0.0532  0.5814  1.0000
   3.000   0.5919   0.00841   0.00259  -0.0529  0.5491  1.0000
   3.500   0.6461   0.00877   0.00286  -0.0526  0.5157  1.0000
   4.000   0.7000   0.00915   0.00322  -0.0522  0.4816  1.0000
   4.500   0.7533   0.00958   0.00359  -0.0517  0.4446  1.0000
   5.000   0.8062   0.01002   0.00405  -0.0511  0.3995  1.0000
   5.500   0.8564   0.01073   0.00452  -0.0503  0.3214  1.0000
   6.000   0.9046   0.01181   0.00529  -0.0494  0.2334  1.0000
   6.500   0.9506   0.01329   0.00635  -0.0483  0.1391  1.0000
   7.000   0.9867   0.01636   0.00866  -0.0460  0.0218  1.0000
   8.500   1.0938   0.02532   0.01845  -0.0374  0.0132  1.0000
   9.000   1.1283   0.02959   0.02309  -0.0347  0.0131  1.0000
   9.500   1.1575   0.03585   0.02988  -0.0320  0.0134  1.0000
  10.000   1.1853   0.03954   0.03399  -0.0290  0.0139  1.0000
  10.500   1.1787   0.04772   0.04328  -0.0230  0.0165  1.0000
  11.000   1.1403   0.05607   0.05223  -0.0175  0.0180  1.0000
  11.500   1.0996   0.06559   0.06213  -0.0182  0.0187  1.0000
  12.000   1.0582   0.07778   0.07466  -0.0246  0.0191  1.0000
  12.500   1.0169   0.09318   0.09035  -0.0354  0.0192  1.0000
  13.000   0.9723   0.11378   0.11118  -0.0503  0.0190  1.0000