XFOIL         Version 6.94
  
 Calculated polar for: 74-130 WP2                                      
  
 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
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.3643   0.01869   0.01370  -0.0726  0.7528  0.9514
   1.000   0.3913   0.02008   0.01510  -0.0701  0.7434  0.9606
   1.500   0.4318   0.02102   0.01604  -0.0698  0.7346  0.9673
   2.000   0.4884   0.02135   0.01634  -0.0711  0.7307  0.9715
   2.500   0.5560   0.02098   0.01593  -0.0733  0.7279  0.9735
   3.000   0.6238   0.02081   0.01577  -0.0758  0.7256  0.9753
   3.500   0.6923   0.02070   0.01574  -0.0783  0.7232  0.9781
   4.000   0.7260   0.02249   0.01768  -0.0786  0.7053  0.9882
   4.500   0.7983   0.02169   0.01696  -0.0813  0.7015  0.9909
   5.000   0.8777   0.02001   0.01537  -0.0843  0.6983  0.9916
   5.500   0.9664   0.01717   0.01270  -0.0877  0.6943  0.9898
   6.500   1.1085   0.01287   0.00870  -0.0912  0.6560  1.0000
   7.000   1.1539   0.01231   0.00844  -0.0895  0.6172  1.0000
   7.500   1.1421   0.01430   0.00896  -0.0789  0.3944  1.0000
   8.000   1.1097   0.02005   0.01366  -0.0711  0.2243  1.0000
   8.500   1.1015   0.02465   0.01740  -0.0659  0.0929  1.0000
   9.000   1.1172   0.02765   0.02021  -0.0629  0.0497  1.0000
   9.500   1.1388   0.03029   0.02281  -0.0604  0.0297  1.0000
  10.000   1.1473   0.03398   0.02651  -0.0565  0.0119  1.0000
  11.500   1.1868   0.04603   0.03947  -0.0443  0.0080  1.0000
  12.000   1.2202   0.05076   0.04471  -0.0411  0.0082  1.0000
  12.500   1.2292   0.05801   0.05260  -0.0376  0.0086  1.0000
  13.000   1.2100   0.06729   0.06251  -0.0341  0.0091  1.0000
  13.500   1.1992   0.07542   0.07101  -0.0320  0.0100  1.0000