XFOIL         Version 6.94
  
 Calculated polar for: NASA/LANGLEY NLF 0414F 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   0.1797   0.02630   0.02153  -0.0531  0.8093  0.8505
   1.000   0.2812   0.02613   0.02127  -0.0532  0.7916  0.8628
   1.500   0.3512   0.02623   0.02133  -0.0554  0.7899  0.8652
   2.500   0.4235   0.02605   0.02123  -0.0516  0.7618  0.8713
   3.000   0.4928   0.02547   0.02070  -0.0537  0.7559  0.8733
   4.000   0.6479   0.02166   0.01704  -0.0580  0.7372  0.8749
   4.500   0.7317   0.01847   0.01395  -0.0596  0.7307  0.8761
   5.500   0.8513   0.01558   0.01128  -0.0608  0.6538  0.8775
   6.000   0.8806   0.01506   0.00926  -0.0541  0.4208  0.8781
   6.500   0.8775   0.01881   0.01190  -0.0485  0.2235  0.8799
   7.000   0.8987   0.02122   0.01365  -0.0455  0.1210  0.8802
   7.500   0.9290   0.02312   0.01524  -0.0433  0.0791  0.8804
   8.000   0.9623   0.02483   0.01689  -0.0415  0.0641  0.8809
   8.500   0.9948   0.02660   0.01864  -0.0395  0.0563  0.8816
   9.500   1.0615   0.03027   0.02237  -0.0358  0.0470  0.8818
  10.000   1.0983   0.03203   0.02412  -0.0342  0.0439  0.8819
  11.000   1.1813   0.03568   0.02798  -0.0319  0.0390  0.8825
  11.500   1.2268   0.03792   0.03041  -0.0311  0.0369  0.8832
  12.000   1.2671   0.04049   0.03327  -0.0299  0.0353  0.8834
  12.500   1.3071   0.04331   0.03626  -0.0290  0.0341  0.8837
  13.000   1.3384   0.04764   0.04095  -0.0273  0.0330  0.8840
  13.500   1.3355   0.05261   0.04655  -0.0230  0.0322  0.8842