XFOIL         Version 6.94
  
 Calculated polar for: NASA/LANGLEY NLF(2)-0415 AIRFOI                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.0637   0.02192   0.01686  -0.0534  0.8320  0.6458
   0.500   0.1276   0.02201   0.01689  -0.0557  0.8305  0.6468
   1.000   0.1963   0.02210   0.01695  -0.0589  0.8292  0.6478
   1.500   0.2645   0.02236   0.01722  -0.0620  0.8282  0.6489
   3.000   0.4098   0.02113   0.01607  -0.0586  0.7972  0.6534
   3.500   0.5471   0.01771   0.01280  -0.0721  0.7941  0.6545
   4.000   0.7887   0.01117   0.00644  -0.1060  0.7835  0.6559
   4.500   0.7694   0.01100   0.00638  -0.0900  0.7540  0.6581
   5.000   0.7451   0.01171   0.00599  -0.0731  0.5295  0.6601
   5.500   0.6991   0.01468   0.00803  -0.0554  0.3398  0.6620
   6.000   0.7021   0.01680   0.00938  -0.0474  0.1819  0.6641
   6.500   0.7307   0.01806   0.01033  -0.0438  0.1170  0.6664
   7.000   0.7650   0.01912   0.01126  -0.0410  0.0874  0.6689
   8.000   0.8376   0.02118   0.01327  -0.0363  0.0585  0.6741
  10.000   0.9895   0.02528   0.01765  -0.0289  0.0226  0.6892
  11.000   1.0531   0.02895   0.02156  -0.0239  0.0150  0.6977
  11.500   1.0834   0.03086   0.02367  -0.0214  0.0135  0.7039
  12.000   1.1109   0.03331   0.02637  -0.0188  0.0125  0.7104
  12.500   1.1378   0.03598   0.02933  -0.0163  0.0118  0.7176
  13.000   1.1605   0.03888   0.03258  -0.0136  0.0113  0.7268
  13.500   1.1766   0.04245   0.03648  -0.0106  0.0109  0.7380
  14.000   1.1827   0.04745   0.04191  -0.0074  0.0105  0.7521
  14.500   1.1737   0.05335   0.04840  -0.0036  0.0104  0.7722
  15.000   1.1524   0.06017   0.05591  -0.0001  0.0102  0.8132