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 =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500  -0.0143   0.02805   0.02234  -0.0317  0.8528  0.6759
   1.000   0.0692   0.02849   0.02268  -0.0377  0.8495  0.6763
   2.000   0.1648   0.02829   0.02245  -0.0368  0.8309  0.6785
   2.500   0.2423   0.02867   0.02287  -0.0415  0.8286  0.6798
   3.500   0.3434   0.02780   0.02216  -0.0404  0.8013  0.6834
   4.000   0.4861   0.02394   0.01847  -0.0532  0.7897  0.6852
   5.500   0.7488   0.01411   0.00806  -0.0629  0.5140  0.6919
   6.000   0.7043   0.01744   0.01023  -0.0464  0.2986  0.6941
   7.000   0.7420   0.02101   0.01277  -0.0365  0.1097  0.6987
   7.500   0.7751   0.02223   0.01393  -0.0337  0.0875  0.7017
   8.000   0.8094   0.02357   0.01526  -0.0311  0.0742  0.7050
   8.500   0.8481   0.02466   0.01643  -0.0293  0.0626  0.7091
   9.000   0.8864   0.02581   0.01763  -0.0277  0.0522  0.7137
   9.500   0.9239   0.02703   0.01888  -0.0261  0.0425  0.7178
  10.000   0.9604   0.02835   0.02039  -0.0242  0.0317  0.7226
  10.500   0.9952   0.03028   0.02246  -0.0221  0.0241  0.7280
  11.000   1.0326   0.03253   0.02491  -0.0204  0.0205  0.7343
  11.500   1.0745   0.03521   0.02775  -0.0196  0.0186  0.7411
  12.000   1.1108   0.03842   0.03153  -0.0179  0.0171  0.7495
  12.500   1.1370   0.04146   0.03476  -0.0158  0.0159  0.7598
  13.000   1.1433   0.04648   0.04057  -0.0114  0.0154  0.7715
  13.500   1.1303   0.05301   0.04787  -0.0061  0.0152  0.7861
  14.000   1.0992   0.06082   0.05643  -0.0011  0.0151  0.8071
  14.500   1.0576   0.06980   0.06614   0.0024  0.0152  0.8622