XFOIL         Version 6.94
  
 Calculated polar for: USA 26 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   0.2731   0.01280   0.00316  -0.0509  0.4205  0.0590
   1.000   0.3627   0.01363   0.00272  -0.0463  0.2498  0.0458
   1.500   0.4259   0.01225   0.00293  -0.0471  0.4021  0.1463
   2.000   0.4766   0.01194   0.00326  -0.0459  0.4156  0.2920
   2.500   0.5231   0.01222   0.00323  -0.0439  0.3564  0.3054
   3.000   0.5862   0.01299   0.00348  -0.0445  0.3886  0.0700
   3.500   0.6231   0.01353   0.00346  -0.0412  0.2709  0.1420
   4.500   0.8452   0.01222   0.00492  -0.0642  0.3427  1.0000
   5.500   0.8850   0.01280   0.00526  -0.0488  0.2882  0.9890
   6.500   1.0114   0.01373   0.00601  -0.0524  0.2527  1.0000
   7.500   1.1006   0.01450   0.00734  -0.0473  0.2752  1.0000
   8.500   1.1573   0.01626   0.00842  -0.0376  0.1532  1.0000
   9.000   1.1624   0.01749   0.00816  -0.0265  0.1764  0.0443
  10.000   1.2004   0.02128   0.01127  -0.0159  0.0556  0.3254
  11.000   1.2870   0.02224   0.01305  -0.0119  0.0836  0.2949
  13.000   1.3740   0.02996   0.02115   0.0023  0.0501  0.1104
  14.000   1.3228   0.04336   0.03412   0.0083  0.0047  0.1148
  14.500   1.2989   0.05246   0.04354   0.0066  0.0004  0.1042
  15.000   1.3529   0.05065   0.04337   0.0074  0.0089  1.0000
  15.500   1.3091   0.06403   0.05548   0.0035  0.0035  0.1576
  16.000   1.3238   0.06760   0.06059   0.0032  0.0080  1.0000