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 =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.3999   0.01053   0.00270  -0.0723  0.5508  0.0493
   0.500   0.4448   0.01053   0.00250  -0.0698  0.5156  0.0572
   1.000   0.4856   0.00998   0.00244  -0.0668  0.4899  0.2706
   1.500   0.5320   0.00999   0.00258  -0.0648  0.4660  0.3589
   2.500   0.7652   0.00965   0.00342  -0.0919  0.4125  1.0000
   3.000   0.8058   0.00995   0.00347  -0.0888  0.3720  1.0000
   3.500   0.8487   0.01024   0.00366  -0.0862  0.3539  1.0000
   4.000   0.8878   0.01068   0.00383  -0.0829  0.3101  1.0000
   4.500   0.9293   0.01104   0.00407  -0.0801  0.2866  1.0000
   5.000   0.9703   0.01141   0.00441  -0.0771  0.2608  1.0000
   6.000   1.0073   0.01514   0.00671  -0.0636  0.0072  1.0000
   6.500   1.0429   0.01577   0.00747  -0.0596  0.0049  1.0000
   7.000   1.0772   0.01646   0.00828  -0.0554  0.0049  1.0000
   7.500   1.1076   0.01735   0.00936  -0.0505  0.0049  1.0000
   8.000   1.1296   0.01842   0.01071  -0.0439  0.0051  1.0000
   8.500   1.1431   0.02002   0.01256  -0.0363  0.0056  1.0000
   9.000   1.1478   0.02221   0.01499  -0.0282  0.0059  1.0000
   9.500   1.1693   0.02368   0.01657  -0.0231  0.0067  1.0000
  10.000   1.1823   0.02581   0.01888  -0.0174  0.0079  1.0000
  10.500   1.1789   0.02936   0.02270  -0.0107  0.0103  1.0000