XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0406 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   0.2126   0.00761   0.00324  -0.0523  0.9392  0.7979
   0.500   0.2628   0.00757   0.00318  -0.0493  0.8867  0.8225
   1.000   0.3101   0.00764   0.00316  -0.0459  0.8194  0.8465
   1.500   0.3589   0.00781   0.00318  -0.0431  0.7420  0.8691
   2.000   0.4049   0.00832   0.00319  -0.0400  0.5936  0.8935
   2.500   0.4478   0.00946   0.00339  -0.0372  0.3665  0.9202
   3.000   0.4860   0.01068   0.00373  -0.0337  0.1622  0.9604
   3.500   0.5464   0.01198   0.00454  -0.0353  0.0899  1.0000
   4.000   0.6072   0.01305   0.00547  -0.0367  0.0713  1.0000
   4.500   0.6648   0.01433   0.00666  -0.0373  0.0630  1.0000
   5.000   0.7210   0.01563   0.00793  -0.0376  0.0567  1.0000
   5.500   0.7758   0.01727   0.00954  -0.0374  0.0516  1.0000
   6.000   0.8301   0.01899   0.01123  -0.0373  0.0474  1.0000
   6.500   0.8842   0.02076   0.01323  -0.0369  0.0442  1.0000
   7.000   0.9362   0.02319   0.01594  -0.0363  0.0410  1.0000
   7.500   0.9868   0.02574   0.01869  -0.0356  0.0387  1.0000
   8.000   1.0308   0.02987   0.02351  -0.0342  0.0366  1.0000
   8.500   1.0684   0.03476   0.02907  -0.0323  0.0349  1.0000
   9.000   1.0381   0.05470   0.05098  -0.0272  0.0395  1.0000
   9.500   1.0239   0.06362   0.06038  -0.0259  0.0368  1.0000