XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0606 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   0.4468   0.00660   0.00271  -0.1005  0.9230  0.7605
   0.500   0.5047   0.00666   0.00258  -0.0995  0.8255  0.7767
   1.000   0.5481   0.00802   0.00282  -0.0959  0.5672  0.7940
   1.500   0.5940   0.00968   0.00337  -0.0940  0.3013  0.8091
   2.000   0.6429   0.01124   0.00404  -0.0930  0.1062  0.8237
   2.500   0.6952   0.01216   0.00475  -0.0920  0.0663  0.8382
   3.000   0.7474   0.01295   0.00553  -0.0909  0.0546  0.8539
   3.500   0.7988   0.01376   0.00635  -0.0897  0.0474  0.8692
   4.000   0.8490   0.01469   0.00731  -0.0882  0.0423  0.8855
   4.500   0.8979   0.01560   0.00830  -0.0865  0.0379  0.9031
   5.000   0.9442   0.01678   0.00962  -0.0842  0.0349  0.9228
   5.500   0.9862   0.01811   0.01098  -0.0812  0.0326  0.9517
   6.000   1.0375   0.01947   0.01259  -0.0801  0.0303  1.0000
   6.500   1.0903   0.02189   0.01500  -0.0800  0.0284  1.0000
   7.000   1.1413   0.02404   0.01762  -0.0790  0.0265  1.0000
   7.500   1.1891   0.02672   0.02057  -0.0778  0.0255  1.0000
   8.000   1.2302   0.03065   0.02483  -0.0759  0.0245  1.0000
   8.500   1.2532   0.03744   0.03266  -0.0713  0.0238  1.0000
   9.000   1.2286   0.05106   0.04764  -0.0632  0.0232  1.0000
   9.500   1.0051   0.04582   0.04320  -0.0317  0.0253  1.0000