XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0612 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
 ------- -------- --------- --------- -------- ------- -------
   1.000   0.5547   0.00908   0.00439  -0.1055  0.7227  0.7202
   1.500   0.5913   0.01109   0.00496  -0.1015  0.4021  0.7229
   2.000   0.6376   0.01273   0.00559  -0.1001  0.1897  0.7272
   2.500   0.6898   0.01376   0.00612  -0.0996  0.1091  0.7315
   3.000   0.7431   0.01441   0.00666  -0.0990  0.0886  0.7349
   3.500   0.7987   0.01501   0.00723  -0.0990  0.0778  0.7396
   4.000   0.8509   0.01573   0.00794  -0.0981  0.0691  0.7426
   4.500   0.9043   0.01633   0.00856  -0.0975  0.0616  0.7469
   6.000   1.0579   0.01911   0.01144  -0.0946  0.0386  0.7602
   6.500   1.1057   0.02024   0.01272  -0.0928  0.0360  0.7635
   7.000   1.1537   0.02157   0.01415  -0.0913  0.0342  0.7682
   7.500   1.2000   0.02297   0.01569  -0.0895  0.0328  0.7725
   8.000   1.2440   0.02480   0.01766  -0.0874  0.0314  0.7772
   8.500   1.2811   0.02909   0.02240  -0.0847  0.0297  0.7815
   9.000   1.3226   0.03048   0.02404  -0.0822  0.0290  0.7858
   9.500   1.3604   0.03249   0.02634  -0.0794  0.0279  0.7909
  10.000   1.3893   0.03526   0.02948  -0.0754  0.0264  0.7953
  10.500   1.4069   0.03912   0.03379  -0.0702  0.0253  0.8009
  11.000   1.4014   0.04371   0.03889  -0.0620  0.0247  0.8055
  11.500   1.3770   0.04996   0.04569  -0.0535  0.0244  0.8106
  12.000   1.3358   0.05823   0.05451  -0.0470  0.0243  0.8144
  12.500   1.2720   0.07109   0.06792  -0.0464  0.0244  0.8184