XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0614 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   0.4183   0.01340   0.00877  -0.0995  0.8723  0.7041
   0.500   0.4820   0.01292   0.00824  -0.1007  0.8450  0.7108
   1.000   0.5358   0.01276   0.00806  -0.0990  0.8083  0.7137
   1.500   0.5978   0.01254   0.00766  -0.0999  0.7421  0.7203
   2.000   0.6360   0.01361   0.00772  -0.0954  0.5210  0.7227
   2.500   0.6727   0.01559   0.00847  -0.0923  0.2809  0.7270
   3.000   0.7226   0.01687   0.00904  -0.0920  0.1621  0.7314
   3.500   0.7701   0.01789   0.00981  -0.0905  0.1210  0.7346
   4.000   0.8266   0.01871   0.01041  -0.0912  0.0999  0.7402
   4.500   0.8745   0.01953   0.01120  -0.0896  0.0880  0.7427
   5.000   0.9237   0.02051   0.01213  -0.0884  0.0799  0.7469
   6.000   1.0230   0.02253   0.01426  -0.0864  0.0681  0.7546
   6.500   1.0759   0.02344   0.01519  -0.0862  0.0623  0.7600
   7.500   1.1667   0.02600   0.01791  -0.0825  0.0528  0.7686
   9.000   1.2944   0.03182   0.02428  -0.0763  0.0427  0.7821
   9.500   1.3325   0.03451   0.02717  -0.0738  0.0414  0.7868
  10.000   1.3663   0.03893   0.03193  -0.0713  0.0402  0.7915
  10.500   1.3813   0.04387   0.03740  -0.0663  0.0393  0.7960
  11.000   1.3842   0.04752   0.04148  -0.0596  0.0391  0.8014
  11.500   1.3744   0.05253   0.04697  -0.0526  0.0390  0.8058
  12.000   1.3514   0.05900   0.05395  -0.0466  0.0389  0.8107
  12.500   1.3109   0.06754   0.06302  -0.0423  0.0390  0.8143
  13.000   1.2528   0.07987   0.07587  -0.0428  0.0392  0.8180