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NASA SC(2)-0414 AIRFOIL (sc20414-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0414 AIRFOIL (sc20414-il)
Reynolds number: 50,000
Max Cl/Cd: 25.9 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20414-il-50000.txt
Download as CSV file: xf-sc20414-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0414 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6917   0.08109   0.07275  -0.0423   1.0000   0.1822
  -9.750  -0.7065   0.07550   0.06717  -0.0430   1.0000   0.1801
  -9.500  -0.7413   0.06961   0.06124  -0.0434   1.0000   0.1770
  -9.250  -0.7798   0.06356   0.05499  -0.0437   1.0000   0.1737
  -9.000  -0.8009   0.05852   0.04961  -0.0431   1.0000   0.1720
  -8.750  -0.8048   0.05456   0.04536  -0.0424   1.0000   0.1724
  -8.500  -0.8004   0.05130   0.04184  -0.0416   1.0000   0.1749
  -8.250  -0.7935   0.04812   0.03831  -0.0410   1.0000   0.1779
  -8.000  -0.7832   0.04500   0.03473  -0.0407   1.0000   0.1809
  -7.750  -0.7681   0.04215   0.03154  -0.0403   1.0000   0.1848
  -7.500  -0.7490   0.04023   0.02970  -0.0392   1.0000   0.1909
  -7.250  -0.7301   0.03807   0.02719  -0.0387   1.0000   0.1974
  -7.000  -0.7101   0.03611   0.02517  -0.0378   1.0000   0.2043
  -6.750  -0.6894   0.03449   0.02342  -0.0370   1.0000   0.2137
  -6.500  -0.6687   0.03294   0.02192  -0.0357   1.0000   0.2232
  -6.250  -0.6476   0.03154   0.02047  -0.0344   1.0000   0.2358
  -6.000  -0.6269   0.03030   0.01929  -0.0329   1.0000   0.2507
  -5.750  -0.6074   0.02920   0.01841  -0.0308   1.0000   0.2691
  -5.500  -0.5886   0.02810   0.01757  -0.0287   1.0000   0.2938
  -5.250  -0.5706   0.02681   0.01663  -0.0269   1.0000   0.3314
  -5.000  -0.5564   0.02532   0.01610  -0.0243   1.0000   0.4006
  -4.750  -0.5538   0.02708   0.01916  -0.0138   1.0000   0.5080
  -4.500  -0.5425   0.03093   0.02302  -0.0045   1.0000   0.5946
  -4.250  -0.5300   0.03420   0.02617   0.0041   1.0000   0.6347
  -4.000  -0.5161   0.03651   0.02834   0.0112   1.0000   0.6683
  -3.750  -0.5028   0.03797   0.02968   0.0170   1.0000   0.7009
  -3.500  -0.4808   0.03942   0.03101   0.0230   1.0000   0.7273
  -3.250  -0.4659   0.03986   0.03133   0.0273   1.0000   0.7574
  -3.000  -0.4403   0.04022   0.03155   0.0307   1.0000   0.7881
  -2.750  -0.4044   0.04017   0.03136   0.0321   1.0000   0.8215
  -2.250  -0.2150   0.03811   0.02875   0.0132   1.0000   0.9154
  -2.000  -0.1532   0.03661   0.02710   0.0061   1.0000   0.9422
  -1.750  -0.1034   0.03533   0.02573   0.0005   1.0000   0.9606
  -1.500  -0.0686   0.03443   0.02478  -0.0027   1.0000   0.9727
  -1.250  -0.0409   0.03379   0.02411  -0.0047   1.0000   0.9819
  -1.000  -0.0062   0.03312   0.02342  -0.0080   1.0000   0.9904
  -0.750   0.0235   0.03266   0.02296  -0.0105   1.0000   0.9974
  -0.500   0.0365   0.03249   0.02281  -0.0099   1.0000   1.0000
  -0.250   0.0387   0.03245   0.02280  -0.0073   1.0000   1.0000
   0.000   0.0409   0.03243   0.02281  -0.0047   1.0000   1.0000
   0.250   0.0430   0.03242   0.02284  -0.0021   1.0000   1.0000
   0.500   0.0452   0.03243   0.02290   0.0004   1.0000   1.0000
   0.750   0.0472   0.03246   0.02297   0.0030   1.0000   1.0000
   1.000   0.0491   0.03251   0.02306   0.0055   1.0000   1.0000
   1.250   0.0509   0.03257   0.02318   0.0080   1.0000   1.0000
   1.500   0.0524   0.03264   0.02331   0.0105   1.0000   1.0000
   1.750   0.0536   0.03274   0.02347   0.0129   1.0000   1.0000
   2.000   0.0544   0.03286   0.02366   0.0154   1.0000   1.0000
   2.250   0.0547   0.03300   0.02387   0.0178   1.0000   1.0000
   2.500   0.0545   0.03317   0.02413   0.0201   1.0000   1.0000
   2.750   0.0538   0.03337   0.02442   0.0224   1.0000   1.0000
   3.000   0.1659   0.03489   0.02618   0.0050   0.9635   1.0000
   3.250   0.2969   0.03488   0.02651  -0.0127   0.9124   1.0000
   3.500   0.4877   0.03102   0.02327  -0.0349   0.8486   1.0000
   3.750   0.5814   0.02654   0.01923  -0.0386   0.7740   1.0000
   4.000   0.6185   0.02388   0.01550  -0.0315   0.5183   1.0000
   4.250   0.6288   0.02585   0.01599  -0.0276   0.3845   1.0000
   4.500   0.6493   0.02723   0.01685  -0.0264   0.3297   1.0000
   4.750   0.6727   0.02834   0.01772  -0.0257   0.2957   1.0000
   5.000   0.6981   0.02949   0.01860  -0.0254   0.2723   1.0000
   5.250   0.7206   0.03061   0.01966  -0.0246   0.2541   1.0000
   5.500   0.7420   0.03178   0.02080  -0.0238   0.2398   1.0000
   5.750   0.7598   0.03298   0.02217  -0.0223   0.2281   1.0000
   6.000   0.7787   0.03442   0.02364  -0.0211   0.2189   1.0000
   6.250   0.7942   0.03574   0.02512  -0.0193   0.2102   1.0000
   6.500   0.8101   0.03742   0.02686  -0.0178   0.2036   1.0000
   6.750   0.8197   0.03902   0.02884  -0.0151   0.1980   1.0000
   7.000   0.8327   0.04056   0.03044  -0.0132   0.1921   1.0000
   7.250   0.8430   0.04260   0.03258  -0.0110   0.1879   1.0000
   7.500   0.8439   0.04455   0.03498  -0.0073   0.1854   1.0000
   7.750   0.8450   0.04669   0.03750  -0.0041   0.1826   1.0000
   8.000   0.8507   0.04908   0.04022  -0.0019   0.1793   1.0000
   8.250   0.8621   0.05174   0.04310  -0.0008   0.1762   1.0000
   8.500   0.8747   0.05500   0.04658  -0.0003   0.1744   1.0000
   8.750   0.8890   0.05888   0.05057  -0.0004   0.1723   1.0000
   9.000   0.8872   0.06311   0.05526   0.0006   0.1720   1.0000
   9.250   0.8854   0.06780   0.06029   0.0009   0.1722   1.0000
   9.500   0.8008   0.07711   0.07045   0.0017   0.1844   1.0000
   9.750   0.7926   0.08322   0.07667  -0.0002   0.1873   1.0000
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