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NASA SC(2)-0503 AIRFOIL (sc20503-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0503 AIRFOIL (sc20503-il)
Reynolds number: 1,000,000
Max Cl/Cd: 42.82 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20503-il-1000000-n5.txt
Download as CSV file: xf-sc20503-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0503 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5313   0.08372   0.08220   0.0062   1.0000   0.0051
  -8.250  -0.5313   0.07955   0.07804   0.0048   1.0000   0.0051
  -8.000  -0.5317   0.07538   0.07388   0.0034   1.0000   0.0051
  -7.750  -0.5329   0.07124   0.06976   0.0017   1.0000   0.0051
  -7.500  -0.5341   0.06697   0.06551  -0.0007   1.0000   0.0051
  -7.000  -0.5797   0.06903   0.06740  -0.0118   1.0000   0.0051
  -4.000  -0.2665   0.01563   0.01130  -0.0409   1.0000   0.0045
  -3.750  -0.2331   0.01213   0.00736  -0.0416   1.0000   0.0034
  -3.500  -0.2011   0.00950   0.00443  -0.0419   1.0000   0.0027
  -3.250  -0.1691   0.00787   0.00256  -0.0426   1.0000   0.0023
  -3.000  -0.1395   0.00718   0.00173  -0.0430   1.0000   0.0022
  -2.750  -0.1112   0.00681   0.00125  -0.0430   1.0000   0.0023
  -2.500  -0.0837   0.00660   0.00096  -0.0429   1.0000   0.0026
  -2.250  -0.0569   0.00647   0.00082  -0.0427   1.0000   0.0034
  -2.000  -0.0302   0.00637   0.00075  -0.0425   1.0000   0.0044
  -1.750  -0.0019   0.00596   0.00067  -0.0429   1.0000   0.0894
  -1.500   0.0296   0.00506   0.00057  -0.0445   1.0000   0.3563
  -1.250   0.0606   0.00427   0.00056  -0.0459   1.0000   0.6056
  -1.000   0.0887   0.00401   0.00063  -0.0461   1.0000   0.7064
  -0.750   0.1191   0.00394   0.00065  -0.0468   0.9982   0.7314
  -0.500   0.1630   0.00501   0.00061  -0.0502   0.6290   0.7592
   0.000   0.2133   0.00721   0.00089  -0.0502   0.0118   0.7885
   0.250   0.2409   0.00732   0.00099  -0.0502   0.0040   0.8018
   0.500   0.2684   0.00742   0.00109  -0.0501   0.0030   0.8146
   0.750   0.2959   0.00755   0.00130  -0.0500   0.0024   0.8275
   1.250   0.3507   0.00819   0.00221  -0.0495   0.0022   0.8503
   1.500   0.3775   0.00895   0.00316  -0.0490   0.0023   0.8598
   1.750   0.4036   0.01065   0.00512  -0.0479   0.0027   0.8692
   2.000   0.4305   0.01311   0.00788  -0.0467   0.0034   0.8793
   2.250   0.4574   0.01632   0.01153  -0.0453   0.0045   0.8910
   4.250   0.6466   0.04491   0.04260  -0.0384   0.0050   1.0000
   4.500   0.6688   0.04921   0.04708  -0.0386   0.0050   1.0000
   4.750   0.6901   0.05370   0.05173  -0.0393   0.0050   1.0000
   5.000   0.7105   0.05837   0.05654  -0.0404   0.0050   1.0000
   5.250   0.7297   0.06319   0.06148  -0.0420   0.0050   1.0000
   5.500   0.7475   0.06816   0.06655  -0.0441   0.0049   1.0000
   5.750   0.7639   0.07324   0.07172  -0.0467   0.0049   1.0000
   6.000   0.7787   0.07839   0.07694  -0.0499   0.0049   1.0000
   6.250   0.7918   0.08359   0.08220  -0.0536   0.0049   1.0000
   6.500   0.8030   0.08879   0.08743  -0.0578   0.0048   1.0000
   6.750   0.8122   0.09395   0.09262  -0.0623   0.0048   1.0000
   7.000   0.8192   0.09899   0.09767  -0.0671   0.0048   1.0000
   7.250   0.8203   0.10347   0.10214  -0.0712   0.0048   1.0000
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