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S5010 (s5010-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: S5010 (s5010-il)
Reynolds number: 50,000
Max Cl/Cd: 24.09 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s5010-il-50000.txt
Download as CSV file: xf-s5010-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S5010                                           
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5061   0.11100   0.10502   0.0246   1.0000   0.2368
  -8.500  -0.5272   0.11028   0.10444   0.0209   1.0000   0.2459
  -8.250  -0.5129   0.10581   0.09999   0.0214   1.0000   0.2601
  -8.000  -0.5013   0.10177   0.09599   0.0215   1.0000   0.2742
  -7.750  -0.4934   0.09814   0.09242   0.0213   1.0000   0.2886
  -7.500  -0.4931   0.09519   0.08957   0.0207   1.0000   0.3040
  -7.250  -0.4691   0.09046   0.08485   0.0223   1.0000   0.3253
  -7.000  -0.4735   0.08825   0.08275   0.0221   1.0000   0.3479
  -6.750  -0.4551   0.08434   0.07888   0.0237   1.0000   0.3756
  -6.500  -0.4409   0.08065   0.07525   0.0246   1.0000   0.3991
  -6.250  -0.4345   0.07764   0.07232   0.0255   1.0000   0.4262
  -5.250  -0.4314   0.04771   0.04111  -0.0227   1.0000   0.1561
  -5.000  -0.4177   0.04345   0.03643  -0.0226   1.0000   0.1399
  -4.750  -0.4120   0.04059   0.03281  -0.0206   1.0000   0.1288
  -4.500  -0.4038   0.03806   0.03018  -0.0185   1.0000   0.1260
  -4.250  -0.3941   0.03576   0.02755  -0.0165   1.0000   0.1230
  -4.000  -0.3820   0.03376   0.02517  -0.0147   1.0000   0.1227
  -3.750  -0.3676   0.03202   0.02306  -0.0131   1.0000   0.1250
  -3.500  -0.3508   0.03039   0.02099  -0.0117   1.0000   0.1267
  -3.250  -0.3320   0.02887   0.01910  -0.0104   1.0000   0.1279
  -3.000  -0.3119   0.02740   0.01739  -0.0094   1.0000   0.1308
  -2.750  -0.2919   0.02625   0.01621  -0.0086   1.0000   0.1396
  -2.500  -0.2602   0.02508   0.01497  -0.0095   0.9961   0.1542
  -2.250  -0.2080   0.02349   0.01349  -0.0137   0.9850   0.1924
  -2.000  -0.0955   0.01918   0.01206  -0.0246   0.9843   1.0000
  -1.750  -0.0369   0.01964   0.01190  -0.0316   0.9692   1.0000
  -1.500   0.0190   0.02008   0.01189  -0.0380   0.9539   1.0000
  -1.250   0.0763   0.02049   0.01194  -0.0444   0.9391   1.0000
  -1.000   0.1291   0.02089   0.01204  -0.0497   0.9240   1.0000
  -0.750   0.1715   0.02135   0.01226  -0.0530   0.9081   1.0000
  -0.500   0.2070   0.02187   0.01259  -0.0549   0.8922   1.0000
  -0.250   0.2384   0.02244   0.01301  -0.0559   0.8765   1.0000
   0.000   0.2663   0.02305   0.01348  -0.0562   0.8611   1.0000
   0.250   0.2916   0.02370   0.01400  -0.0559   0.8459   1.0000
   0.500   0.3150   0.02437   0.01457  -0.0552   0.8308   1.0000
   0.750   0.3371   0.02505   0.01517  -0.0543   0.8158   1.0000
   1.000   0.3582   0.02575   0.01579  -0.0531   0.8010   1.0000
   1.250   0.3787   0.02647   0.01644  -0.0518   0.7862   1.0000
   1.500   0.3989   0.02719   0.01709  -0.0504   0.7715   1.0000
   1.750   0.4189   0.02791   0.01777  -0.0490   0.7568   1.0000
   2.000   0.4389   0.02864   0.01846  -0.0474   0.7420   1.0000
   2.250   0.4588   0.02936   0.01916  -0.0459   0.7273   1.0000
   2.500   0.4787   0.03008   0.01987  -0.0444   0.7125   1.0000
   2.750   0.4989   0.03078   0.02055  -0.0428   0.6977   1.0000
   3.000   0.5190   0.03148   0.02125  -0.0412   0.6828   1.0000
   3.250   0.5395   0.03213   0.02194  -0.0395   0.6679   1.0000
   3.500   0.5601   0.03278   0.02259  -0.0378   0.6529   1.0000
   3.750   0.5810   0.03337   0.02320  -0.0361   0.6379   1.0000
   4.000   0.6022   0.03393   0.02379  -0.0343   0.6229   1.0000
   4.250   0.6215   0.03473   0.02467  -0.0329   0.6067   1.0000
   4.500   0.6398   0.03566   0.02567  -0.0316   0.5898   1.0000
   4.750   0.6596   0.03641   0.02648  -0.0301   0.5735   1.0000
   5.000   0.6803   0.03703   0.02716  -0.0284   0.5574   1.0000
   5.250   0.7016   0.03755   0.02779  -0.0267   0.5415   1.0000
   5.500   0.7239   0.03790   0.02821  -0.0248   0.5257   1.0000
   5.750   0.7468   0.03814   0.02853  -0.0228   0.5101   1.0000
   6.000   0.7704   0.03823   0.02869  -0.0207   0.4944   1.0000
   6.250   0.7950   0.03817   0.02873  -0.0185   0.4788   1.0000
   6.500   0.8209   0.03790   0.02852  -0.0161   0.4631   1.0000
   6.750   0.8367   0.03934   0.03011  -0.0152   0.4442   1.0000
   7.000   0.8571   0.03998   0.03088  -0.0136   0.4265   1.0000
   7.250   0.8814   0.04002   0.03103  -0.0116   0.4096   1.0000
   7.500   0.9090   0.03956   0.03060  -0.0093   0.3927   1.0000
   7.750   0.9355   0.03935   0.03041  -0.0072   0.3755   1.0000
   8.000   0.9484   0.04130   0.03258  -0.0064   0.3570   1.0000
   8.250   0.9685   0.04213   0.03352  -0.0049   0.3391   1.0000
   8.500   0.9954   0.04203   0.03345  -0.0030   0.3212   1.0000
   8.750   1.0203   0.04236   0.03377  -0.0012   0.3036   1.0000
   9.000   1.0319   0.04452   0.03615  -0.0003   0.2873   1.0000
   9.250   1.0445   0.04658   0.03838   0.0008   0.2720   1.0000
   9.500   1.0566   0.04856   0.04052   0.0019   0.2570   1.0000
   9.750   1.0685   0.05083   0.04296   0.0030   0.2439   1.0000
  10.000   1.0804   0.05290   0.04515   0.0041   0.2306   1.0000
  10.250   1.0027   0.06504   0.05760   0.0006   0.2376   1.0000
  10.500   0.9766   0.07159   0.06414  -0.0011   0.2345   1.0000
  11.000   1.0961   0.06469   0.05745   0.0076   0.1910   1.0000
  13.000   0.7446   0.15726   0.14911  -0.0491   0.3159   1.0000
  13.250   0.7689   0.16316   0.15514  -0.0490   0.3042   1.0000
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