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Ornithopter airfoil. (s1020-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: Ornithopter airfoil. (s1020-il)
Reynolds number: 50,000
Max Cl/Cd: 5.99 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s1020-il-50000.txt
Download as CSV file: xf-s1020-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Ornithopter 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3908   0.13038   0.12471  -0.0196   1.0000   0.2467
  -8.000  -0.3798   0.12659   0.12094  -0.0177   1.0000   0.2523
  -7.750  -0.3937   0.12576   0.12020  -0.0160   1.0000   0.2613
  -7.500  -0.3981   0.12329   0.11781  -0.0144   1.0000   0.2665
  -7.250  -0.4036   0.12182   0.11639  -0.0124   1.0000   0.2763
  -7.000  -0.4133   0.11977   0.11443  -0.0107   1.0000   0.2820
  -6.750  -0.4213   0.11842   0.11314  -0.0085   1.0000   0.2925
  -6.500  -0.4221   0.11581   0.11058  -0.0066   1.0000   0.2996
  -6.250  -0.4470   0.11551   0.11039  -0.0040   1.0000   0.3093
  -6.000  -0.4365   0.11213   0.10703  -0.0023   1.0000   0.3168
  -5.750  -0.4827   0.11299   0.10806   0.0016   1.0000   0.3259
  -5.500  -0.4560   0.10859   0.10363   0.0025   1.0000   0.3337
  -5.250  -0.4932   0.10825   0.10344   0.0062   1.0000   0.3433
  -5.000  -0.4811   0.10510   0.10029   0.0080   1.0000   0.3507
  -4.750  -0.5155   0.06930   0.06340  -0.0480   1.0000   0.1378
  -4.500  -0.4872   0.06095   0.05414  -0.0559   1.0000   0.1234
  -4.250  -0.4678   0.05740   0.05042  -0.0573   1.0000   0.1212
  -4.000  -0.4427   0.05345   0.04597  -0.0599   1.0000   0.1179
  -3.750  -0.4149   0.04997   0.04186  -0.0623   1.0000   0.1148
  -3.500  -0.3877   0.04735   0.03865  -0.0638   1.0000   0.1139
  -3.250  -0.3627   0.04552   0.03641  -0.0645   1.0000   0.1147
  -3.000  -0.3388   0.04421   0.03477  -0.0648   1.0000   0.1179
  -2.750  -0.3141   0.04328   0.03331  -0.0650   1.0000   0.1239
  -2.500  -0.2931   0.04235   0.03249  -0.0648   1.0000   0.1312
  -2.250  -0.2701   0.04163   0.03161  -0.0644   1.0000   0.1412
  -2.000  -0.2482   0.04100   0.03109  -0.0639   1.0000   0.1593
  -1.750  -0.2217   0.04016   0.03057  -0.0644   1.0000   0.2147
  -1.500  -0.1993   0.03960   0.03215  -0.0626   1.0000   0.5332
  -1.250  -0.1938   0.04053   0.03329  -0.0575   1.0000   0.6168
  -1.000  -0.1883   0.04122   0.03406  -0.0526   1.0000   0.6720
  -0.750  -0.1828   0.04175   0.03460  -0.0478   1.0000   0.7220
  -0.500  -0.1775   0.04211   0.03495  -0.0432   1.0000   0.7682
  -0.250  -0.1727   0.04229   0.03512  -0.0387   1.0000   0.8144
   0.000  -0.1682   0.04229   0.03511  -0.0343   1.0000   0.8649
   0.250  -0.1558   0.04197   0.03487  -0.0323   1.0000   0.9366
   0.500  -0.1335   0.04158   0.03433  -0.0355   1.0000   1.0000
   0.750  -0.0963   0.04291   0.03526  -0.0416   1.0000   1.0000
   1.000  -0.0543   0.04486   0.03679  -0.0480   0.9966   1.0000
   1.250  -0.0155   0.04691   0.03847  -0.0533   0.9922   1.0000
   1.500   0.0254   0.04937   0.04056  -0.0587   0.9857   1.0000
   1.750   0.0527   0.05085   0.04179  -0.0613   0.9795   1.0000
   2.000   0.0904   0.05355   0.04419  -0.0656   0.9735   1.0000
   2.250   0.1132   0.05472   0.04518  -0.0673   0.9652   1.0000
   2.500   0.1395   0.05657   0.04684  -0.0694   0.9588   1.0000
   2.750   0.1732   0.05898   0.04906  -0.0727   0.9502   1.0000
   3.000   0.1920   0.06012   0.05007  -0.0735   0.9410   1.0000
   3.250   0.2175   0.06216   0.05198  -0.0754   0.9340   1.0000
   3.500   0.2485   0.06443   0.05412  -0.0781   0.9238   1.0000
   3.750   0.2641   0.06556   0.05517  -0.0783   0.9139   1.0000
   4.000   0.2907   0.06793   0.05744  -0.0804   0.9067   1.0000
   4.250   0.3175   0.06993   0.05936  -0.0823   0.8952   1.0000
   4.500   0.3308   0.07116   0.06056  -0.0822   0.8853   1.0000
   4.750   0.3592   0.07382   0.06315  -0.0844   0.8775   1.0000
   5.000   0.3818   0.07563   0.06492  -0.0856   0.8655   1.0000
   5.250   0.3935   0.07700   0.06628  -0.0854   0.8553   1.0000
   5.500   0.4205   0.07970   0.06895  -0.0874   0.8470   1.0000
   5.750   0.4430   0.08168   0.07093  -0.0885   0.8349   1.0000
   6.000   0.4526   0.08311   0.07236  -0.0881   0.8242   1.0000
   6.250   0.4752   0.08567   0.07492  -0.0895   0.8156   1.0000
   6.500   0.5027   0.08823   0.07750  -0.0913   0.8037   1.0000
   6.750   0.5098   0.08959   0.07889  -0.0906   0.7919   1.0000
   7.000   0.5249   0.09185   0.08118  -0.0911   0.7825   1.0000
   7.250   0.5586   0.09520   0.08457  -0.0937   0.7719   1.0000
   7.500   0.5659   0.09658   0.08600  -0.0931   0.7592   1.0000
   7.750   0.5745   0.09854   0.08801  -0.0929   0.7483   1.0000
   8.000   0.5953   0.10139   0.09092  -0.0941   0.7390   1.0000
   8.250   0.6265   0.10459   0.09420  -0.0962   0.7263   1.0000
   8.500   0.6281   0.10598   0.09565  -0.0952   0.7136   1.0000
   8.750   0.6356   0.10821   0.09795  -0.0951   0.7026   1.0000
   9.000   0.6541   0.11112   0.10093  -0.0961   0.6926   1.0000
   9.250   0.6833   0.11443   0.10433  -0.0978   0.6800   1.0000
   9.500   0.6891   0.11627   0.10628  -0.0975   0.6669   1.0000
   9.750   0.6928   0.11850   0.10858  -0.0972   0.6552   1.0000
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