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NASA SC(2)-1010 AIRFOIL (sc21010-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-1010 AIRFOIL (sc21010-il)
Reynolds number: 50,000
Max Cl/Cd: 32.06 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc21010-il-50000.txt
Download as CSV file: xf-sc21010-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-1010 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4401   0.10194   0.09448  -0.0159   0.9999   0.3453
 -10.250  -0.4227   0.09799   0.09051  -0.0148   0.9999   0.3547
 -10.000  -0.4254   0.09550   0.08810  -0.0133   0.9999   0.3642
  -9.750  -0.4550   0.09567   0.08847  -0.0100   0.9999   0.3755
  -9.500  -0.4294   0.09110   0.08382  -0.0092   0.9999   0.3879
  -9.250  -0.5520   0.05894   0.05141  -0.0574   0.9999   0.1603
  -9.000  -0.5313   0.05199   0.04398  -0.0640   0.9999   0.1540
  -8.750  -0.5019   0.04558   0.03697  -0.0706   0.9999   0.1505
  -8.500  -0.4673   0.04084   0.03153  -0.0760   0.9999   0.1539
  -8.250  -0.4279   0.03681   0.02659  -0.0810   0.9999   0.1577
  -8.000  -0.4010   0.03435   0.02414  -0.0818   0.9999   0.1642
  -7.750  -0.3661   0.03195   0.02122  -0.0842   0.9999   0.1744
  -7.500  -0.3370   0.03013   0.01931  -0.0849   0.9999   0.1856
  -7.250  -0.3066   0.02851   0.01754  -0.0857   0.9999   0.2027
  -7.000  -0.2780   0.02700   0.01612  -0.0859   0.9999   0.2240
  -6.750  -0.2491   0.02564   0.01498  -0.0862   0.9999   0.2586
  -6.500  -0.2156   0.02418   0.01404  -0.0875   0.9999   0.3312
  -6.250  -0.1897   0.02371   0.01463  -0.0859   0.9999   0.4766
  -6.000  -0.1828   0.02460   0.01588  -0.0790   0.9999   0.5617
  -5.750  -0.1782   0.02550   0.01687  -0.0719   0.9999   0.6118
  -5.500  -0.1737   0.02627   0.01766  -0.0649   0.9999   0.6513
  -5.250  -0.1669   0.02685   0.01819  -0.0588   0.9999   0.6884
  -5.000  -0.1616   0.02717   0.01849  -0.0525   0.9999   0.7194
  -4.750  -0.1556   0.02728   0.01855  -0.0467   0.9999   0.7496
  -4.500  -0.1480   0.02719   0.01841  -0.0415   0.9999   0.7808
  -4.250  -0.1398   0.02691   0.01807  -0.0367   0.9999   0.8135
  -4.000  -0.1296   0.02643   0.01751  -0.0327   0.9999   0.8478
  -3.750  -0.1210   0.02561   0.01663  -0.0285   0.9999   0.8797
  -3.500  -0.1097   0.02462   0.01558  -0.0254   0.9999   0.9120
  -3.250  -0.0926   0.02362   0.01451  -0.0239   0.9999   0.9450
  -3.000  -0.0655   0.02278   0.01362  -0.0249   0.9999   0.9803
  -2.750  -0.0448   0.02232   0.01313  -0.0259   0.9999   1.0001
  -2.500  -0.0062   0.02255   0.01324  -0.0306   0.9999   1.0001
  -2.250   0.0337   0.02289   0.01347  -0.0352   0.9999   1.0001
  -2.000   0.0735   0.02330   0.01380  -0.0397   0.9999   1.0001
  -1.750   0.1127   0.02378   0.01423  -0.0440   0.9999   1.0001
  -1.500   0.1510   0.02432   0.01473  -0.0479   0.9999   1.0001
  -1.250   0.1883   0.02490   0.01530  -0.0516   0.9999   1.0001
  -1.000   0.2243   0.02554   0.01595  -0.0550   0.9999   1.0001
  -0.750   0.2592   0.02622   0.01667  -0.0582   0.9999   1.0001
  -0.500   0.2929   0.02695   0.01745  -0.0610   0.9999   1.0001
  -0.250   0.3254   0.02774   0.01832  -0.0637   0.9999   1.0001
   0.000   0.3566   0.02859   0.01927  -0.0661   0.9999   1.0001
   0.250   0.3867   0.02951   0.02030  -0.0684   0.9999   1.0001
   0.500   0.4155   0.03050   0.02142  -0.0705   0.9999   1.0001
   0.750   0.4431   0.03158   0.02267  -0.0724   0.9999   1.0001
   1.000   0.4693   0.03277   0.02403  -0.0743   0.9999   1.0001
   1.250   0.4940   0.03408   0.02553  -0.0760   0.9999   1.0001
   1.500   0.5503   0.03555   0.02735  -0.0833   0.9822   1.0001
   1.750   0.6975   0.03264   0.02522  -0.0985   0.8898   1.0001
   2.000   0.7880   0.02458   0.01797  -0.0946   0.7592   1.0001
   2.250   0.8210   0.02670   0.01612  -0.0867   0.2430   1.0001
   2.500   0.8653   0.02906   0.01801  -0.0889   0.1984   1.0001
   2.750   0.9111   0.03123   0.02006  -0.0913   0.1744   1.0001
   3.000   0.9511   0.03356   0.02240  -0.0929   0.1595   1.0001
   3.250   0.9876   0.03623   0.02521  -0.0940   0.1506   1.0001
   3.500   1.0199   0.03876   0.02794  -0.0944   0.1429   1.0001
   3.750   1.0492   0.04186   0.03130  -0.0946   0.1374   1.0001
   4.000   1.0753   0.04509   0.03508  -0.0939   0.1357   1.0001
   4.250   1.0972   0.04848   0.03900  -0.0928   0.1338   1.0001
   4.500   1.1158   0.05203   0.04304  -0.0914   0.1318   1.0001
   4.750   1.1305   0.05619   0.04775  -0.0898   0.1327   1.0001
   5.000   1.1429   0.06092   0.05293  -0.0882   0.1357   1.0001
   5.250   1.1589   0.06610   0.05833  -0.0874   0.1386   1.0001
   5.500   1.1448   0.07206   0.06531  -0.0839   0.1497   1.0001
   5.750   1.1633   0.07817   0.07141  -0.0839   0.1543   1.0001
   6.000   0.8889   0.11860   0.11316  -0.1219   0.4113   1.0001
   6.250   0.8999   0.12386   0.11842  -0.1227   0.4071   1.0001
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