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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 50,000
Max Cl/Cd: 25.35 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20606-il-50000-n5.txt
Download as CSV file: xf-sc20606-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0606 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6150   0.10777   0.10043  -0.0073   1.0000   0.0620
  -9.750  -0.6104   0.10355   0.09624  -0.0079   1.0000   0.0603
  -9.250  -0.6243   0.09048   0.08332  -0.0217   1.0000   0.0500
  -9.000  -0.6210   0.08575   0.07861  -0.0239   1.0000   0.0496
  -8.750  -0.6180   0.08096   0.07383  -0.0265   1.0000   0.0491
  -8.500  -0.6147   0.07614   0.06898  -0.0290   1.0000   0.0484
  -8.250  -0.6106   0.07117   0.06394  -0.0319   1.0000   0.0476
  -8.000  -0.6048   0.06602   0.05866  -0.0350   1.0000   0.0467
  -7.750  -0.5960   0.06078   0.05321  -0.0381   1.0000   0.0458
  -7.500  -0.5834   0.05555   0.04765  -0.0411   1.0000   0.0451
  -7.250  -0.5664   0.05044   0.04212  -0.0439   1.0000   0.0445
  -7.000  -0.5454   0.04566   0.03683  -0.0462   1.0000   0.0443
  -6.750  -0.5210   0.04133   0.03194  -0.0481   1.0000   0.0444
  -6.500  -0.4945   0.03750   0.02744  -0.0495   1.0000   0.0449
  -6.250  -0.4668   0.03432   0.02365  -0.0504   1.0000   0.0471
  -6.000  -0.4387   0.03166   0.02040  -0.0509   1.0000   0.0510
  -5.750  -0.4133   0.02940   0.01796  -0.0509   1.0000   0.0547
  -5.500  -0.3875   0.02737   0.01566  -0.0502   1.0000   0.0577
  -5.250  -0.3626   0.02567   0.01365  -0.0490   1.0000   0.0620
  -5.000  -0.3385   0.02425   0.01221  -0.0483   1.0000   0.0727
  -4.750  -0.3138   0.02283   0.01064  -0.0475   1.0000   0.0838
  -4.500  -0.2882   0.02155   0.00928  -0.0472   1.0000   0.1047
  -4.250  -0.2588   0.01994   0.00788  -0.0483   1.0000   0.1449
  -4.000  -0.2299   0.01708   0.00715  -0.0502   1.0000   0.4861
  -3.750  -0.2172   0.01702   0.00751  -0.0451   1.0000   0.6595
  -3.500  -0.2052   0.01721   0.00765  -0.0400   1.0000   0.7388
  -3.250  -0.1999   0.01734   0.00782  -0.0330   1.0000   0.7993
  -3.000  -0.1941   0.01721   0.00764  -0.0266   1.0000   0.8461
  -2.750  -0.1826   0.01683   0.00716  -0.0222   1.0000   0.8819
  -2.500  -0.1657   0.01634   0.00653  -0.0193   1.0000   0.9125
  -2.250  -0.1415   0.01586   0.00590  -0.0185   1.0000   0.9389
  -2.000  -0.1098   0.01547   0.00531  -0.0196   1.0000   0.9623
  -1.750  -0.0765   0.01516   0.00488  -0.0213   1.0000   0.9887
  -1.500  -0.0506   0.01498   0.00462  -0.0218   1.0000   1.0000
  -1.250  -0.0212   0.01495   0.00448  -0.0231   1.0000   1.0000
  -1.000   0.0089   0.01496   0.00440  -0.0245   1.0000   1.0000
  -0.750   0.0395   0.01500   0.00439  -0.0258   1.0000   1.0000
  -0.500   0.0700   0.01507   0.00443  -0.0272   1.0000   1.0000
  -0.250   0.1006   0.01516   0.00453  -0.0284   1.0000   1.0000
   0.000   0.1309   0.01528   0.00470  -0.0296   1.0000   1.0000
   0.250   0.1610   0.01542   0.00491  -0.0307   1.0000   1.0000
   0.500   0.1909   0.01559   0.00517  -0.0317   1.0000   1.0000
   0.750   0.2204   0.01577   0.00549  -0.0327   1.0000   1.0000
   1.000   0.2496   0.01598   0.00591  -0.0335   1.0000   1.0000
   1.250   0.2785   0.01622   0.00635  -0.0343   1.0000   1.0000
   1.500   0.3071   0.01649   0.00686  -0.0350   1.0000   1.0000
   1.750   0.3353   0.01679   0.00745  -0.0357   1.0000   1.0000
   2.000   0.3632   0.01713   0.00814  -0.0363   1.0000   1.0000
   2.250   0.4758   0.01877   0.00751  -0.0471   0.2288   1.0000
   2.500   0.5000   0.02097   0.00892  -0.0470   0.1215   1.0000
   2.750   0.5285   0.02251   0.01030  -0.0474   0.0894   1.0000
   3.000   0.5587   0.02415   0.01202  -0.0476   0.0759   1.0000
   3.250   0.5885   0.02574   0.01378  -0.0478   0.0628   1.0000
   3.500   0.6200   0.02760   0.01597  -0.0477   0.0572   1.0000
   3.750   0.6504   0.02981   0.01845  -0.0476   0.0537   1.0000
   4.000   0.6781   0.03227   0.02107  -0.0477   0.0500   1.0000
   4.250   0.7061   0.03476   0.02415  -0.0472   0.0461   1.0000
   4.500   0.7324   0.03770   0.02767  -0.0467   0.0439   1.0000
   4.750   0.7566   0.04113   0.03175  -0.0460   0.0433   1.0000
   5.000   0.7784   0.04490   0.03609  -0.0453   0.0431   1.0000
   5.250   0.7978   0.04901   0.04075  -0.0446   0.0432   1.0000
   5.500   0.8145   0.05347   0.04572  -0.0440   0.0436   1.0000
   5.750   0.8286   0.05819   0.05090  -0.0436   0.0443   1.0000
   6.000   0.8402   0.06309   0.05619  -0.0435   0.0451   1.0000
   6.250   0.8493   0.06811   0.06152  -0.0437   0.0459   1.0000
   6.500   0.8558   0.07319   0.06686  -0.0442   0.0463   1.0000
   6.750   0.8597   0.07836   0.07225  -0.0451   0.0465   1.0000
   7.000   0.8611   0.08359   0.07764  -0.0465   0.0467   1.0000
   7.250   0.8605   0.08886   0.08303  -0.0482   0.0470   1.0000
   7.500   0.8586   0.09411   0.08835  -0.0503   0.0473   1.0000
   7.750   0.8551   0.09915   0.09342  -0.0523   0.0478   1.0000
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