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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 500,000
Max Cl/Cd: 72.36 at α=0.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20606-il-500000.txt
Download as CSV file: xf-sc20606-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.6102   0.08437   0.08207  -0.0131   1.0000   0.0180
  -9.000  -0.6139   0.07975   0.07748  -0.0158   1.0000   0.0183
  -8.750  -0.6185   0.07440   0.07217  -0.0203   1.0000   0.0185
  -8.500  -0.6188   0.06658   0.06432  -0.0306   1.0000   0.0186
  -8.250  -0.6156   0.06022   0.05785  -0.0360   1.0000   0.0191
  -8.000  -0.6054   0.05408   0.05151  -0.0400   1.0000   0.0204
  -7.750  -0.5849   0.05059   0.04773  -0.0417   1.0000   0.0216
  -6.500  -0.4734   0.02297   0.01801  -0.0541   1.0000   0.0228
  -6.250  -0.4392   0.01737   0.01185  -0.0559   1.0000   0.0175
  -6.000  -0.4092   0.01517   0.00936  -0.0564   1.0000   0.0176
  -5.750  -0.3805   0.01388   0.00790  -0.0568   1.0000   0.0184
  -5.500  -0.3515   0.01284   0.00675  -0.0571   1.0000   0.0194
  -5.250  -0.3227   0.01208   0.00591  -0.0575   1.0000   0.0206
  -5.000  -0.2944   0.01159   0.00536  -0.0577   1.0000   0.0220
  -4.750  -0.2597   0.01027   0.00395  -0.0597   1.0000   0.0258
  -4.500  -0.2302   0.00987   0.00349  -0.0602   1.0000   0.0309
  -4.250  -0.1994   0.00937   0.00298  -0.0610   1.0000   0.0438
  -4.000  -0.1688   0.00893   0.00263  -0.0619   1.0000   0.0699
  -3.750  -0.1341   0.00816   0.00234  -0.0641   1.0000   0.1857
  -3.500  -0.0905   0.00677   0.00209  -0.0690   1.0000   0.5004
  -3.250  -0.0597   0.00654   0.00210  -0.0698   1.0000   0.5877
  -3.000  -0.0307   0.00645   0.00213  -0.0701   1.0000   0.6341
  -2.750  -0.0024   0.00644   0.00216  -0.0702   1.0000   0.6609
  -2.500   0.0256   0.00645   0.00219  -0.0702   1.0000   0.6833
  -2.250   0.0532   0.00646   0.00227  -0.0701   1.0000   0.7072
  -2.000   0.0805   0.00651   0.00237  -0.0699   1.0000   0.7288
  -1.750   0.1077   0.00656   0.00247  -0.0697   1.0000   0.7443
  -1.500   0.1348   0.00662   0.00258  -0.0694   1.0000   0.7583
  -1.250   0.1617   0.00669   0.00269  -0.0692   1.0000   0.7713
  -1.000   0.1890   0.00677   0.00281  -0.0691   1.0000   0.7824
  -0.750   0.2158   0.00682   0.00294  -0.0688   1.0000   0.7911
  -0.500   0.2429   0.00690   0.00308  -0.0687   1.0000   0.7995
  -0.250   0.2792   0.00684   0.00309  -0.0706   0.9965   0.8072
   0.000   0.3362   0.00643   0.00276  -0.0768   0.9829   0.8127
   0.250   0.3764   0.00620   0.00264  -0.0792   0.9687   0.8187
   0.500   0.4262   0.00589   0.00234  -0.0834   0.9146   0.8245
   0.750   0.4441   0.00718   0.00237  -0.0803   0.5967   0.8300
   1.000   0.4627   0.00904   0.00277  -0.0792   0.2404   0.8372
   1.250   0.4863   0.01000   0.00311  -0.0786   0.0850   0.8431
   1.500   0.5129   0.01050   0.00347  -0.0783   0.0507   0.8503
   1.750   0.5390   0.01091   0.00388  -0.0778   0.0344   0.8565
   2.000   0.5651   0.01153   0.00454  -0.0773   0.0265   0.8638
   2.250   0.5909   0.01198   0.00507  -0.0767   0.0234   0.8701
   2.500   0.6168   0.01259   0.00574  -0.0761   0.0211   0.8773
   2.750   0.6401   0.01388   0.00715  -0.0749   0.0195   0.8837
   3.000   0.6644   0.01556   0.00898  -0.0739   0.0187   0.8909
   3.250   0.6899   0.01599   0.00953  -0.0732   0.0178   0.8978
   3.500   0.7156   0.01729   0.01100  -0.0724   0.0173   0.9056
   3.750   0.7395   0.01890   0.01288  -0.0711   0.0168   0.9129
   4.000   0.7636   0.02138   0.01573  -0.0697   0.0167   0.9208
   4.250   0.7844   0.02528   0.02001  -0.0676   0.0197   0.9293
  12.500   0.7011   0.16656   0.16484  -0.0808   0.0124   1.0000
  12.750   0.7018   0.16971   0.16799  -0.0821   0.0123   1.0000
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