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NASA SC(2)-0606 AIRFOIL (sc20606-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0606 AIRFOIL (sc20606-il)
Reynolds number: 500,000
Max Cl/Cd: 53.61 at α=-0.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20606-il-500000-n5.txt
Download as CSV file: xf-sc20606-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6488   0.08096   0.07867  -0.0142   1.0000   0.0070
  -9.500  -0.6594   0.07431   0.07207  -0.0186   1.0000   0.0069
  -9.250  -0.6765   0.06145   0.05920  -0.0334   1.0000   0.0067
  -9.000  -0.6822   0.05224   0.04979  -0.0400   1.0000   0.0067
  -8.750  -0.6786   0.04439   0.04164  -0.0440   1.0000   0.0068
  -8.500  -0.6677   0.03648   0.03330  -0.0473   1.0000   0.0070
  -8.250  -0.6486   0.02938   0.02562  -0.0503   1.0000   0.0073
  -8.000  -0.6245   0.02666   0.02259  -0.0515   1.0000   0.0077
  -7.750  -0.5984   0.02384   0.01938  -0.0528   1.0000   0.0083
  -7.500  -0.5698   0.02032   0.01535  -0.0544   1.0000   0.0085
  -7.250  -0.5412   0.01791   0.01256  -0.0554   1.0000   0.0087
  -7.000  -0.5130   0.01622   0.01060  -0.0560   1.0000   0.0091
  -6.750  -0.4849   0.01499   0.00916  -0.0565   1.0000   0.0096
  -6.500  -0.4571   0.01406   0.00808  -0.0568   1.0000   0.0100
  -6.250  -0.4293   0.01333   0.00721  -0.0571   1.0000   0.0103
  -6.000  -0.4000   0.01235   0.00610  -0.0579   1.0000   0.0107
  -5.750  -0.3699   0.01145   0.00510  -0.0588   1.0000   0.0116
  -5.500  -0.3416   0.01100   0.00462  -0.0592   1.0000   0.0128
  -5.250  -0.3129   0.01057   0.00414  -0.0596   1.0000   0.0138
  -5.000  -0.2840   0.01018   0.00366  -0.0601   1.0000   0.0149
  -4.750  -0.2554   0.00986   0.00329  -0.0604   1.0000   0.0166
  -4.500  -0.2262   0.00952   0.00295  -0.0609   1.0000   0.0224
  -4.250  -0.1980   0.00929   0.00273  -0.0611   1.0000   0.0307
  -4.000  -0.1699   0.00910   0.00257  -0.0614   1.0000   0.0408
  -3.750  -0.1411   0.00887   0.00241  -0.0618   1.0000   0.0618
  -3.500  -0.1104   0.00848   0.00225  -0.0629   1.0000   0.1163
  -3.250  -0.0785   0.00805   0.00212  -0.0642   1.0000   0.1987
  -3.000  -0.0456   0.00761   0.00200  -0.0658   1.0000   0.2987
  -2.750  -0.0107   0.00708   0.00191  -0.0680   1.0000   0.4289
  -2.500   0.0227   0.00672   0.00193  -0.0695   1.0000   0.5455
  -2.250   0.0527   0.00659   0.00200  -0.0701   1.0000   0.6055
  -2.000   0.0814   0.00656   0.00208  -0.0704   1.0000   0.6395
  -1.750   0.1097   0.00656   0.00215  -0.0705   1.0000   0.6640
  -1.500   0.1392   0.00656   0.00221  -0.0709   0.9993   0.6795
  -1.250   0.1825   0.00638   0.00209  -0.0743   0.9929   0.6957
  -1.000   0.2244   0.00620   0.00198  -0.0773   0.9848   0.7128
  -0.750   0.2640   0.00602   0.00187  -0.0797   0.9723   0.7253
  -0.500   0.3068   0.00581   0.00172  -0.0828   0.9480   0.7342
  -0.250   0.3415   0.00637   0.00151  -0.0833   0.7318   0.7418
   0.000   0.3615   0.00758   0.00176  -0.0817   0.4884   0.7493
   0.250   0.3866   0.00830   0.00196  -0.0814   0.3434   0.7564
   0.500   0.4125   0.00890   0.00215  -0.0813   0.2283   0.7630
   0.750   0.4388   0.00946   0.00236  -0.0812   0.1323   0.7696
   1.000   0.4652   0.00996   0.00258  -0.0810   0.0645   0.7761
   1.250   0.4926   0.01025   0.00281  -0.0810   0.0421   0.7826
   1.500   0.5198   0.01048   0.00304  -0.0808   0.0312   0.7881
   1.750   0.5472   0.01075   0.00330  -0.0807   0.0226   0.7944
   2.000   0.5741   0.01112   0.00368  -0.0804   0.0163   0.8007
   2.250   0.6013   0.01145   0.00407  -0.0802   0.0148   0.8080
   2.500   0.6280   0.01184   0.00458  -0.0798   0.0136   0.8140
   2.750   0.6545   0.01230   0.00512  -0.0794   0.0128   0.8200
   3.000   0.6810   0.01273   0.00559  -0.0791   0.0117   0.8257
   3.250   0.7067   0.01338   0.00632  -0.0786   0.0107   0.8322
   3.500   0.7312   0.01456   0.00766  -0.0778   0.0101   0.8384
   3.750   0.7568   0.01524   0.00848  -0.0773   0.0098   0.8447
   4.000   0.7825   0.01603   0.00941  -0.0768   0.0095   0.8513
   4.250   0.8076   0.01701   0.01057  -0.0761   0.0092   0.8578
   4.500   0.8328   0.01829   0.01206  -0.0754   0.0087   0.8641
   4.750   0.8568   0.01993   0.01399  -0.0744   0.0084   0.8696
   5.000   0.8800   0.02214   0.01656  -0.0733   0.0082   0.8758
   5.250   0.9013   0.02487   0.01974  -0.0718   0.0080   0.8818
   5.500   0.9227   0.02696   0.02219  -0.0707   0.0075   0.8888
   5.750   0.9436   0.02870   0.02420  -0.0696   0.0071   0.8958
   6.000   0.9601   0.03223   0.02817  -0.0678   0.0068   0.9034
   6.250   0.9726   0.03682   0.03323  -0.0656   0.0066   0.9116
   6.500   0.9812   0.04237   0.03923  -0.0634   0.0064   0.9212
   6.750   0.9819   0.04978   0.04710  -0.0609   0.0062   0.9344
   7.000   0.9740   0.05880   0.05654  -0.0588   0.0061   0.9673
   7.250   0.9654   0.06845   0.06646  -0.0592   0.0060   1.0000
   7.500   0.9604   0.07598   0.07416  -0.0608   0.0060   1.0000
   7.750   0.9517   0.08325   0.08152  -0.0637   0.0060   1.0000
   8.000   0.9409   0.08936   0.08770  -0.0667   0.0060   1.0000
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