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S6062 8% (s6062-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: S6062 8% (s6062-il)
Reynolds number: 100,000
Max Cl/Cd: 50.44 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-s6062-il-100000.txt
Download as CSV file: xf-s6062-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: S6062 8%                                        
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5270   0.11791   0.11292   0.0005   1.0000   0.0805
  -9.750  -0.5375   0.11541   0.11051  -0.0050   1.0000   0.0826
  -9.500  -0.5490   0.11270   0.10789  -0.0107   1.0000   0.0831
  -9.250  -0.5219   0.10634   0.10148  -0.0054   1.0000   0.0859
  -9.000  -0.5129   0.10281   0.09796  -0.0054   1.0000   0.0894
  -8.750  -0.5116   0.09939   0.09458  -0.0075   1.0000   0.0929
  -8.500  -0.5278   0.09635   0.09167  -0.0144   1.0000   0.0960
  -8.250  -0.4269   0.08244   0.07806  -0.0149   1.0000   0.1081
  -8.000  -0.4543   0.07871   0.07445  -0.0215   1.0000   0.1099
  -7.750  -0.4814   0.07409   0.06985  -0.0289   1.0000   0.1103
  -7.500  -0.4281   0.07000   0.06573  -0.0180   1.0000   0.1179
  -7.250  -0.5207   0.07537   0.07088  -0.0259   1.0000   0.1123
  -7.000  -0.5080   0.07238   0.06792  -0.0241   1.0000   0.1167
  -6.750  -0.5147   0.06721   0.06259  -0.0323   1.0000   0.1251
  -6.500  -0.4985   0.06408   0.05954  -0.0293   1.0000   0.1290
  -6.250  -0.4946   0.05984   0.05515  -0.0327   1.0000   0.1399
  -6.000  -0.4866   0.05663   0.05178  -0.0341   1.0000   0.1528
  -5.750  -0.4757   0.05363   0.04875  -0.0336   1.0000   0.1669
  -5.250  -0.4402   0.03877   0.03275  -0.0364   1.0000   0.0887
  -5.000  -0.4198   0.03191   0.02442  -0.0342   1.0000   0.0578
  -4.750  -0.4006   0.02918   0.02142  -0.0327   1.0000   0.0556
  -4.500  -0.3814   0.02632   0.01822  -0.0312   1.0000   0.0537
  -4.250  -0.3605   0.02397   0.01550  -0.0297   1.0000   0.0527
  -4.000  -0.3387   0.02200   0.01320  -0.0281   1.0000   0.0529
  -3.750  -0.3166   0.02040   0.01138  -0.0267   1.0000   0.0542
  -3.500  -0.2946   0.01912   0.00991  -0.0252   1.0000   0.0567
  -3.250  -0.2743   0.01772   0.00861  -0.0240   1.0000   0.0654
  -3.000  -0.2537   0.01655   0.00748  -0.0226   1.0000   0.0816
  -2.750  -0.2348   0.01378   0.00608  -0.0215   1.0000   0.2900
  -2.500  -0.2252   0.01207   0.00625  -0.0175   1.0000   0.6751
  -2.250  -0.2115   0.01182   0.00646  -0.0123   1.0000   0.8837
  -2.000  -0.1301   0.01178   0.00594  -0.0220   1.0000   1.0000
  -1.750  -0.1216   0.01181   0.00575  -0.0195   1.0000   1.0000
  -1.500  -0.1077   0.01191   0.00563  -0.0179   1.0000   1.0000
  -1.250  -0.0913   0.01206   0.00558  -0.0168   1.0000   1.0000
  -1.000  -0.0738   0.01225   0.00559  -0.0158   1.0000   1.0000
  -0.750  -0.0556   0.01248   0.00564  -0.0151   1.0000   1.0000
  -0.500  -0.0107   0.01288   0.00586  -0.0194   0.9913   1.0000
  -0.250   0.0390   0.01329   0.00612  -0.0244   0.9812   1.0000
   0.000   0.0896   0.01364   0.00634  -0.0295   0.9709   1.0000
   0.250   0.1364   0.01389   0.00652  -0.0337   0.9591   1.0000
   0.500   0.1831   0.01410   0.00668  -0.0378   0.9471   1.0000
   0.750   0.2309   0.01424   0.00680  -0.0419   0.9354   1.0000
   1.000   0.2809   0.01429   0.00687  -0.0463   0.9240   1.0000
   1.250   0.3269   0.01429   0.00691  -0.0496   0.9117   1.0000
   1.500   0.3672   0.01427   0.00694  -0.0517   0.8977   1.0000
   1.750   0.4016   0.01424   0.00695  -0.0524   0.8819   1.0000
   2.000   0.4325   0.01420   0.00695  -0.0522   0.8653   1.0000
   2.250   0.4609   0.01412   0.00696  -0.0514   0.8483   1.0000
   2.500   0.4875   0.01404   0.00690  -0.0501   0.8313   1.0000
   2.750   0.5102   0.01404   0.00696  -0.0483   0.8108   1.0000
   3.000   0.5342   0.01398   0.00692  -0.0465   0.7912   1.0000
   3.250   0.5573   0.01395   0.00694  -0.0446   0.7697   1.0000
   3.500   0.5810   0.01388   0.00694  -0.0427   0.7476   1.0000
   3.750   0.6041   0.01388   0.00698  -0.0409   0.7219   1.0000
   4.000   0.6274   0.01388   0.00701  -0.0391   0.6943   1.0000
   4.250   0.6508   0.01390   0.00705  -0.0373   0.6643   1.0000
   4.500   0.6740   0.01399   0.00719  -0.0356   0.6302   1.0000
   4.750   0.6970   0.01413   0.00731  -0.0339   0.5925   1.0000
   5.000   0.7183   0.01430   0.00741  -0.0319   0.5393   1.0000
   5.250   0.7375   0.01462   0.00748  -0.0298   0.4652   1.0000
   5.500   0.7554   0.01529   0.00776  -0.0279   0.3716   1.0000
   5.750   0.7721   0.01642   0.00838  -0.0263   0.2604   1.0000
   6.000   0.7793   0.01959   0.01022  -0.0238   0.0931   1.0000
   6.250   0.7960   0.02151   0.01200  -0.0220   0.0706   1.0000
   6.500   0.8158   0.02308   0.01360  -0.0205   0.0616   1.0000
   6.750   0.8361   0.02496   0.01539  -0.0194   0.0553   1.0000
   7.000   0.8605   0.02680   0.01738  -0.0184   0.0524   1.0000
   7.250   0.8858   0.02898   0.01976  -0.0175   0.0506   1.0000
   7.500   0.9106   0.03150   0.02254  -0.0167   0.0497   1.0000
   7.750   0.9333   0.03449   0.02593  -0.0155   0.0497   1.0000
   8.000   0.9523   0.03799   0.02998  -0.0141   0.0507   1.0000
   8.250   0.9672   0.04182   0.03435  -0.0125   0.0516   1.0000
   8.500   0.9780   0.04561   0.03864  -0.0109   0.0515   1.0000
   8.750   0.9845   0.04977   0.04328  -0.0093   0.0518   1.0000
   9.000   0.9872   0.05451   0.04843  -0.0079   0.0534   1.0000
   9.250   0.9908   0.05989   0.05399  -0.0070   0.0555   1.0000
   9.500   0.9611   0.06648   0.06141  -0.0051   0.0636   1.0000
   9.750   0.9388   0.07296   0.06820  -0.0047   0.0726   1.0000
  10.000   0.8991   0.07942   0.07482  -0.0068   0.0757   1.0000
  10.250   0.8085   0.07385   0.06947  -0.0033   0.0655   1.0000
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