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E210 (13.64%) (e210-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: E210 (13.64%) (e210-il)
Reynolds number: 50,000
Max Cl/Cd: 9.81 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e210-il-50000.txt
Download as CSV file: xf-e210-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E210  (13.64%)                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3052   0.11527   0.10870  -0.0306   1.0000   0.2582
  -8.750  -0.3015   0.11200   0.10551  -0.0296   1.0000   0.2642
  -8.500  -0.3099   0.11068   0.10428  -0.0284   1.0000   0.2742
  -8.250  -0.3027   0.10717   0.10083  -0.0270   1.0000   0.2805
  -8.000  -0.3227   0.10684   0.10066  -0.0253   1.0000   0.2905
  -7.750  -0.3059   0.10282   0.09667  -0.0234   1.0000   0.3015
  -7.500  -0.3138   0.10095   0.09492  -0.0214   1.0000   0.3108
  -7.250  -0.3359   0.10079   0.09492  -0.0184   1.0000   0.3213
  -7.000  -0.3216   0.09718   0.09134  -0.0164   1.0000   0.3318
  -6.750  -0.3716   0.09906   0.09350  -0.0117   1.0000   0.3376
  -6.500  -0.3454   0.09453   0.08896  -0.0103   1.0000   0.3490
  -6.250  -0.3651   0.09355   0.08814  -0.0066   1.0000   0.3561
  -6.000  -0.3735   0.09214   0.08683  -0.0034   1.0000   0.3636
  -5.750  -0.4007   0.09178   0.08664   0.0011   1.0000   0.3711
  -5.500  -0.4854   0.06079   0.05470  -0.0483   1.0000   0.1385
  -5.250  -0.4733   0.05716   0.05096  -0.0488   1.0000   0.1349
  -5.000  -0.4577   0.05279   0.04625  -0.0513   1.0000   0.1326
  -4.750  -0.4375   0.04857   0.04153  -0.0541   1.0000   0.1314
  -4.500  -0.4140   0.04478   0.03715  -0.0565   1.0000   0.1299
  -4.250  -0.3902   0.04187   0.03369  -0.0580   1.0000   0.1298
  -4.000  -0.3672   0.03974   0.03108  -0.0588   1.0000   0.1324
  -3.750  -0.3313   0.03785   0.02863  -0.0616   0.9959   0.1399
  -3.500  -0.2833   0.03638   0.02676  -0.0657   0.9861   0.1520
  -3.250  -0.2412   0.03513   0.02555  -0.0689   0.9762   0.1732
  -3.000  -0.2001   0.03392   0.02441  -0.0716   0.9665   0.2157
  -2.750  -0.1584   0.03258   0.02393  -0.0746   0.9578   0.3306
  -2.500  -0.1315   0.03283   0.02514  -0.0734   0.9471   0.4968
  -2.250  -0.1131   0.03356   0.02611  -0.0700   0.9362   0.5942
  -2.000  -0.0957   0.03415   0.02675  -0.0666   0.9259   0.6624
  -1.750  -0.0761   0.03461   0.02720  -0.0634   0.9167   0.7192
  -1.500  -0.0645   0.03478   0.02733  -0.0595   0.9063   0.7646
  -1.250  -0.0534   0.03488   0.02736  -0.0558   0.8969   0.8079
  -1.000  -0.0343   0.03490   0.02732  -0.0530   0.8887   0.8557
  -0.750  -0.0241   0.03478   0.02718  -0.0498   0.8786   0.9023
  -0.500   0.0545   0.03520   0.02735  -0.0598   0.8685   0.9818
  -0.250   0.0989   0.03548   0.02731  -0.0661   0.8594   1.0000
   0.000   0.1237   0.03601   0.02761  -0.0691   0.8490   1.0000
   0.250   0.1654   0.03673   0.02802  -0.0742   0.8403   1.0000
   0.500   0.2033   0.03751   0.02853  -0.0783   0.8311   1.0000
   0.750   0.2314   0.03846   0.02925  -0.0806   0.8221   1.0000
   1.000   0.2745   0.03927   0.02980  -0.0845   0.8141   1.0000
   1.250   0.2905   0.04043   0.03080  -0.0848   0.8049   1.0000
   1.500   0.3348   0.04125   0.03140  -0.0883   0.7977   1.0000
   1.750   0.3423   0.04263   0.03267  -0.0873   0.7886   1.0000
   2.000   0.3785   0.04359   0.03348  -0.0895   0.7812   1.0000
   2.250   0.3881   0.04506   0.03486  -0.0887   0.7730   1.0000
   2.500   0.4223   0.04611   0.03579  -0.0906   0.7659   1.0000
   2.750   0.4292   0.04773   0.03735  -0.0895   0.7580   1.0000
   3.000   0.4654   0.04879   0.03832  -0.0915   0.7511   1.0000
   3.250   0.4657   0.05069   0.04019  -0.0899   0.7437   1.0000
   3.500   0.5070   0.05167   0.04110  -0.0922   0.7366   1.0000
   3.750   0.5000   0.05390   0.04331  -0.0901   0.7301   1.0000
   4.000   0.5283   0.05525   0.04463  -0.0912   0.7230   1.0000
   4.250   0.5345   0.05728   0.04665  -0.0904   0.7169   1.0000
   4.500   0.5468   0.05916   0.04852  -0.0902   0.7107   1.0000
   4.750   0.5713   0.06079   0.05015  -0.0909   0.7037   1.0000
   5.000   0.5706   0.06318   0.05255  -0.0898   0.6991   1.0000
   5.250   0.6115   0.06440   0.05378  -0.0917   0.6902   1.0000
   5.500   0.6015   0.06719   0.05661  -0.0901   0.6868   1.0000
   5.750   0.6043   0.06967   0.05912  -0.0896   0.6831   1.0000
   6.000   0.6357   0.07131   0.06078  -0.0907   0.6736   1.0000
   6.250   0.6326   0.07417   0.06368  -0.0900   0.6716   1.0000
   6.500   0.6346   0.07701   0.06658  -0.0898   0.6706   1.0000
   6.750   0.6408   0.08007   0.06969  -0.0903   0.6722   1.0000
   7.000   0.6557   0.08334   0.07305  -0.0916   0.6752   1.0000
   7.250   0.6953   0.08358   0.07334  -0.0913   0.6439   1.0000
   8.750   0.7692   0.09836   0.08863  -0.0911   0.5822   1.0000
   9.000   0.7798   0.10060   0.09096  -0.0907   0.5677   1.0000
   9.250   0.7794   0.10471   0.09519  -0.0914   0.5664   1.0000
   9.500   0.7852   0.10696   0.09752  -0.0907   0.5495   1.0000
   9.750   0.8118   0.10701   0.09768  -0.0890   0.5167   1.0000
  10.000   0.8244   0.10922   0.10000  -0.0886   0.5003   1.0000
  10.250   0.8738   0.10828   0.09927  -0.0872   0.4743   1.0000
  10.500   0.8539   0.11325   0.10429  -0.0877   0.4659   1.0000
  10.750   0.8956   0.11274   0.10396  -0.0861   0.4435   1.0000
  11.000   0.9072   0.11498   0.10633  -0.0857   0.4284   1.0000
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