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E226 (10.19%) (e226-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: E226 (10.19%) (e226-il)
Reynolds number: 50,000
Max Cl/Cd: 30.45 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e226-il-50000.txt
Download as CSV file: xf-e226-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E226  (10.19%)                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5216   0.09994   0.09310  -0.0007   1.0000   0.2868
  -8.500  -0.6131   0.07524   0.06859  -0.0301   1.0000   0.1246
  -8.250  -0.6498   0.06689   0.05994  -0.0346   1.0000   0.1125
  -8.000  -0.6467   0.06233   0.05530  -0.0345   1.0000   0.1108
  -7.750  -0.6481   0.05756   0.05033  -0.0344   1.0000   0.1091
  -7.500  -0.6480   0.05287   0.04531  -0.0340   1.0000   0.1070
  -7.250  -0.6453   0.04836   0.04037  -0.0330   1.0000   0.1056
  -7.000  -0.6378   0.04424   0.03576  -0.0317   1.0000   0.1046
  -6.750  -0.6262   0.04055   0.03158  -0.0301   1.0000   0.1045
  -6.500  -0.6109   0.03737   0.02798  -0.0285   1.0000   0.1059
  -6.250  -0.5931   0.03453   0.02470  -0.0269   1.0000   0.1088
  -6.000  -0.5742   0.03195   0.02175  -0.0253   1.0000   0.1156
  -5.750  -0.5540   0.02991   0.01941  -0.0236   1.0000   0.1273
  -5.500  -0.5328   0.02763   0.01731  -0.0221   1.0000   0.1455
  -5.250  -0.5151   0.02561   0.01550  -0.0201   1.0000   0.1864
  -5.000  -0.5040   0.02411   0.01454  -0.0173   1.0000   0.2609
  -4.750  -0.4917   0.02283   0.01373  -0.0142   1.0000   0.3425
  -4.500  -0.4779   0.02191   0.01319  -0.0109   1.0000   0.4159
  -4.250  -0.4634   0.02123   0.01279  -0.0075   1.0000   0.4836
  -4.000  -0.4485   0.02070   0.01250  -0.0040   1.0000   0.5476
  -3.750  -0.4340   0.02030   0.01228  -0.0001   1.0000   0.6078
  -3.500  -0.4203   0.02002   0.01214   0.0042   1.0000   0.6648
  -3.250  -0.4073   0.01983   0.01201   0.0088   1.0000   0.7203
  -3.000  -0.3941   0.01974   0.01196   0.0136   1.0000   0.7751
  -2.750  -0.3742   0.01979   0.01200   0.0176   1.0000   0.8318
  -2.500  -0.3193   0.02024   0.01216   0.0155   1.0000   0.8979
  -2.250  -0.1452   0.02062   0.01164  -0.0093   1.0000   0.9610
  -2.000  -0.0291   0.01952   0.01003  -0.0278   1.0000   1.0000
  -1.750  -0.0398   0.01906   0.00958  -0.0241   1.0000   1.0000
  -1.500  -0.0561   0.01870   0.00923  -0.0193   1.0000   1.0000
  -1.250  -0.0736   0.01843   0.00894  -0.0140   1.0000   1.0000
  -1.000  -0.0851   0.01828   0.00869  -0.0093   1.0000   1.0000
  -0.750  -0.0877   0.01827   0.00855  -0.0058   1.0000   1.0000
  -0.500  -0.0838   0.01837   0.00852  -0.0032   1.0000   1.0000
  -0.250  -0.0753   0.01857   0.00859  -0.0014   1.0000   1.0000
   0.000  -0.0641   0.01885   0.00875   0.0001   1.0000   1.0000
   0.250  -0.0513   0.01919   0.00899   0.0012   1.0000   1.0000
   0.500  -0.0377   0.01960   0.00930   0.0021   1.0000   1.0000
   0.750  -0.0234   0.02007   0.00969   0.0028   1.0000   1.0000
   1.000  -0.0088   0.02060   0.01015   0.0034   1.0000   1.0000
   1.250   0.0142   0.02131   0.01083   0.0022   0.9966   1.0000
   1.500   0.0749   0.02259   0.01211  -0.0058   0.9764   1.0000
   1.750   0.1350   0.02373   0.01329  -0.0133   0.9560   1.0000
   2.000   0.1870   0.02460   0.01424  -0.0190   0.9336   1.0000
   2.250   0.2464   0.02540   0.01515  -0.0256   0.9123   1.0000
   2.500   0.2926   0.02598   0.01587  -0.0294   0.8886   1.0000
   2.750   0.3496   0.02639   0.01645  -0.0346   0.8657   1.0000
   3.000   0.4046   0.02655   0.01683  -0.0390   0.8417   1.0000
   3.250   0.4497   0.02664   0.01712  -0.0411   0.8169   1.0000
   3.500   0.5032   0.02625   0.01697  -0.0436   0.7934   1.0000
   3.750   0.5366   0.02617   0.01707  -0.0429   0.7668   1.0000
   4.000   0.5681   0.02600   0.01706  -0.0417   0.7395   1.0000
   4.250   0.5985   0.02574   0.01694  -0.0399   0.7116   1.0000
   4.500   0.6268   0.02546   0.01681  -0.0376   0.6824   1.0000
   4.750   0.6534   0.02519   0.01662  -0.0350   0.6513   1.0000
   5.000   0.6794   0.02491   0.01639  -0.0322   0.6185   1.0000
   5.250   0.7023   0.02485   0.01635  -0.0293   0.5823   1.0000
   5.500   0.7257   0.02479   0.01627  -0.0264   0.5436   1.0000
   5.750   0.7464   0.02500   0.01643  -0.0235   0.5012   1.0000
   6.000   0.7668   0.02533   0.01663  -0.0207   0.4564   1.0000
   6.250   0.7867   0.02584   0.01690  -0.0180   0.4094   1.0000
   6.500   0.8035   0.02671   0.01769  -0.0153   0.3605   1.0000
   6.750   0.8204   0.02778   0.01856  -0.0129   0.3128   1.0000
   7.000   0.8371   0.02913   0.01968  -0.0105   0.2673   1.0000
   7.250   0.8530   0.03082   0.02117  -0.0083   0.2244   1.0000
   7.500   0.8663   0.03259   0.02275  -0.0059   0.1836   1.0000
   7.750   0.8799   0.03456   0.02442  -0.0039   0.1483   1.0000
   8.000   0.8961   0.03706   0.02703  -0.0021   0.1255   1.0000
   8.250   0.9141   0.03991   0.03004  -0.0007   0.1120   1.0000
   8.500   0.9302   0.04351   0.03400   0.0008   0.1053   1.0000
   8.750   0.9425   0.04680   0.03770   0.0025   0.1000   1.0000
   9.000   0.9575   0.05030   0.04122   0.0034   0.0949   1.0000
   9.250   0.9576   0.05422   0.04576   0.0056   0.0934   1.0000
   9.500   0.9554   0.05846   0.05046   0.0074   0.0929   1.0000
   9.750   0.9489   0.06297   0.05533   0.0089   0.0932   1.0000
  10.000   0.9391   0.06750   0.06014   0.0102   0.0937   1.0000
  10.250   0.9271   0.07211   0.06495   0.0111   0.0944   1.0000
  10.500   0.9143   0.07674   0.06970   0.0119   0.0951   1.0000
  10.750   0.8214   0.08693   0.08018   0.0048   0.1037   1.0000
  11.000   0.8037   0.09458   0.08781   0.0005   0.1067   1.0000
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