Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

E222 (10.17%) (e222-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: E222 (10.17%) (e222-il)
Reynolds number: 200,000
Max Cl/Cd: 76.61 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e222-il-200000.txt
Download as CSV file: xf-e222-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E222  (10.17%)                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.4324   0.11507   0.11132  -0.0307   1.0000   0.0521
 -10.500  -0.4438   0.11105   0.10738  -0.0358   1.0000   0.0533
 -10.250  -0.4477   0.10711   0.10348  -0.0385   1.0000   0.0535
 -10.000  -0.4482   0.10146   0.09788  -0.0379   1.0000   0.0544
  -9.750  -0.4361   0.09924   0.09563  -0.0354   1.0000   0.0557
  -9.500  -0.4295   0.09680   0.09321  -0.0346   1.0000   0.0573
  -9.250  -0.4269   0.09402   0.09045  -0.0346   1.0000   0.0594
  -9.000  -0.4302   0.09054   0.08702  -0.0358   1.0000   0.0622
  -8.750  -0.3690   0.07490   0.07172  -0.0400   1.0000   0.0682
  -8.500  -0.3658   0.07232   0.06917  -0.0385   1.0000   0.0696
  -8.250  -0.4911   0.07620   0.07298  -0.0468   1.0000   0.0655
  -8.000  -0.5178   0.07134   0.06801  -0.0499   1.0000   0.0658
  -7.750  -0.5354   0.06812   0.06461  -0.0502   1.0000   0.0660
  -7.500  -0.5302   0.06252   0.05925  -0.0482   1.0000   0.0678
  -7.250  -0.5169   0.06242   0.05926  -0.0441   1.0000   0.0700
  -7.000  -0.5187   0.05967   0.05648  -0.0434   1.0000   0.0720
  -6.750  -0.5235   0.05553   0.05219  -0.0444   1.0000   0.0757
  -6.500  -0.5316   0.04958   0.04581  -0.0466   1.0000   0.0812
  -6.250  -0.5212   0.04724   0.04352  -0.0453   1.0000   0.0828
  -6.000  -0.5057   0.03066   0.02513  -0.0481   1.0000   0.0396
  -5.750  -0.4848   0.02667   0.02058  -0.0480   1.0000   0.0368
  -5.500  -0.4603   0.02366   0.01696  -0.0477   1.0000   0.0350
  -5.250  -0.4281   0.02170   0.01466  -0.0487   0.9982   0.0348
  -5.000  -0.3904   0.02022   0.01297  -0.0509   0.9951   0.0356
  -4.750  -0.3534   0.01925   0.01188  -0.0529   0.9911   0.0378
  -4.500  -0.3159   0.01819   0.01070  -0.0549   0.9871   0.0396
  -4.250  -0.2761   0.01729   0.00969  -0.0574   0.9839   0.0412
  -4.000  -0.2419   0.01617   0.00856  -0.0589   0.9785   0.0443
  -3.750  -0.2027   0.01548   0.00781  -0.0614   0.9740   0.0518
  -3.500  -0.1621   0.01393   0.00694  -0.0648   0.9712   0.1747
  -3.250  -0.1294   0.01335   0.00670  -0.0663   0.9642   0.2680
  -3.000  -0.0897   0.01291   0.00657  -0.0691   0.9597   0.3595
  -2.750  -0.0550   0.01254   0.00649  -0.0707   0.9535   0.4456
  -2.500  -0.0184   0.01221   0.00643  -0.0725   0.9476   0.5277
  -2.250   0.0229   0.01193   0.00631  -0.0750   0.9441   0.5928
  -2.000   0.0528   0.01171   0.00622  -0.0752   0.9353   0.6401
  -1.750   0.0931   0.01146   0.00607  -0.0773   0.9314   0.6884
  -1.500   0.1245   0.01127   0.00597  -0.0775   0.9236   0.7300
  -1.250   0.1626   0.01102   0.00579  -0.0790   0.9188   0.7705
  -1.000   0.1984   0.01078   0.00562  -0.0799   0.9137   0.8066
  -0.750   0.2272   0.01059   0.00547  -0.0793   0.9052   0.8402
  -0.500   0.2579   0.01036   0.00528  -0.0791   0.8983   0.8725
  -0.250   0.2852   0.01014   0.00507  -0.0780   0.8896   0.9038
   0.000   0.3150   0.00993   0.00487  -0.0775   0.8800   0.9356
   0.250   0.3610   0.00967   0.00457  -0.0807   0.8733   0.9621
   0.500   0.4108   0.00949   0.00435  -0.0850   0.8628   0.9855
   0.750   0.4541   0.00933   0.00414  -0.0881   0.8511   1.0000
   1.000   0.4794   0.00927   0.00400  -0.0876   0.8367   1.0000
   1.250   0.5060   0.00925   0.00391  -0.0873   0.8218   1.0000
   1.500   0.5331   0.00926   0.00385  -0.0870   0.8063   1.0000
   1.750   0.5603   0.00929   0.00381  -0.0867   0.7902   1.0000
   2.000   0.5875   0.00935   0.00379  -0.0864   0.7735   1.0000
   2.250   0.6148   0.00943   0.00380  -0.0860   0.7564   1.0000
   2.500   0.6406   0.00953   0.00386  -0.0855   0.7377   1.0000
   2.750   0.6666   0.00965   0.00392  -0.0849   0.7184   1.0000
   3.000   0.6928   0.00979   0.00398  -0.0844   0.6992   1.0000
   3.250   0.7182   0.00994   0.00410  -0.0837   0.6781   1.0000
   3.500   0.7435   0.01011   0.00420  -0.0830   0.6564   1.0000
   3.750   0.7684   0.01030   0.00434  -0.0823   0.6337   1.0000
   4.000   0.7932   0.01052   0.00447  -0.0815   0.6106   1.0000
   4.250   0.8174   0.01075   0.00467  -0.0807   0.5856   1.0000
   4.500   0.8413   0.01100   0.00487  -0.0798   0.5597   1.0000
   4.750   0.8649   0.01129   0.00508  -0.0789   0.5330   1.0000
   5.000   0.8880   0.01160   0.00532  -0.0779   0.5048   1.0000
   5.250   0.9108   0.01194   0.00560  -0.0769   0.4753   1.0000
   5.500   0.9330   0.01231   0.00590  -0.0758   0.4442   1.0000
   5.750   0.9547   0.01273   0.00623  -0.0746   0.4111   1.0000
   6.000   0.9755   0.01321   0.00659  -0.0734   0.3769   1.0000
   6.250   0.9964   0.01371   0.00702  -0.0722   0.3414   1.0000
   6.500   1.0164   0.01429   0.00749  -0.0709   0.3061   1.0000
   6.750   1.0357   0.01494   0.00799  -0.0696   0.2663   1.0000
   7.000   1.0538   0.01569   0.00855  -0.0682   0.2230   1.0000
   7.250   1.0717   0.01651   0.00917  -0.0668   0.1822   1.0000
   7.500   1.0893   0.01738   0.00987  -0.0654   0.1465   1.0000
   7.750   1.1073   0.01824   0.01062  -0.0640   0.1174   1.0000
   8.000   1.1242   0.01921   0.01146  -0.0625   0.0881   1.0000
   8.250   1.1350   0.02080   0.01275  -0.0601   0.0500   1.0000
   8.500   1.1430   0.02258   0.01445  -0.0571   0.0343   1.0000
   8.750   1.1547   0.02388   0.01581  -0.0547   0.0297   1.0000
   9.000   1.1596   0.02560   0.01760  -0.0514   0.0273   1.0000
   9.250   1.1707   0.02688   0.01902  -0.0489   0.0258   1.0000
   9.500   1.1815   0.02833   0.02057  -0.0466   0.0246   1.0000
   9.750   1.1928   0.02987   0.02220  -0.0444   0.0236   1.0000
  10.000   1.2047   0.03147   0.02387  -0.0425   0.0226   1.0000
  10.250   1.2166   0.03354   0.02598  -0.0409   0.0214   1.0000
  10.500   1.2321   0.03631   0.02890  -0.0398   0.0205   1.0000
  10.750   1.2431   0.03798   0.03079  -0.0379   0.0199   1.0000
  11.000   1.2543   0.04029   0.03336  -0.0362   0.0197   1.0000
  11.250   1.2618   0.04286   0.03620  -0.0343   0.0194   1.0000
  11.500   1.2653   0.04567   0.03930  -0.0322   0.0193   1.0000
  11.750   1.2641   0.04871   0.04263  -0.0300   0.0193   1.0000
  12.000   1.2584   0.05202   0.04624  -0.0279   0.0193   1.0000
  12.250   1.2495   0.05562   0.05012  -0.0261   0.0194   1.0000
  12.500   1.2372   0.05959   0.05437  -0.0249   0.0195   1.0000
  12.750   1.2230   0.06394   0.05897  -0.0244   0.0196   1.0000
  13.000   1.2067   0.06874   0.06402  -0.0246   0.0197   1.0000
  13.250   1.1886   0.07410   0.06961  -0.0258   0.0198   1.0000
  13.500   1.1694   0.08001   0.07573  -0.0278   0.0200   1.0000
  13.750   1.1490   0.08660   0.08253  -0.0308   0.0201   1.0000
  14.000   1.1280   0.09383   0.08994  -0.0347   0.0203   1.0000
  14.250   1.1068   0.10177   0.09805  -0.0393   0.0204   1.0000
  14.500   1.0852   0.11063   0.10703  -0.0444   0.0207   1.0000
  14.750   0.9250   0.11587   0.11281  -0.0458   0.0211   1.0000
<< Back to E222 (10.17%) (e222-il)

Polar data table (+)

Polar graphs


<< Back to E222 (10.17%) (e222-il)