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EPPLER 1213 AIRFOIL (e1213-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 1213 AIRFOIL (e1213-il)
Reynolds number: 50,000
Max Cl/Cd: 11.52 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1213-il-50000.txt
Download as CSV file: xf-e1213-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1213 AIRFOIL                             
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.3946   0.13353   0.12575   0.0074   1.0000   0.3349
 -11.250  -0.3569   0.12735   0.11950   0.0075   1.0000   0.3457
 -11.000  -0.3821   0.12771   0.11994   0.0067   1.0000   0.3543
 -10.750  -0.3412   0.12142   0.11358   0.0067   1.0000   0.3649
 -10.500  -0.3628   0.12125   0.11348   0.0061   1.0000   0.3743
 -10.250  -0.3252   0.11574   0.10792   0.0060   1.0000   0.3854
 -10.000  -0.3386   0.11457   0.10681   0.0057   1.0000   0.3950
  -9.750  -0.3098   0.11038   0.10259   0.0056   1.0000   0.4073
  -9.500  -0.3115   0.10811   0.10036   0.0054   1.0000   0.4166
  -9.250  -0.3007   0.10585   0.09812   0.0055   1.0000   0.4307
  -9.000  -0.2840   0.10214   0.09442   0.0051   1.0000   0.4396
  -8.750  -0.3079   0.10279   0.09518   0.0062   1.0000   0.4538
  -8.500  -0.2603   0.09689   0.08922   0.0051   1.0000   0.4646
  -8.250  -0.2615   0.09502   0.08742   0.0054   1.0000   0.4773
  -8.000  -0.2670   0.09447   0.08694   0.0065   1.0000   0.4935
  -7.750  -0.2284   0.08948   0.08196   0.0052   1.0000   0.5035
  -7.500  -0.2265   0.08768   0.08023   0.0058   1.0000   0.5183
  -7.250  -0.2382   0.08759   0.08027   0.0076   1.0000   0.5352
  -7.000  -0.1929   0.08269   0.07538   0.0056   1.0000   0.5474
  -6.750  -0.1819   0.08050   0.07330   0.0058   1.0000   0.5625
  -6.500  -0.1788   0.07903   0.07197   0.0066   1.0000   0.5791
  -6.250  -0.1894   0.07903   0.07218   0.0090   1.0000   0.5965
  -6.000  -0.4514   0.07198   0.06551   0.0067   0.9894   0.4137
  -5.750  -0.4018   0.06678   0.06015  -0.0034   0.9696   0.4105
  -5.500  -0.3724   0.06106   0.05425  -0.0136   0.9494   0.4090
  -5.250  -0.3530   0.05572   0.04869  -0.0220   0.9290   0.4092
  -5.000  -0.3397   0.05043   0.04305  -0.0297   0.9101   0.4124
  -4.750  -0.2794   0.04903   0.04166  -0.0340   0.8937   0.4207
  -4.500  -0.2544   0.04644   0.03885  -0.0373   0.8757   0.4270
  -4.250  -0.2433   0.04357   0.03563  -0.0397   0.8583   0.4326
  -4.000  -0.2066   0.04249   0.03452  -0.0407   0.8417   0.4392
  -3.750  -0.1856   0.04141   0.03331  -0.0406   0.8246   0.4460
  -3.500  -0.1751   0.03995   0.03155  -0.0405   0.8079   0.4526
  -3.250  -0.1523   0.03933   0.03092  -0.0397   0.7917   0.4584
  -3.000  -0.1309   0.03875   0.03025  -0.0390   0.7767   0.4656
  -2.750  -0.1086   0.03763   0.02880  -0.0391   0.7644   0.4743
  -2.500  -0.0889   0.03745   0.02869  -0.0376   0.7485   0.4806
  -2.250  -0.0713   0.03717   0.02831  -0.0367   0.7342   0.4884
  -2.000  -0.0478   0.03661   0.02758  -0.0361   0.7224   0.4972
  -1.750  -0.0269   0.03649   0.02746  -0.0350   0.7086   0.5059
  -1.500  -0.0102   0.03640   0.02722  -0.0344   0.6954   0.5151
  -1.250   0.0196   0.03599   0.02680  -0.0333   0.6855   0.5252
  -1.000   0.0333   0.03629   0.02701  -0.0326   0.6712   0.5357
  -0.750   0.0558   0.03637   0.02715  -0.0314   0.6600   0.5467
  -0.500   0.0783   0.03636   0.02709  -0.0307   0.6486   0.5592
  -0.250   0.0953   0.03685   0.02759  -0.0299   0.6367   0.5722
   0.000   0.1233   0.03665   0.02740  -0.0289   0.6270   0.5887
   0.250   0.1350   0.03759   0.02841  -0.0279   0.6148   0.6033
   0.500   0.1654   0.03726   0.02808  -0.0268   0.6064   0.6252
   0.750   0.1724   0.03859   0.02953  -0.0258   0.5940   0.6438
   1.000   0.2036   0.03820   0.02917  -0.0245   0.5864   0.6763
   1.250   0.2041   0.04000   0.03120  -0.0231   0.5745   0.7010
   1.500   0.2351   0.03946   0.03087  -0.0209   0.5670   0.7505
   1.750   0.2310   0.04163   0.03337  -0.0193   0.5560   0.7985
   2.000   0.3435   0.04155   0.03357  -0.0292   0.5422   0.9280
   2.250   0.4725   0.04100   0.03267  -0.0460   0.5270   1.0000
   2.500   0.4213   0.04498   0.03661  -0.0421   0.5195   1.0000
   2.750   0.4200   0.04677   0.03823  -0.0401   0.5119   1.0000
   3.000   0.4467   0.04707   0.03831  -0.0387   0.5054   1.0000
   3.250   0.3867   0.05331   0.04456  -0.0365   0.4986   1.0000
   3.500   0.3818   0.05596   0.04710  -0.0352   0.4930   1.0000
   3.750   0.4164   0.05643   0.04743  -0.0345   0.4857   1.0000
   4.000   0.3688   0.06161   0.05255  -0.0326   0.4836   1.0000
   4.250   0.3531   0.06534   0.05620  -0.0321   0.4831   1.0000
   4.500   0.3468   0.06875   0.05955  -0.0321   0.4837   1.0000
   4.750   0.3459   0.07200   0.06273  -0.0323   0.4846   1.0000
   5.000   0.3568   0.07522   0.06591  -0.0330   0.4884   1.0000
   5.250   0.2417   0.08428   0.07515  -0.0379   0.5989   1.0000
   5.500   0.2661   0.08688   0.07767  -0.0385   0.5881   1.0000
   5.750   0.2670   0.08866   0.07938  -0.0376   0.5791   1.0000
   6.000   0.2908   0.09126   0.08191  -0.0380   0.5685   1.0000
   6.250   0.2976   0.09371   0.08430  -0.0379   0.5623   1.0000
   6.500   0.3058   0.09535   0.08589  -0.0374   0.5511   1.0000
   6.750   0.3496   0.10062   0.09111  -0.0394   0.5459   1.0000
   7.000   0.3253   0.09981   0.09025  -0.0371   0.5336   1.0000
   7.250   0.3634   0.10427   0.09467  -0.0384   0.5270   1.0000
   7.500   0.3458   0.10454   0.09489  -0.0370   0.5174   1.0000
   7.750   0.3680   0.10754   0.09787  -0.0375   0.5101   1.0000
   8.000   0.3906   0.11168   0.10197  -0.0383   0.5059   1.0000
   8.250   0.3799   0.11180   0.10207  -0.0373   0.4943   1.0000
   8.500   0.4159   0.11649   0.10673  -0.0384   0.4884   1.0000
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