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

EPPLER 555 AIRFOIL (e555-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 555 AIRFOIL (e555-il)
Reynolds number: 50,000
Max Cl/Cd: 32.81 at α=8.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e555-il-50000-n5.txt
Download as CSV file: xf-e555-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 555 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.750  -0.5167   0.10082   0.09322  -0.0560   1.0000   0.0407
 -12.500  -0.5420   0.09246   0.08479  -0.0604   1.0000   0.0401
 -12.250  -0.5662   0.08528   0.07749  -0.0641   1.0000   0.0395
 -12.000  -0.5888   0.07903   0.07111  -0.0670   1.0000   0.0393
 -11.750  -0.6071   0.07385   0.06578  -0.0689   1.0000   0.0392
 -11.500  -0.6210   0.06952   0.06128  -0.0699   1.0000   0.0392
 -11.250  -0.6314   0.06574   0.05734  -0.0704   1.0000   0.0394
 -11.000  -0.6389   0.06238   0.05382  -0.0702   1.0000   0.0400
 -10.750  -0.6442   0.05940   0.05063  -0.0696   1.0000   0.0409
 -10.500  -0.6473   0.05668   0.04769  -0.0686   1.0000   0.0421
 -10.250  -0.6490   0.05420   0.04492  -0.0672   1.0000   0.0434
 -10.000  -0.6382   0.05266   0.04348  -0.0656   1.0000   0.0457
  -9.750  -0.6291   0.05109   0.04183  -0.0636   1.0000   0.0483
  -9.500  -0.6138   0.04971   0.04019  -0.0612   1.0000   0.0515
  -9.250  -0.6044   0.04856   0.03914  -0.0588   1.0000   0.0547
  -9.000  -0.6000   0.04735   0.03792  -0.0563   1.0000   0.0587
  -8.750  -0.5957   0.04624   0.03675  -0.0533   1.0000   0.0631
  -8.500  -0.6006   0.04502   0.03565  -0.0502   1.0000   0.0670
  -8.250  -0.6060   0.04392   0.03452  -0.0466   1.0000   0.0713
  -8.000  -0.6138   0.04263   0.03331  -0.0433   1.0000   0.0756
  -7.750  -0.6212   0.04117   0.03191  -0.0405   1.0000   0.0811
  -7.500  -0.6272   0.03964   0.03042  -0.0380   1.0000   0.0877
  -7.250  -0.6322   0.03801   0.02884  -0.0358   1.0000   0.0962
  -7.000  -0.6359   0.03617   0.02712  -0.0341   1.0000   0.1078
  -6.750  -0.6362   0.03410   0.02518  -0.0331   1.0000   0.1265
  -6.500  -0.6337   0.03138   0.02284  -0.0336   1.0000   0.1619
  -6.250  -0.6139   0.02992   0.02293  -0.0350   0.9944   0.3082
  -6.000  -0.5750   0.03108   0.02384  -0.0373   0.9858   0.4030
  -5.750  -0.5398   0.03286   0.02535  -0.0374   0.9770   0.4409
  -5.500  -0.5062   0.03436   0.02660  -0.0372   0.9685   0.4669
  -5.250  -0.4729   0.03590   0.02796  -0.0362   0.9605   0.4863
  -5.000  -0.4438   0.03675   0.02862  -0.0353   0.9515   0.5023
  -4.500  -0.3796   0.03649   0.02783  -0.0381   0.9347   0.5290
  -4.250  -0.3494   0.03661   0.02780  -0.0379   0.9266   0.5344
  -4.000  -0.3150   0.03577   0.02666  -0.0410   0.9183   0.5440
  -3.750  -0.2861   0.03549   0.02622  -0.0415   0.9094   0.5488
  -3.500  -0.2510   0.03507   0.02562  -0.0434   0.9020   0.5544
  -3.250  -0.2200   0.03416   0.02444  -0.0463   0.8924   0.5619
  -3.000  -0.1840   0.03389   0.02404  -0.0478   0.8857   0.5656
  -2.750  -0.1569   0.03352   0.02354  -0.0484   0.8757   0.5701
  -2.500  -0.1158   0.03280   0.02259  -0.0522   0.8695   0.5764
  -2.250  -0.0893   0.03247   0.02215  -0.0527   0.8591   0.5806
  -2.000  -0.0502   0.03208   0.02165  -0.0549   0.8532   0.5845
  -1.750  -0.0240   0.03176   0.02122  -0.0555   0.8425   0.5892
  -1.500   0.0197   0.03115   0.02044  -0.0594   0.8369   0.5951
  -1.250   0.0428   0.03104   0.02029  -0.0588   0.8257   0.5985
  -1.000   0.0846   0.03064   0.01980  -0.0615   0.8203   0.6028
  -0.750   0.1100   0.03043   0.01951  -0.0620   0.8089   0.6079
  -0.500   0.1447   0.03012   0.01913  -0.0637   0.8009   0.6127
  -0.250   0.1746   0.02995   0.01894  -0.0642   0.7918   0.6166
   0.000   0.2043   0.02978   0.01872  -0.0649   0.7823   0.6213
   0.250   0.2419   0.02943   0.01828  -0.0673   0.7744   0.6273
   0.500   0.2660   0.02941   0.01830  -0.0667   0.7638   0.6310
   0.750   0.3031   0.02912   0.01798  -0.0682   0.7569   0.6360
   1.000   0.3289   0.02911   0.01795  -0.0686   0.7455   0.6415
   1.250   0.3652   0.02885   0.01768  -0.0700   0.7382   0.6463
   1.500   0.3894   0.02889   0.01777  -0.0695   0.7271   0.6511
   1.750   0.4185   0.02887   0.01774  -0.0702   0.7172   0.6572
   2.000   0.4528   0.02868   0.01757  -0.0712   0.7090   0.6625
   2.250   0.4756   0.02883   0.01778  -0.0706   0.6977   0.6677
   2.500   0.5132   0.02860   0.01754  -0.0723   0.6901   0.6746
   2.750   0.5356   0.02877   0.01779  -0.0715   0.6787   0.6799
   3.000   0.5614   0.02888   0.01798  -0.0713   0.6683   0.6862
   3.250   0.5967   0.02873   0.01785  -0.0725   0.6599   0.6933
   3.500   0.6168   0.02901   0.01824  -0.0714   0.6483   0.6994
   3.750   0.6472   0.02904   0.01833  -0.0720   0.6387   0.7073
   4.000   0.6742   0.02908   0.01847  -0.0716   0.6288   0.7142
   4.250   0.6972   0.02936   0.01885  -0.0711   0.6173   0.7224
   4.500   0.7269   0.02931   0.01891  -0.0711   0.6081   0.7303
   4.750   0.7523   0.02950   0.01919  -0.0709   0.5969   0.7398
   5.000   0.7722   0.02978   0.01963  -0.0696   0.5854   0.7484
   5.250   0.8010   0.02979   0.01975  -0.0695   0.5753   0.7589
   5.500   0.8256   0.02990   0.01998  -0.0689   0.5641   0.7699
   5.750   0.8436   0.03022   0.02047  -0.0673   0.5519   0.7813
   6.000   0.8661   0.03036   0.02077  -0.0663   0.5405   0.7946
   6.250   0.8954   0.03017   0.02069  -0.0659   0.5300   0.8096
   6.500   0.9094   0.03055   0.02126  -0.0637   0.5170   0.8267
   6.750   0.9249   0.03081   0.02173  -0.0616   0.5045   0.8476
   7.000   0.9429   0.03086   0.02196  -0.0597   0.4924   0.8764
   7.250   0.9677   0.03076   0.02203  -0.0589   0.4797   0.9336
   7.500   0.9974   0.03097   0.02231  -0.0597   0.4652   1.0000
   7.750   1.0237   0.03137   0.02276  -0.0600   0.4503   1.0000
   8.000   1.0452   0.03191   0.02336  -0.0595   0.4349   1.0000
   8.250   1.0651   0.03246   0.02397  -0.0587   0.4193   1.0000
   8.500   1.0832   0.03305   0.02462  -0.0576   0.4035   1.0000
   8.750   1.0987   0.03367   0.02528  -0.0561   0.3875   1.0000
   9.000   1.1121   0.03435   0.02601  -0.0543   0.3715   1.0000
   9.250   1.1243   0.03511   0.02679  -0.0525   0.3552   1.0000
   9.500   1.1330   0.03607   0.02779  -0.0503   0.3387   1.0000
   9.750   1.1403   0.03716   0.02894  -0.0482   0.3222   1.0000
  10.000   1.1465   0.03837   0.03020  -0.0462   0.3058   1.0000
  10.250   1.1514   0.03974   0.03161  -0.0442   0.2895   1.0000
  10.500   1.1555   0.04124   0.03314  -0.0424   0.2736   1.0000
  10.750   1.1588   0.04289   0.03482  -0.0407   0.2580   1.0000
  11.000   1.1619   0.04468   0.03663  -0.0392   0.2431   1.0000
  11.250   1.1641   0.04663   0.03860  -0.0379   0.2288   1.0000
  11.500   1.1664   0.04869   0.04064  -0.0368   0.2153   1.0000
  11.750   1.1674   0.05097   0.04294  -0.0358   0.2023   1.0000
  12.000   1.1668   0.05359   0.04564  -0.0351   0.1900   1.0000
  12.250   1.1665   0.05630   0.04841  -0.0345   0.1786   1.0000
  12.500   1.1668   0.05899   0.05112  -0.0341   0.1680   1.0000
  12.750   1.1682   0.06162   0.05373  -0.0338   0.1582   1.0000
  13.000   1.1649   0.06505   0.05732  -0.0338   0.1489   1.0000
  13.250   1.1650   0.06804   0.06034  -0.0338   0.1404   1.0000
  13.500   1.1631   0.07144   0.06386  -0.0341   0.1324   1.0000
  13.750   1.1610   0.07498   0.06749  -0.0345   0.1252   1.0000
  14.000   1.1590   0.07854   0.07114  -0.0351   0.1183   1.0000
  14.250   1.1565   0.08237   0.07507  -0.0359   0.1123   1.0000
  14.500   1.1504   0.08683   0.07971  -0.0371   0.1066   1.0000
  14.750   1.1527   0.08993   0.08277  -0.0377   0.1011   1.0000
  15.000   1.1396   0.09601   0.08917  -0.0400   0.0971   1.0000
  15.250   1.1342   0.10068   0.09396  -0.0418   0.0927   1.0000
  15.500   1.1350   0.10427   0.09756  -0.0429   0.0883   1.0000
  15.750   1.1133   0.11257   0.10618  -0.0471   0.0862   1.0000
  16.000   1.0866   0.12238   0.11627  -0.0525   0.0846   1.0000
  16.250   1.0416   0.13736   0.13146  -0.0614   0.0842   1.0000
<< Back to EPPLER 555 AIRFOIL (e555-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 555 AIRFOIL (e555-il)