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

EPPLER 550 AIRFOIL (e550-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 550 AIRFOIL (e550-il)
Reynolds number: 1,000,000
Max Cl/Cd: 124.41 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e550-il-1000000.txt
Download as CSV file: xf-e550-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 550 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -16.500  -0.6637   0.11225   0.11018  -0.0348   1.0000   0.0050
 -16.250  -0.7141   0.09639   0.09409  -0.0435   1.0000   0.0048
 -16.000  -0.7569   0.08386   0.08129  -0.0505   1.0000   0.0047
 -15.750  -0.7749   0.07688   0.07417  -0.0542   1.0000   0.0047
 -15.500  -0.7959   0.07003   0.06713  -0.0574   1.0000   0.0046
 -15.250  -0.8198   0.06330   0.06018  -0.0601   1.0000   0.0046
 -15.000  -0.8350   0.05824   0.05495  -0.0616   1.0000   0.0046
 -14.750  -0.8475   0.05383   0.05036  -0.0626   1.0000   0.0045
 -14.500  -0.8556   0.05020   0.04657  -0.0631   1.0000   0.0045
 -14.250  -0.8608   0.04705   0.04329  -0.0632   1.0000   0.0045
 -14.000  -0.8646   0.04424   0.04035  -0.0630   1.0000   0.0045
 -13.750  -0.8669   0.04167   0.03767  -0.0626   1.0000   0.0046
 -13.500  -0.8666   0.03944   0.03533  -0.0621   1.0000   0.0046
 -13.250  -0.8651   0.03737   0.03315  -0.0614   1.0000   0.0046
 -13.000  -0.8633   0.03539   0.03106  -0.0605   1.0000   0.0046
 -12.750  -0.8589   0.03369   0.02928  -0.0595   1.0000   0.0046
 -12.500  -0.8530   0.03218   0.02770  -0.0585   1.0000   0.0046
 -12.250  -0.8487   0.03057   0.02600  -0.0573   1.0000   0.0047
 -12.000  -0.8420   0.02921   0.02457  -0.0561   1.0000   0.0047
 -11.750  -0.8360   0.02786   0.02315  -0.0546   1.0000   0.0047
 -11.500  -0.8281   0.02672   0.02196  -0.0532   1.0000   0.0051
 -11.250  -0.8000   0.02520   0.02036  -0.0560   0.9974   0.0051
 -11.000  -0.7875   0.02386   0.01896  -0.0554   0.9640   0.0049
 -10.750  -0.7007   0.02218   0.01714  -0.0695   0.9346   0.0054
 -10.500  -0.6131   0.01998   0.01458  -0.0854   0.8656   0.0062
 -10.250  -0.6055   0.01929   0.01360  -0.0832   0.8098   0.0061
 -10.000  -0.5957   0.01864   0.01280  -0.0814   0.7786   0.0063
  -9.750  -0.5853   0.01796   0.01200  -0.0796   0.7550   0.0064
  -9.500  -0.5731   0.01736   0.01130  -0.0780   0.7351   0.0068
  -9.250  -0.5619   0.01671   0.01055  -0.0763   0.7181   0.0070
  -9.000  -0.5519   0.01605   0.00980  -0.0743   0.7032   0.0071
  -8.750  -0.5424   0.01549   0.00914  -0.0721   0.6894   0.0073
  -8.250  -0.5263   0.01439   0.00788  -0.0667   0.6652   0.0080
  -8.000  -0.5095   0.01398   0.00740  -0.0653   0.6541   0.0085
  -7.500  -0.4705   0.01328   0.00658  -0.0631   0.6338   0.0098
  -7.250  -0.4512   0.01287   0.00611  -0.0620   0.6249   0.0112
  -7.000  -0.4295   0.01252   0.00572  -0.0613   0.6163   0.0129
  -6.750  -0.4078   0.01218   0.00535  -0.0605   0.6084   0.0162
  -6.500  -0.3853   0.01181   0.00499  -0.0599   0.6008   0.0233
  -6.250  -0.3629   0.01147   0.00467  -0.0593   0.5933   0.0341
  -6.000  -0.3394   0.01108   0.00435  -0.0589   0.5868   0.0498
  -5.750  -0.3160   0.01072   0.00406  -0.0585   0.5799   0.0713
  -5.500  -0.2925   0.01030   0.00375  -0.0581   0.5739   0.1033
  -5.250  -0.2693   0.00977   0.00342  -0.0578   0.5678   0.1516
  -5.000  -0.2473   0.00910   0.00302  -0.0575   0.5621   0.2284
  -4.750  -0.2268   0.00800   0.00245  -0.0571   0.5575   0.3649
  -4.500  -0.2033   0.00725   0.00218  -0.0570   0.5527   0.4967
  -4.250  -0.1748   0.00726   0.00217  -0.0572   0.5478   0.5231
  -4.000  -0.1459   0.00732   0.00218  -0.0574   0.5432   0.5359
  -3.750  -0.1168   0.00732   0.00217  -0.0576   0.5392   0.5430
  -3.500  -0.0876   0.00739   0.00217  -0.0579   0.5348   0.5504
  -3.250  -0.0589   0.00745   0.00217  -0.0581   0.5305   0.5567
  -3.000  -0.0299   0.00753   0.00222  -0.0583   0.5267   0.5628
  -2.750  -0.0004   0.00763   0.00227  -0.0586   0.5236   0.5701
  -2.500   0.0285   0.00775   0.00242  -0.0587   0.5201   0.5790
  -2.250   0.0575   0.00786   0.00246  -0.0589   0.5167   0.5832
  -2.000   0.0863   0.00792   0.00244  -0.0592   0.5132   0.5845
  -1.750   0.1155   0.00792   0.00240  -0.0595   0.5103   0.5856
  -1.500   0.1446   0.00786   0.00233  -0.0598   0.5076   0.5872
  -1.250   0.1737   0.00784   0.00230  -0.0601   0.5047   0.5887
  -1.000   0.2027   0.00785   0.00228  -0.0604   0.5018   0.5901
  -0.750   0.2315   0.00789   0.00228  -0.0607   0.4990   0.5915
  -0.500   0.2600   0.00795   0.00231  -0.0609   0.4957   0.5928
  -0.250   0.2893   0.00796   0.00231  -0.0612   0.4938   0.5941
   0.000   0.3185   0.00796   0.00231  -0.0616   0.4915   0.5954
   0.250   0.3477   0.00798   0.00231  -0.0619   0.4891   0.5967
   0.500   0.3767   0.00801   0.00233  -0.0622   0.4867   0.5981
   0.750   0.4055   0.00806   0.00235  -0.0625   0.4843   0.5995
   1.000   0.4341   0.00814   0.00239  -0.0628   0.4815   0.6006
   1.250   0.4628   0.00822   0.00244  -0.0630   0.4789   0.6016
   1.500   0.4918   0.00819   0.00244  -0.0634   0.4771   0.6034
   1.750   0.5207   0.00819   0.00246  -0.0637   0.4751   0.6050
   2.000   0.5495   0.00821   0.00250  -0.0640   0.4728   0.6065
   2.250   0.5782   0.00824   0.00254  -0.0643   0.4704   0.6080
   2.500   0.6068   0.00830   0.00259  -0.0645   0.4680   0.6095
   2.750   0.6350   0.00840   0.00267  -0.0647   0.4650   0.6110
   3.000   0.6635   0.00847   0.00275  -0.0650   0.4625   0.6127
   3.250   0.6923   0.00849   0.00280  -0.0653   0.4602   0.6144
   3.500   0.7210   0.00852   0.00285  -0.0655   0.4575   0.6160
   3.750   0.7495   0.00857   0.00291  -0.0658   0.4547   0.6174
   4.000   0.7777   0.00863   0.00297  -0.0660   0.4518   0.6187
   4.250   0.8053   0.00872   0.00306  -0.0661   0.4484   0.6210
   4.500   0.8336   0.00875   0.00313  -0.0664   0.4458   0.6229
   4.750   0.8620   0.00877   0.00321  -0.0666   0.4430   0.6248
   5.000   0.8903   0.00881   0.00328  -0.0668   0.4397   0.6267
   5.250   0.9180   0.00887   0.00336  -0.0670   0.4362   0.6286
   5.500   0.9450   0.00901   0.00348  -0.0670   0.4319   0.6306
   5.750   0.9734   0.00904   0.00357  -0.0672   0.4289   0.6326
   6.000   1.0015   0.00908   0.00366  -0.0674   0.4251   0.6344
   6.250   1.0289   0.00914   0.00374  -0.0675   0.4208   0.6369
   6.500   1.0552   0.00926   0.00387  -0.0674   0.4160   0.6396
   6.750   1.0833   0.00928   0.00397  -0.0676   0.4119   0.6420
   7.000   1.1105   0.00935   0.00408  -0.0677   0.4061   0.6445
   7.250   1.1365   0.00949   0.00422  -0.0676   0.4002   0.6470
   7.500   1.1640   0.00955   0.00434  -0.0677   0.3934   0.6494
   7.750   1.1893   0.00971   0.00448  -0.0674   0.3850   0.6516
   8.000   1.2157   0.00980   0.00463  -0.0674   0.3756   0.6550
   8.250   1.2404   0.00997   0.00482  -0.0671   0.3635   0.6581
   8.500   1.2637   0.01022   0.00504  -0.0665   0.3474   0.6613
   8.750   1.2852   0.01057   0.00533  -0.0657   0.3281   0.6645
   9.000   1.3038   0.01105   0.00571  -0.0644   0.3031   0.6677
   9.250   1.3194   0.01164   0.00618  -0.0626   0.2757   0.6717
   9.500   1.3337   0.01225   0.00670  -0.0606   0.2511   0.6758
   9.750   1.3476   0.01284   0.00721  -0.0586   0.2294   0.6800
  10.000   1.3576   0.01345   0.00774  -0.0558   0.2093   0.6840
  10.250   1.3640   0.01406   0.00830  -0.0525   0.1918   0.6886
  10.500   1.3698   0.01475   0.00894  -0.0491   0.1748   0.6937
  10.750   1.3743   0.01551   0.00965  -0.0458   0.1584   0.6991
  11.000   1.3779   0.01633   0.01044  -0.0425   0.1443   0.7048
  11.250   1.3796   0.01727   0.01136  -0.0392   0.1316   0.7116
  11.750   1.3818   0.01951   0.01362  -0.0334   0.1106   0.7269
  12.000   1.3820   0.02090   0.01502  -0.0310   0.1015   0.7361
  12.250   1.3795   0.02259   0.01673  -0.0287   0.0920   0.7472
  12.500   1.3796   0.02428   0.01848  -0.0269   0.0852   0.7600
  12.750   1.3764   0.02633   0.02057  -0.0252   0.0764   0.7758
  13.000   1.3748   0.02837   0.02269  -0.0238   0.0707   0.7961
  13.250   1.3726   0.03048   0.02490  -0.0225   0.0648   0.8252
  13.500   1.3670   0.03250   0.02724  -0.0206   0.0592   0.9320
  13.750   1.3680   0.03524   0.03000  -0.0211   0.0529   1.0000
  14.000   1.3666   0.03774   0.03253  -0.0205   0.0479   1.0000
  14.250   1.3637   0.04044   0.03523  -0.0201   0.0441   1.0000
  14.500   1.3624   0.04306   0.03788  -0.0198   0.0399   1.0000
  14.750   1.3596   0.04592   0.04076  -0.0196   0.0366   1.0000
  15.000   1.3599   0.04856   0.04343  -0.0196   0.0335   1.0000
  15.250   1.3574   0.05155   0.04644  -0.0197   0.0305   1.0000
  15.500   1.3598   0.05410   0.04903  -0.0198   0.0287   1.0000
  15.750   1.3566   0.05731   0.05225  -0.0201   0.0255   1.0000
  16.000   1.3583   0.06005   0.05504  -0.0205   0.0239   1.0000
  16.250   1.3585   0.06300   0.05803  -0.0209   0.0219   1.0000
  16.500   1.3586   0.06606   0.06113  -0.0215   0.0204   1.0000
  16.750   1.3595   0.06905   0.06416  -0.0221   0.0188   1.0000
  17.000   1.3582   0.07236   0.06751  -0.0228   0.0173   1.0000
  17.250   1.3592   0.07546   0.07066  -0.0236   0.0159   1.0000
  17.500   1.3595   0.07870   0.07394  -0.0245   0.0149   1.0000
  17.750   1.3582   0.08221   0.07750  -0.0255   0.0139   1.0000
  18.000   1.3598   0.08535   0.08071  -0.0265   0.0129   1.0000
  18.250   1.3586   0.08896   0.08436  -0.0276   0.0121   1.0000
  18.500   1.3569   0.09267   0.08813  -0.0289   0.0112   1.0000
  18.750   1.3570   0.09621   0.09174  -0.0302   0.0106   1.0000
  19.000   1.3570   0.09976   0.09535  -0.0315   0.0099   1.0000
  19.250   1.3531   0.10397   0.09962  -0.0332   0.0091   1.0000
<< Back to EPPLER 550 AIRFOIL (e550-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 550 AIRFOIL (e550-il)