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

EPPLER 557 AIRFOIL (e557-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 557 AIRFOIL (e557-il)
Reynolds number: 1,000,000
Max Cl/Cd: 122.13 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e557-il-1000000.txt
Download as CSV file: xf-e557-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 557 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
  ------ -------- --------- --------- -------- -------- --------
 -17.250  -0.6711   0.12544   0.12317  -0.0423   1.0000   0.0047
 -17.000  -0.7059   0.11394   0.11149  -0.0478   1.0000   0.0047
 -16.750  -0.7267   0.10586   0.10328  -0.0517   1.0000   0.0046
 -16.500  -0.7482   0.09794   0.09523  -0.0556   1.0000   0.0046
 -16.250  -0.7658   0.09118   0.08835  -0.0588   1.0000   0.0045
 -16.000  -0.7802   0.08513   0.08219  -0.0617   1.0000   0.0045
 -15.750  -0.7925   0.07974   0.07669  -0.0641   1.0000   0.0045
 -15.500  -0.8037   0.07472   0.07158  -0.0662   1.0000   0.0045
 -15.250  -0.8158   0.06980   0.06654  -0.0682   1.0000   0.0045
 -15.000  -0.8227   0.06578   0.06244  -0.0697   1.0000   0.0045
 -14.750  -0.8301   0.06188   0.05845  -0.0710   1.0000   0.0046
 -14.500  -0.8369   0.05825   0.05473  -0.0721   1.0000   0.0047
 -14.250  -0.8442   0.05481   0.05121  -0.0728   1.0000   0.0045
 -14.000  -0.8506   0.05165   0.04797  -0.0733   1.0000   0.0047
 -13.750  -0.8565   0.04880   0.04505  -0.0734   1.0000   0.0047
 -13.500  -0.8653   0.04602   0.04219  -0.0731   1.0000   0.0046
 -13.250  -0.8712   0.04369   0.03981  -0.0724   1.0000   0.0046
 -13.000  -0.8816   0.04124   0.03729  -0.0713   1.0000   0.0047
 -12.750  -0.8768   0.03867   0.03464  -0.0731   0.9990   0.0047
 -12.500  -0.8609   0.03596   0.03184  -0.0771   0.9968   0.0049
 -12.250  -0.8428   0.03346   0.02925  -0.0811   0.9946   0.0049
 -12.000  -0.8257   0.03112   0.02682  -0.0845   0.9917   0.0050
 -11.750  -0.8077   0.02895   0.02457  -0.0877   0.9879   0.0052
 -11.500  -0.7864   0.02672   0.02225  -0.0918   0.9850   0.0052
 -11.250  -0.7649   0.02543   0.02089  -0.0937   0.9809   0.0054
 -11.000  -0.7442   0.02242   0.01777  -0.0989   0.9752   0.0056
 -10.750  -0.7104   0.01967   0.01492  -0.1061   0.9730   0.0060
 -10.500  -0.6882   0.01846   0.01365  -0.1075   0.9664   0.0063
 -10.250  -0.6562   0.01737   0.01249  -0.1103   0.9638   0.0065
 -10.000  -0.6225   0.01648   0.01155  -0.1129   0.9621   0.0070
  -9.750  -0.5881   0.01566   0.01067  -0.1155   0.9607   0.0074
  -9.500  -0.5661   0.01466   0.00961  -0.1158   0.9534   0.0079
  -9.250  -0.5332   0.01393   0.00885  -0.1177   0.9506   0.0088
  -9.000  -0.4975   0.01329   0.00816  -0.1201   0.9485   0.0095
  -8.750  -0.4660   0.01260   0.00745  -0.1217   0.9434   0.0111
  -8.500  -0.4317   0.01201   0.00683  -0.1237   0.9384   0.0133
  -8.250  -0.3913   0.01145   0.00626  -0.1270   0.9353   0.0164
  -8.000  -0.3496   0.01093   0.00574  -0.1305   0.9317   0.0206
  -7.750  -0.3118   0.01050   0.00531  -0.1331   0.9246   0.0253
  -7.500  -0.2676   0.01008   0.00488  -0.1371   0.9194   0.0320
  -7.250  -0.2310   0.00966   0.00448  -0.1395   0.9101   0.0419
  -7.000  -0.1934   0.00924   0.00409  -0.1421   0.9009   0.0572
  -6.750  -0.1614   0.00887   0.00375  -0.1434   0.8890   0.0767
  -6.500  -0.1313   0.00848   0.00343  -0.1443   0.8772   0.1019
  -6.250  -0.1017   0.00800   0.00306  -0.1452   0.8655   0.1436
  -6.000  -0.0729   0.00722   0.00256  -0.1464   0.8537   0.2260
  -5.750  -0.0449   0.00664   0.00224  -0.1471   0.8420   0.3118
  -5.500  -0.0165   0.00652   0.00214  -0.1473   0.8310   0.3414
  -5.250   0.0118   0.00649   0.00207  -0.1474   0.8201   0.3577
  -5.000   0.0398   0.00647   0.00201  -0.1474   0.8089   0.3700
  -4.750   0.0679   0.00647   0.00196  -0.1473   0.7985   0.3804
  -4.500   0.0957   0.00651   0.00193  -0.1473   0.7879   0.3885
  -4.250   0.1237   0.00652   0.00188  -0.1472   0.7773   0.3950
  -4.000   0.1517   0.00654   0.00185  -0.1472   0.7672   0.4008
  -3.500   0.2074   0.00661   0.00180  -0.1470   0.7465   0.4122
  -3.250   0.2352   0.00670   0.00184  -0.1469   0.7365   0.4208
  -3.000   0.2630   0.00675   0.00183  -0.1469   0.7264   0.4261
  -2.750   0.2910   0.00674   0.00180  -0.1468   0.7168   0.4294
  -2.250   0.3467   0.00681   0.00177  -0.1467   0.6975   0.4347
  -2.000   0.3744   0.00686   0.00175  -0.1466   0.6883   0.4374
  -1.750   0.4023   0.00691   0.00175  -0.1466   0.6785   0.4398
  -1.500   0.4302   0.00692   0.00173  -0.1466   0.6695   0.4426
  -1.250   0.4578   0.00695   0.00172  -0.1465   0.6600   0.4456
  -1.000   0.4858   0.00698   0.00174  -0.1465   0.6509   0.4483
  -0.500   0.5413   0.00709   0.00177  -0.1463   0.6326   0.4540
  -0.250   0.5687   0.00718   0.00179  -0.1462   0.6233   0.4564
   0.000   0.5965   0.00720   0.00181  -0.1462   0.6140   0.4595
   0.250   0.6241   0.00725   0.00184  -0.1461   0.6051   0.4626
   0.500   0.6515   0.00731   0.00188  -0.1460   0.5954   0.4656
   0.750   0.6792   0.00738   0.00193  -0.1459   0.5865   0.4685
   1.000   0.7061   0.00748   0.00198  -0.1457   0.5768   0.4714
   1.250   0.7339   0.00755   0.00203  -0.1456   0.5677   0.4740
   1.500   0.7608   0.00762   0.00208  -0.1455   0.5580   0.4779
   1.750   0.7882   0.00769   0.00215  -0.1454   0.5481   0.4812
   2.000   0.8152   0.00778   0.00222  -0.1452   0.5384   0.4844
   2.500   0.8689   0.00799   0.00238  -0.1448   0.5178   0.4903
   2.750   0.8955   0.00809   0.00247  -0.1445   0.5077   0.4945
   3.000   0.9220   0.00820   0.00257  -0.1443   0.4971   0.4984
   3.250   0.9488   0.00830   0.00267  -0.1441   0.4872   0.5023
   3.500   0.9747   0.00845   0.00278  -0.1437   0.4770   0.5058
   3.750   1.0013   0.00856   0.00289  -0.1435   0.4666   0.5096
   4.000   1.0275   0.00868   0.00301  -0.1432   0.4564   0.5139
   4.250   1.0530   0.00884   0.00315  -0.1427   0.4455   0.5182
   4.500   1.0791   0.00898   0.00329  -0.1424   0.4347   0.5226
   4.750   1.1047   0.00913   0.00343  -0.1420   0.4236   0.5274
   5.000   1.1296   0.00931   0.00359  -0.1415   0.4118   0.5325
   5.250   1.1547   0.00948   0.00376  -0.1410   0.4002   0.5377
   5.500   1.1798   0.00966   0.00393  -0.1405   0.3884   0.5426
   5.750   1.2042   0.00986   0.00412  -0.1399   0.3759   0.5483
   6.000   1.2281   0.01008   0.00433  -0.1392   0.3630   0.5544
   6.250   1.2514   0.01033   0.00454  -0.1384   0.3494   0.5607
   6.500   1.2747   0.01057   0.00478  -0.1376   0.3351   0.5679
   6.750   1.2972   0.01086   0.00502  -0.1367   0.3196   0.5749
   7.000   1.3190   0.01115   0.00530  -0.1357   0.3032   0.5827
   7.250   1.3395   0.01148   0.00558  -0.1344   0.2866   0.5907
   7.500   1.3588   0.01182   0.00589  -0.1328   0.2684   0.5995
   7.750   1.3767   0.01222   0.00623  -0.1311   0.2488   0.6089
   8.000   1.3936   0.01268   0.00663  -0.1292   0.2287   0.6196
   8.250   1.4098   0.01316   0.00705  -0.1272   0.2107   0.6304
   8.500   1.4262   0.01364   0.00749  -0.1253   0.1941   0.6423
   8.750   1.4422   0.01414   0.00796  -0.1233   0.1775   0.6553
   9.000   1.4573   0.01468   0.00846  -0.1213   0.1613   0.6691
   9.250   1.4718   0.01523   0.00899  -0.1191   0.1470   0.6848
   9.500   1.4864   0.01579   0.00954  -0.1170   0.1343   0.7025
   9.750   1.4998   0.01639   0.01013  -0.1148   0.1224   0.7232
  10.000   1.5135   0.01697   0.01075  -0.1127   0.1118   0.7474
  10.250   1.5267   0.01757   0.01140  -0.1105   0.1023   0.7771
  10.500   1.5387   0.01820   0.01210  -0.1082   0.0937   0.8194
  10.750   1.5425   0.01861   0.01275  -0.1043   0.0865   1.0000
  11.000   1.5544   0.01938   0.01351  -0.1022   0.0792   1.0000
  11.250   1.5641   0.02029   0.01440  -0.0999   0.0718   1.0000
  11.500   1.5743   0.02121   0.01531  -0.0978   0.0656   1.0000
  11.750   1.5839   0.02219   0.01629  -0.0957   0.0597   1.0000
  12.000   1.5913   0.02335   0.01742  -0.0934   0.0539   1.0000
  12.250   1.6007   0.02443   0.01852  -0.0916   0.0494   1.0000
  12.500   1.6065   0.02578   0.01986  -0.0894   0.0445   1.0000
  12.750   1.6152   0.02700   0.02112  -0.0877   0.0411   1.0000
  13.000   1.6201   0.02855   0.02267  -0.0858   0.0371   1.0000
  13.250   1.6274   0.02997   0.02413  -0.0842   0.0342   1.0000
  13.500   1.6310   0.03175   0.02591  -0.0825   0.0309   1.0000
  13.750   1.6375   0.03337   0.02758  -0.0812   0.0286   1.0000
  14.000   1.6397   0.03543   0.02965  -0.0797   0.0259   1.0000
  14.250   1.6455   0.03723   0.03151  -0.0787   0.0241   1.0000
  14.500   1.6481   0.03941   0.03373  -0.0776   0.0224   1.0000
  14.750   1.6511   0.04164   0.03602  -0.0766   0.0208   1.0000
  15.000   1.6545   0.04392   0.03835  -0.0759   0.0194   1.0000
  15.250   1.6550   0.04656   0.04104  -0.0752   0.0181   1.0000
  15.500   1.6570   0.04916   0.04371  -0.0747   0.0170   1.0000
  15.750   1.6591   0.05183   0.04645  -0.0744   0.0161   1.0000
  16.000   1.6589   0.05484   0.04952  -0.0742   0.0151   1.0000
  16.250   1.6567   0.05821   0.05295  -0.0742   0.0142   1.0000
  16.500   1.6582   0.06119   0.05602  -0.0743   0.0135   1.0000
  16.750   1.6575   0.06455   0.05946  -0.0746   0.0129   1.0000
  17.000   1.6548   0.06827   0.06325  -0.0751   0.0121   1.0000
  17.250   1.6498   0.07241   0.06747  -0.0758   0.0115   1.0000
  17.500   1.6486   0.07612   0.07127  -0.0765   0.0111   1.0000
  17.750   1.6463   0.08007   0.07531  -0.0774   0.0106   1.0000
  18.000   1.6422   0.08436   0.07969  -0.0786   0.0101   1.0000
  18.250   1.6364   0.08900   0.08442  -0.0799   0.0096   1.0000
  18.500   1.6285   0.09407   0.08957  -0.0816   0.0092   1.0000
<< Back to EPPLER 557 AIRFOIL (e557-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 557 AIRFOIL (e557-il)