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

EPPLER 855 AIRFOIL (e855-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 855 AIRFOIL (e855-il)
Reynolds number: 500,000
Max Cl/Cd: 111.99 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e855-il-500000.txt
Download as CSV file: xf-e855-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 855 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.5062   0.08438   0.08198  -0.0589   1.0000   0.0112
 -12.250  -0.5346   0.07531   0.07282  -0.0639   1.0000   0.0110
 -12.000  -0.5615   0.06777   0.06519  -0.0680   1.0000   0.0108
 -11.750  -0.5879   0.06131   0.05861  -0.0713   1.0000   0.0108
 -11.500  -0.6117   0.05613   0.05333  -0.0732   1.0000   0.0106
 -11.250  -0.6415   0.05112   0.04819  -0.0741   1.0000   0.0104
 -11.000  -0.6578   0.04568   0.04254  -0.0787   0.9979   0.0103
 -10.750  -0.6585   0.03971   0.03622  -0.0874   0.9917   0.0103
 -10.500  -0.6530   0.03464   0.03073  -0.0938   0.9844   0.0101
 -10.250  -0.6395   0.03102   0.02676  -0.0962   0.9769   0.0101
 -10.000  -0.6154   0.02785   0.02323  -0.0992   0.9731   0.0099
  -9.750  -0.5968   0.02569   0.02085  -0.0996   0.9656   0.0100
  -9.500  -0.5684   0.02375   0.01872  -0.1015   0.9618   0.0101
  -9.250  -0.5465   0.02221   0.01703  -0.1016   0.9543   0.0103
  -9.000  -0.5173   0.02074   0.01541  -0.1030   0.9496   0.0107
  -8.750  -0.4930   0.01950   0.01406  -0.1032   0.9421   0.0109
  -8.500  -0.4637   0.01840   0.01284  -0.1042   0.9365   0.0114
  -8.250  -0.4397   0.01686   0.01121  -0.1046   0.9281   0.0120
  -8.000  -0.4067   0.01583   0.01013  -0.1064   0.9220   0.0129
  -7.750  -0.3722   0.01497   0.00919  -0.1084   0.9143   0.0140
  -7.500  -0.3319   0.01388   0.00801  -0.1117   0.9074   0.0160
  -7.250  -0.2880   0.01308   0.00713  -0.1157   0.8981   0.0200
  -7.000  -0.2417   0.01232   0.00636  -0.1202   0.8871   0.0277
  -6.750  -0.2017   0.01178   0.00575  -0.1232   0.8712   0.0375
  -6.500  -0.1682   0.01136   0.00526  -0.1247   0.8532   0.0478
  -6.250  -0.1389   0.01104   0.00487  -0.1253   0.8350   0.0601
  -6.000  -0.1126   0.01068   0.00451  -0.1253   0.8173   0.0790
  -5.750  -0.0874   0.01037   0.00419  -0.1251   0.8005   0.1027
  -5.500  -0.0631   0.01001   0.00387  -0.1247   0.7849   0.1348
  -5.250  -0.0393   0.00957   0.00355  -0.1243   0.7702   0.1843
  -5.000  -0.0165   0.00899   0.00321  -0.1239   0.7564   0.2660
  -4.750   0.0076   0.00863   0.00304  -0.1235   0.7433   0.3427
  -4.500   0.0334   0.00855   0.00297  -0.1232   0.7308   0.3841
  -4.250   0.0598   0.00855   0.00293  -0.1228   0.7193   0.4092
  -4.000   0.0865   0.00859   0.00288  -0.1226   0.7082   0.4260
  -3.750   0.1132   0.00862   0.00285  -0.1223   0.6974   0.4396
  -3.500   0.1400   0.00868   0.00283  -0.1220   0.6874   0.4526
  -3.250   0.1669   0.00875   0.00280  -0.1218   0.6776   0.4639
  -3.000   0.1939   0.00879   0.00279  -0.1216   0.6684   0.4731
  -2.750   0.2210   0.00885   0.00276  -0.1214   0.6601   0.4808
  -2.500   0.2481   0.00889   0.00275  -0.1212   0.6514   0.4886
  -2.250   0.2752   0.00895   0.00274  -0.1211   0.6436   0.4970
  -2.000   0.3023   0.00903   0.00277  -0.1209   0.6356   0.5073
  -1.750   0.3294   0.00908   0.00278  -0.1207   0.6288   0.5153
  -1.500   0.3567   0.00913   0.00278  -0.1206   0.6216   0.5223
  -1.000   0.4113   0.00919   0.00277  -0.1204   0.6085   0.5324
  -0.750   0.4386   0.00925   0.00277  -0.1204   0.6023   0.5377
  -0.500   0.4661   0.00929   0.00278  -0.1203   0.5967   0.5427
  -0.250   0.4934   0.00931   0.00280  -0.1202   0.5909   0.5476
   0.000   0.5209   0.00940   0.00282  -0.1202   0.5855   0.5532
   0.250   0.5483   0.00942   0.00285  -0.1202   0.5802   0.5584
   0.500   0.5754   0.00945   0.00289  -0.1201   0.5748   0.5636
   0.750   0.6031   0.00956   0.00294  -0.1201   0.5700   0.5696
   1.000   0.6304   0.00959   0.00300  -0.1201   0.5654   0.5751
   1.250   0.6575   0.00963   0.00306  -0.1200   0.5605   0.5811
   1.500   0.6850   0.00973   0.00311  -0.1200   0.5560   0.5876
   1.750   0.7122   0.00979   0.00321  -0.1199   0.5515   0.5940
   2.000   0.7394   0.00984   0.00329  -0.1198   0.5471   0.6013
   2.250   0.7664   0.00990   0.00337  -0.1198   0.5429   0.6082
   2.500   0.7943   0.01004   0.00348  -0.1199   0.5388   0.6165
   2.750   0.8208   0.01006   0.00359  -0.1197   0.5347   0.6246
   3.000   0.8477   0.01013   0.00369  -0.1196   0.5304   0.6340
   3.250   0.8746   0.01021   0.00382  -0.1195   0.5266   0.6434
   3.500   0.9022   0.01035   0.00396  -0.1196   0.5227   0.6542
   3.750   0.9283   0.01039   0.00410  -0.1193   0.5187   0.6663
   4.000   0.9546   0.01045   0.00424  -0.1191   0.5147   0.6795
   4.250   0.9812   0.01054   0.00438  -0.1189   0.5108   0.6940
   4.500   1.0080   0.01066   0.00456  -0.1188   0.5069   0.7110
   4.750   1.0328   0.01069   0.00472  -0.1183   0.5023   0.7303
   5.000   1.0573   0.01072   0.00485  -0.1177   0.4972   0.7531
   5.250   1.0823   0.01083   0.00502  -0.1172   0.4921   0.7802
   5.500   1.1041   0.01079   0.00518  -0.1159   0.4867   0.8162
   5.750   1.1224   0.01072   0.00529  -0.1139   0.4804   0.8737
   6.000   1.1486   0.01067   0.00539  -0.1135   0.4736   1.0000
   6.250   1.1725   0.01078   0.00552  -0.1129   0.4672   1.0000
   6.500   1.1969   0.01097   0.00569  -0.1124   0.4614   1.0000
   6.750   1.2202   0.01106   0.00585  -0.1116   0.4549   1.0000
   7.000   1.2431   0.01125   0.00602  -0.1108   0.4485   1.0000
   7.250   1.2661   0.01137   0.00621  -0.1100   0.4420   1.0000
   7.500   1.2873   0.01154   0.00639  -0.1088   0.4343   1.0000
   7.750   1.3080   0.01168   0.00659  -0.1076   0.4263   1.0000
   8.000   1.3263   0.01190   0.00679  -0.1058   0.4176   1.0000
   8.250   1.3452   0.01206   0.00701  -0.1042   0.4067   1.0000
   8.500   1.3627   0.01230   0.00727  -0.1024   0.3943   1.0000
   8.750   1.3798   0.01258   0.00756  -0.1005   0.3809   1.0000
   9.000   1.3950   0.01293   0.00789  -0.0984   0.3645   1.0000
   9.250   1.4071   0.01340   0.00830  -0.0957   0.3441   1.0000
   9.500   1.4163   0.01403   0.00883  -0.0927   0.3159   1.0000
   9.750   1.4198   0.01492   0.00956  -0.0888   0.2807   1.0000
  10.000   1.4188   0.01608   0.01051  -0.0845   0.2445   1.0000
  10.250   1.4170   0.01735   0.01162  -0.0803   0.2135   1.0000
  10.500   1.4151   0.01868   0.01282  -0.0763   0.1857   1.0000
  10.750   1.4147   0.02004   0.01407  -0.0729   0.1625   1.0000
  11.000   1.4129   0.02156   0.01550  -0.0695   0.1428   1.0000
  11.250   1.4131   0.02308   0.01697  -0.0666   0.1266   1.0000
  11.500   1.4126   0.02476   0.01860  -0.0640   0.1118   1.0000
  11.750   1.4129   0.02651   0.02033  -0.0617   0.1003   1.0000
  12.000   1.4118   0.02847   0.02227  -0.0595   0.0891   1.0000
  12.250   1.4117   0.03049   0.02427  -0.0577   0.0789   1.0000
  12.500   1.4131   0.03250   0.02629  -0.0562   0.0708   1.0000
  12.750   1.4121   0.03480   0.02861  -0.0548   0.0630   1.0000
  13.000   1.4106   0.03726   0.03106  -0.0536   0.0553   1.0000
  13.250   1.4121   0.03954   0.03338  -0.0527   0.0499   1.0000
  13.500   1.4101   0.04222   0.03607  -0.0518   0.0448   1.0000
  13.750   1.4117   0.04465   0.03855  -0.0511   0.0406   1.0000
  14.000   1.4101   0.04747   0.04139  -0.0505   0.0365   1.0000
  14.250   1.4103   0.05018   0.04416  -0.0501   0.0333   1.0000
  14.500   1.4106   0.05295   0.04699  -0.0498   0.0305   1.0000
  14.750   1.4067   0.05628   0.05034  -0.0496   0.0274   1.0000
  15.000   1.4091   0.05894   0.05309  -0.0495   0.0251   1.0000
  15.250   1.4075   0.06215   0.05633  -0.0496   0.0228   1.0000
  15.500   1.4047   0.06558   0.05984  -0.0497   0.0207   1.0000
  15.750   1.4050   0.06871   0.06305  -0.0500   0.0188   1.0000
  16.000   1.4010   0.07244   0.06681  -0.0504   0.0168   1.0000
  16.250   1.3994   0.07594   0.07040  -0.0508   0.0154   1.0000
  16.500   1.3983   0.07942   0.07394  -0.0514   0.0135   1.0000
  16.750   1.3899   0.08403   0.07861  -0.0523   0.0124   1.0000
  17.000   1.3904   0.08744   0.08211  -0.0531   0.0109   1.0000
  17.250   1.3870   0.09146   0.08620  -0.0541   0.0101   1.0000
  17.500   1.3777   0.09646   0.09129  -0.0554   0.0092   1.0000
  17.750   1.3773   0.10019   0.09511  -0.0565   0.0081   1.0000
  18.000   1.3742   0.10436   0.09937  -0.0579   0.0076   1.0000
  18.250   1.3683   0.10904   0.10413  -0.0595   0.0073   1.0000
<< Back to EPPLER 855 AIRFOIL (e855-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 855 AIRFOIL (e855-il)