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EPPLER 417 AIRFOIL (e417-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 417 AIRFOIL (e417-il)
Reynolds number: 500,000
Max Cl/Cd: 103.14 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e417-il-500000.txt
Download as CSV file: xf-e417-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 417 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.5458   0.02812   0.02238  -0.0868   0.9396   0.0191
  -7.250  -0.5208   0.02573   0.01992  -0.0866   0.9387   0.0182
  -7.000  -0.4943   0.02387   0.01793  -0.0866   0.9378   0.0174
  -6.750  -0.4661   0.02250   0.01645  -0.0870   0.9370   0.0171
  -6.500  -0.4368   0.02143   0.01531  -0.0878   0.9362   0.0170
  -6.250  -0.4055   0.02052   0.01434  -0.0891   0.9355   0.0171
  -6.000  -0.3921   0.01990   0.01367  -0.0870   0.9308   0.0175
  -5.750  -0.3669   0.01920   0.01293  -0.0873   0.9280   0.0179
  -5.500  -0.3356   0.01851   0.01218  -0.0888   0.9264   0.0183
  -5.250  -0.3021   0.01793   0.01154  -0.0907   0.9252   0.0188
  -5.000  -0.2664   0.01711   0.01066  -0.0933   0.9242   0.0204
  -4.750  -0.2305   0.01664   0.01015  -0.0956   0.9234   0.0224
  -4.500  -0.1942   0.01627   0.00972  -0.0978   0.9227   0.0246
  -4.250  -0.1551   0.01546   0.00913  -0.1011   0.9222   0.0722
  -4.000  -0.1049   0.01284   0.00817  -0.1096   0.9227   0.4734
  -3.750  -0.0680   0.01275   0.00825  -0.1117   0.9220   0.5438
  -3.500  -0.0315   0.01280   0.00825  -0.1136   0.9214   0.5659
  -3.250  -0.0205   0.01300   0.00846  -0.1103   0.9141   0.5782
  -3.000   0.0115   0.01314   0.00858  -0.1110   0.9121   0.5939
  -2.750   0.0449   0.01345   0.00889  -0.1118   0.9105   0.6120
  -2.500   0.0788   0.01349   0.00898  -0.1129   0.9094   0.6180
  -2.250   0.1152   0.01343   0.00887  -0.1147   0.9086   0.6224
  -1.750   0.1683   0.01340   0.00876  -0.1145   0.9012   0.6285
  -1.500   0.2007   0.01327   0.00864  -0.1155   0.8989   0.6307
  -1.250   0.2367   0.01308   0.00844  -0.1172   0.8974   0.6330
  -1.000   0.2737   0.01286   0.00821  -0.1191   0.8962   0.6356
  -0.750   0.3121   0.01259   0.00792  -0.1213   0.8953   0.6383
  -0.500   0.3523   0.01226   0.00756  -0.1239   0.8945   0.6411
  -0.250   0.3934   0.01184   0.00715  -0.1266   0.8940   0.6434
   0.000   0.4100   0.01190   0.00723  -0.1243   0.8865   0.6453
   0.250   0.4447   0.01163   0.00700  -0.1257   0.8846   0.6472
   0.500   0.4814   0.01134   0.00674  -0.1275   0.8834   0.6495
   0.750   0.5200   0.01101   0.00643  -0.1297   0.8822   0.6520
   1.000   0.5633   0.01055   0.00598  -0.1328   0.8811   0.6546
   1.250   0.6096   0.01009   0.00553  -0.1366   0.8798   0.6570
   1.750   0.6632   0.00966   0.00520  -0.1360   0.8679   0.6610
   2.000   0.7091   0.00930   0.00489  -0.1397   0.8651   0.6630
   2.250   0.7237   0.00921   0.00483  -0.1367   0.8540   0.6651
   2.500   0.7500   0.00906   0.00471  -0.1363   0.8441   0.6676
   2.750   0.7846   0.00886   0.00454  -0.1377   0.8341   0.6703
   3.000   0.8138   0.00876   0.00443  -0.1379   0.8197   0.6727
   3.250   0.8420   0.00864   0.00433  -0.1378   0.7995   0.6746
   3.500   0.8748   0.00859   0.00423  -0.1388   0.7715   0.6764
   3.750   0.9025   0.00875   0.00423  -0.1386   0.7252   0.6785
   4.250   0.9185   0.00988   0.00469  -0.1302   0.5931   0.6834
   4.500   0.9237   0.01061   0.00508  -0.1257   0.5265   0.6859
   4.750   0.9330   0.01133   0.00548  -0.1222   0.4638   0.6881
   5.000   0.9454   0.01195   0.00587  -0.1194   0.4089   0.6900
   5.250   0.9599   0.01257   0.00628  -0.1171   0.3580   0.6921
   5.500   0.9753   0.01325   0.00672  -0.1150   0.3036   0.6944
   5.750   0.9908   0.01400   0.00720  -0.1131   0.2485   0.6969
   6.000   1.0068   0.01479   0.00771  -0.1113   0.1932   0.6994
   6.250   1.0225   0.01566   0.00826  -0.1095   0.1378   0.7018
   6.500   1.0376   0.01658   0.00888  -0.1077   0.0856   0.7039
   6.750   1.0490   0.01782   0.00975  -0.1051   0.0299   0.7059
   7.000   1.0666   0.01856   0.01047  -0.1035   0.0197   0.7081
   7.250   1.0857   0.01915   0.01111  -0.1021   0.0175   0.7107
   7.500   1.1030   0.01993   0.01196  -0.1004   0.0159   0.7134
   7.750   1.1217   0.02058   0.01268  -0.0991   0.0153   0.7160
   8.000   1.1394   0.02130   0.01347  -0.0976   0.0148   0.7184
   8.250   1.1560   0.02203   0.01430  -0.0959   0.0142   0.7206
   8.500   1.1715   0.02285   0.01520  -0.0941   0.0136   0.7229
   8.750   1.1858   0.02376   0.01618  -0.0922   0.0130   0.7255
   9.000   1.1982   0.02486   0.01736  -0.0900   0.0126   0.7283
   9.250   1.2085   0.02618   0.01876  -0.0876   0.0124   0.7311
   9.500   1.2176   0.02777   0.02044  -0.0851   0.0121   0.7336
   9.750   1.2288   0.02932   0.02211  -0.0830   0.0120   0.7358
  10.000   1.2435   0.03044   0.02336  -0.0813   0.0118   0.7383
  10.250   1.2581   0.03176   0.02479  -0.0798   0.0117   0.7410
  10.500   1.2730   0.03312   0.02627  -0.0783   0.0116   0.7439
  10.750   1.2880   0.03463   0.02790  -0.0769   0.0115   0.7469
  11.000   1.3029   0.03628   0.02970  -0.0755   0.0114   0.7496
  11.250   1.3173   0.03810   0.03171  -0.0741   0.0113   0.7523
  11.500   1.3307   0.04019   0.03399  -0.0726   0.0113   0.7552
  11.750   1.3420   0.04258   0.03660  -0.0710   0.0113   0.7581
  12.000   1.3501   0.04539   0.03966  -0.0692   0.0114   0.7609
  12.250   1.3540   0.04870   0.04322  -0.0671   0.0116   0.7635
  12.500   1.3528   0.05218   0.04698  -0.0647   0.0117   0.7658
  12.750   1.3592   0.05708   0.05209  -0.0636   0.0121   0.7683
  13.000   1.3535   0.05973   0.05494  -0.0611   0.0121   0.7713
  13.250   1.3452   0.06277   0.05818  -0.0589   0.0121   0.7745
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