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EPPLER 422 AIRFOIL (e422-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 422 AIRFOIL (e422-il)
Reynolds number: 50,000
Max Cl/Cd: 23.17 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e422-il-50000-n5.txt
Download as CSV file: xf-e422-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 422 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.1440   0.13739   0.13108  -0.0406   1.0000   0.0898
 -10.750  -0.1438   0.13568   0.12951  -0.0401   1.0000   0.0916
 -10.500  -0.1386   0.13345   0.12740  -0.0418   0.9940   0.0936
 -10.000  -0.1141   0.12795   0.12196  -0.0535   0.9456   0.0959
  -9.750  -0.0897   0.12261   0.11663  -0.0575   0.9237   0.0967
  -9.500  -0.0592   0.11710   0.11105  -0.0617   0.9051   0.0981
  -9.250  -0.0325   0.11257   0.10645  -0.0666   0.8868   0.0997
  -9.000  -0.0086   0.10841   0.10218  -0.0717   0.8681   0.1007
  -8.750   0.0121   0.10454   0.09820  -0.0762   0.8492   0.1010
  -8.500   0.0273   0.10106   0.09461  -0.0800   0.8305   0.0996
  -8.000   0.0383   0.09177   0.08505  -0.0870   0.7988   0.0578
  -7.750   0.0497   0.08889   0.08208  -0.0880   0.7837   0.0565
  -7.250   0.0614   0.08356   0.07664  -0.0902   0.7573   0.0552
  -7.000   0.0645   0.08103   0.07408  -0.0911   0.7459   0.0550
  -6.750   0.0665   0.07859   0.07161  -0.0919   0.7356   0.0547
  -6.500   0.0699   0.07594   0.06896  -0.0934   0.7253   0.0546
  -6.250   0.0759   0.07309   0.06607  -0.0953   0.7165   0.0545
  -6.000   0.0813   0.07005   0.06302  -0.0975   0.7076   0.0540
  -5.750   0.0890   0.06671   0.05963  -0.1002   0.6995   0.0539
  -5.500   0.0983   0.06287   0.05571  -0.1036   0.6923   0.0534
  -5.250   0.1089   0.05844   0.05121  -0.1076   0.6844   0.0528
  -5.000   0.1242   0.05272   0.04525  -0.1129   0.6785   0.0525
  -4.750   0.1398   0.04762   0.03988  -0.1169   0.6716   0.0529
  -4.500   0.1589   0.04388   0.03579  -0.1194   0.6645   0.0547
  -4.250   0.1821   0.03870   0.02977  -0.1227   0.6593   0.0579
  -4.000   0.2030   0.03731   0.02820  -0.1227   0.6519   0.0612
  -3.750   0.2271   0.03480   0.02508  -0.1232   0.6455   0.0655
  -3.500   0.2540   0.03367   0.02371  -0.1233   0.6404   0.0722
  -3.250   0.2760   0.03284   0.02270  -0.1228   0.6332   0.0800
  -3.000   0.3009   0.03207   0.02173  -0.1225   0.6271   0.0920
  -2.750   0.3290   0.03144   0.02087  -0.1224   0.6223   0.1106
  -2.500   0.3516   0.03150   0.02090  -0.1217   0.6159   0.1320
  -2.250   0.3755   0.03154   0.02085  -0.1211   0.6100   0.1564
  -2.000   0.4024   0.03151   0.02063  -0.1207   0.6053   0.1820
  -1.750   0.4277   0.03147   0.02033  -0.1203   0.6005   0.2053
  -1.500   0.4480   0.03165   0.02046  -0.1192   0.5942   0.2223
  -1.250   0.4728   0.03161   0.02028  -0.1187   0.5895   0.2392
  -1.000   0.5019   0.03145   0.01990  -0.1186   0.5857   0.2561
  -0.750   0.5246   0.03168   0.02003  -0.1180   0.5806   0.2701
  -0.500   0.5469   0.03197   0.02022  -0.1174   0.5751   0.2849
  -0.250   0.5738   0.03203   0.02011  -0.1171   0.5707   0.3013
   0.000   0.6037   0.03196   0.01987  -0.1172   0.5674   0.3192
   0.250   0.6235   0.03249   0.02037  -0.1164   0.5624   0.3357
   0.500   0.6424   0.03306   0.02095  -0.1154   0.5573   0.3532
   0.750   0.6667   0.03326   0.02116  -0.1149   0.5533   0.3744
   1.000   0.6948   0.03325   0.02112  -0.1148   0.5500   0.4024
   1.250   0.7160   0.03362   0.02158  -0.1140   0.5460   0.4343
   1.500   0.7258   0.03460   0.02283  -0.1122   0.5404   0.4713
   1.750   0.7422   0.03467   0.02342  -0.1107   0.5364   0.5737
   2.250   0.8082   0.03488   0.02352  -0.1120   0.5306   1.0000
   2.500   0.8021   0.03722   0.02592  -0.1088   0.5237   1.0000
   2.750   0.8146   0.03852   0.02711  -0.1073   0.5191   1.0000
   3.000   0.8363   0.03928   0.02771  -0.1066   0.5159   1.0000
   3.250   0.8639   0.03970   0.02794  -0.1064   0.5135   1.0000
   3.500   0.8355   0.04359   0.03198  -0.1018   0.5051   1.0000
   3.750   0.8432   0.04523   0.03355  -0.1001   0.5004   1.0000
   4.000   0.8659   0.04595   0.03415  -0.0995   0.4976   1.0000
   4.250   0.8940   0.04636   0.03443  -0.0993   0.4956   1.0000
   4.750   0.8428   0.05473   0.04293  -0.0928   0.4798   1.0000
   5.000   0.8708   0.05500   0.04308  -0.0923   0.4781   1.0000
   5.750   0.8218   0.06848   0.05667  -0.0895   0.4558   1.0000
   6.000   0.8164   0.07195   0.06015  -0.0890   0.4494   1.0000
   6.250   0.8348   0.07316   0.06131  -0.0884   0.4462   1.0000
   6.500   0.8592   0.07381   0.06190  -0.0878   0.4439   1.0000
   7.000   0.8426   0.08147   0.06961  -0.0870   0.4308   1.0000
   7.250   0.8640   0.08244   0.07054  -0.0864   0.4280   1.0000
   7.500   0.8901   0.08297   0.07104  -0.0859   0.4261   1.0000
   7.750   0.8568   0.08929   0.07746  -0.0859   0.4158   1.0000
   8.000   0.8747   0.09061   0.07877  -0.0855   0.4124   1.0000
   8.250   0.8992   0.09127   0.07940  -0.0849   0.4101   1.0000
   8.500   0.8756   0.09676   0.08497  -0.0851   0.4010   1.0000
   8.750   0.8901   0.09842   0.08664  -0.0848   0.3969   1.0000
   9.000   0.9136   0.09915   0.08738  -0.0842   0.3942   1.0000
   9.250   0.8970   0.10402   0.09233  -0.0845   0.3859   1.0000
   9.500   0.9093   0.10588   0.09421  -0.0843   0.3812   1.0000
   9.750   0.9320   0.10665   0.09498  -0.0837   0.3783   1.0000
  10.000   0.9182   0.11133   0.09973  -0.0842   0.3701   1.0000
  10.250   0.9309   0.11314   0.10157  -0.0840   0.3654   1.0000
  10.500   0.9537   0.11384   0.10232  -0.0834   0.3624   1.0000
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