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EPPLER 908 AIRFOIL (e908-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 908 AIRFOIL (e908-il)
Reynolds number: 500,000
Max Cl/Cd: 107.21 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e908-il-500000.txt
Download as CSV file: xf-e908-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 908 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4199   0.08580   0.08367  -0.0531   0.9852   0.0079
  -8.000  -0.4148   0.08100   0.07888  -0.0578   0.9837   0.0080
  -7.750  -0.4208   0.07726   0.07518  -0.0595   0.9800   0.0081
  -7.500  -0.4280   0.07278   0.07073  -0.0631   0.9751   0.0081
  -7.250  -0.4156   0.06494   0.06281  -0.0744   0.9719   0.0081
  -7.000  -0.4118   0.05931   0.05709  -0.0786   0.9680   0.0081
  -6.750  -0.4081   0.05436   0.05201  -0.0805   0.9632   0.0082
  -6.500  -0.3926   0.04861   0.04606  -0.0843   0.9607   0.0085
  -6.250  -0.3714   0.04313   0.04031  -0.0877   0.9591   0.0090
  -6.000  -0.3619   0.03930   0.03624  -0.0869   0.9551   0.0094
  -5.750  -0.3468   0.03534   0.03198  -0.0865   0.9513   0.0102
  -5.500  -0.3209   0.03215   0.02843  -0.0871   0.9495   0.0122
  -4.000  -0.1591   0.01748   0.01162  -0.0881   0.9385   0.0155
  -3.750  -0.1279   0.01640   0.01046  -0.0885   0.9371   0.0130
  -3.500  -0.0965   0.01501   0.00893  -0.0892   0.9361   0.0122
  -3.250  -0.0636   0.01403   0.00782  -0.0903   0.9351   0.0119
  -3.000  -0.0290   0.01330   0.00688  -0.0917   0.9343   0.0128
  -2.750   0.0049   0.01205   0.00609  -0.0933   0.9336   0.1641
  -2.500   0.0234   0.01128   0.00602  -0.0922   0.9294   0.3552
  -2.250   0.0470   0.01050   0.00601  -0.0920   0.9264   0.5817
  -2.000   0.0766   0.01032   0.00607  -0.0924   0.9245   0.6853
  -1.750   0.1096   0.01024   0.00598  -0.0935   0.9232   0.7092
  -1.500   0.1440   0.01014   0.00581  -0.0950   0.9221   0.7247
  -1.250   0.1785   0.01005   0.00570  -0.0965   0.9212   0.7399
  -1.000   0.2139   0.00993   0.00559  -0.0982   0.9205   0.7518
  -0.750   0.2289   0.01003   0.00570  -0.0955   0.9140   0.7615
  -0.500   0.2610   0.00994   0.00561  -0.0966   0.9121   0.7716
  -0.250   0.2948   0.00983   0.00551  -0.0979   0.9107   0.7832
   0.000   0.3294   0.00967   0.00540  -0.0994   0.9097   0.7961
   0.250   0.3641   0.00949   0.00529  -0.1009   0.9088   0.8119
   0.500   0.3993   0.00929   0.00517  -0.1025   0.9082   0.8308
   0.750   0.4359   0.00911   0.00507  -0.1044   0.9076   0.8450
   1.000   0.4737   0.00894   0.00497  -0.1066   0.9071   0.8558
   1.250   0.4829   0.00895   0.00511  -0.1025   0.8985   0.8702
   1.500   0.5195   0.00875   0.00499  -0.1045   0.8973   0.8813
   1.750   0.5566   0.00854   0.00489  -0.1065   0.8963   0.8935
   2.000   0.5943   0.00829   0.00477  -0.1086   0.8951   0.9077
   2.250   0.6390   0.00696   0.00350  -0.1107   0.8760   0.9192
   2.500   0.6647   0.00620   0.00229  -0.1082   0.7373   0.9417
   2.750   0.6485   0.00824   0.00279  -0.0991   0.4313   1.0000
   3.000   0.6347   0.01117   0.00383  -0.0919   0.0182   1.0000
   3.250   0.6579   0.01157   0.00433  -0.0909   0.0135   1.0000
   3.500   0.6797   0.01206   0.00493  -0.0896   0.0125   1.0000
   3.750   0.7012   0.01253   0.00541  -0.0885   0.0088   1.0000
   4.000   0.7159   0.01371   0.00672  -0.0857   0.0072   1.0000
   4.250   0.7322   0.01585   0.00899  -0.0832   0.0067   1.0000
   4.500   0.7610   0.01734   0.01054  -0.0832   0.0069   1.0000
   4.750   0.7882   0.01795   0.01120  -0.0828   0.0077   1.0000
   6.000   0.9352   0.03347   0.02824  -0.0796   0.0137   1.0000
   6.250   0.9555   0.03520   0.03022  -0.0774   0.0119   1.0000
   6.500   0.9708   0.03771   0.03300  -0.0748   0.0107   1.0000
   6.750   0.9831   0.04040   0.03594  -0.0719   0.0098   1.0000
   7.000   0.9930   0.04312   0.03889  -0.0690   0.0091   1.0000
   7.250   1.0005   0.04588   0.04185  -0.0660   0.0086   1.0000
   7.500   1.0059   0.04865   0.04481  -0.0630   0.0082   1.0000
   7.750   1.0085   0.05156   0.04789  -0.0600   0.0079   1.0000
   8.000   1.0069   0.05485   0.05137  -0.0564   0.0076   1.0000
   8.250   1.0010   0.05850   0.05521  -0.0526   0.0074   1.0000
   8.500   0.9896   0.06257   0.05947  -0.0485   0.0072   1.0000
   8.750   0.9737   0.06637   0.06345  -0.0438   0.0071   1.0000
   9.000   0.9548   0.06987   0.06708  -0.0391   0.0070   1.0000
   9.250   0.9371   0.07327   0.07062  -0.0354   0.0069   1.0000
   9.500   0.9177   0.07713   0.07460  -0.0325   0.0069   1.0000
   9.750   0.9013   0.08088   0.07848  -0.0309   0.0069   1.0000
  10.000   0.8866   0.08484   0.08254  -0.0303   0.0069   1.0000
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