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EH 1.0/9.0 (eh1090-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EH 1.0/9.0 (eh1090-il)
Reynolds number: 200,000
Max Cl/Cd: 56.74 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1090-il-200000.txt
Download as CSV file: xf-eh1090-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 1.0/9.0                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.6443   0.05722   0.05355  -0.0194   1.0000   0.0581
  -7.250  -0.6341   0.05433   0.05062  -0.0189   1.0000   0.0603
  -7.000  -0.6265   0.05069   0.04677  -0.0185   1.0000   0.0633
  -6.750  -0.6261   0.04612   0.04162  -0.0176   1.0000   0.0697
  -6.500  -0.6189   0.03316   0.02729  -0.0130   1.0000   0.0318
  -6.250  -0.6023   0.02915   0.02300  -0.0116   1.0000   0.0300
  -6.000  -0.5847   0.02576   0.01916  -0.0099   1.0000   0.0293
  -5.750  -0.5640   0.02315   0.01615  -0.0084   1.0000   0.0294
  -5.500  -0.5416   0.02111   0.01380  -0.0070   1.0000   0.0299
  -5.250  -0.5182   0.01972   0.01217  -0.0058   1.0000   0.0310
  -5.000  -0.4974   0.01761   0.00999  -0.0046   1.0000   0.0347
  -4.750  -0.4754   0.01651   0.00885  -0.0033   1.0000   0.0380
  -4.500  -0.4539   0.01548   0.00770  -0.0018   1.0000   0.0419
  -4.250  -0.4353   0.01419   0.00646   0.0000   1.0000   0.0504
  -4.000  -0.4160   0.01328   0.00564   0.0016   1.0000   0.0718
  -3.750  -0.3972   0.01246   0.00501   0.0031   1.0000   0.1111
  -3.500  -0.3789   0.01174   0.00456   0.0046   1.0000   0.1634
  -3.250  -0.3622   0.01095   0.00421   0.0062   1.0000   0.2455
  -3.000  -0.3488   0.00998   0.00392   0.0082   1.0000   0.3948
  -2.750  -0.3387   0.00908   0.00390   0.0113   1.0000   0.5859
  -2.500  -0.3022   0.00862   0.00407   0.0102   0.9894   0.7575
  -2.250  -0.2615   0.00863   0.00420   0.0085   0.9794   0.8336
  -2.000  -0.2188   0.00877   0.00434   0.0065   0.9702   0.8824
  -1.750  -0.1749   0.00891   0.00438   0.0042   0.9596   0.9110
  -1.500  -0.1282   0.00904   0.00440   0.0011   0.9481   0.9289
  -1.250  -0.0748   0.00923   0.00446  -0.0032   0.9369   0.9443
  -1.000  -0.0199   0.00934   0.00446  -0.0080   0.9217   0.9560
  -0.750   0.0383   0.00940   0.00441  -0.0135   0.9027   0.9656
  -0.500   0.0901   0.00939   0.00426  -0.0179   0.8767   0.9761
  -0.250   0.1359   0.00933   0.00406  -0.0213   0.8477   0.9862
   0.000   0.1784   0.00925   0.00382  -0.0242   0.8180   0.9949
   0.250   0.2128   0.00916   0.00357  -0.0257   0.7891   1.0000
   0.500   0.2323   0.00916   0.00344  -0.0241   0.7620   1.0000
   0.750   0.2525   0.00918   0.00333  -0.0228   0.7365   1.0000
   1.000   0.2733   0.00921   0.00325  -0.0215   0.7129   1.0000
   1.250   0.2944   0.00927   0.00320  -0.0202   0.6910   1.0000
   1.500   0.3161   0.00932   0.00317  -0.0192   0.6689   1.0000
   2.000   0.3603   0.00947   0.00316  -0.0170   0.6277   1.0000
   2.250   0.3825   0.00958   0.00320  -0.0160   0.6085   1.0000
   2.500   0.4050   0.00969   0.00326  -0.0150   0.5887   1.0000
   2.750   0.4275   0.00982   0.00334  -0.0140   0.5696   1.0000
   3.000   0.4500   0.00998   0.00342  -0.0130   0.5513   1.0000
   3.250   0.4727   0.01012   0.00355  -0.0120   0.5318   1.0000
   3.500   0.4954   0.01029   0.00371  -0.0110   0.5129   1.0000
   3.750   0.5182   0.01046   0.00385  -0.0101   0.4932   1.0000
   4.000   0.5410   0.01064   0.00401  -0.0091   0.4726   1.0000
   4.250   0.5638   0.01085   0.00419  -0.0081   0.4533   1.0000
   4.500   0.5870   0.01105   0.00442  -0.0072   0.4340   1.0000
   4.750   0.6099   0.01129   0.00468  -0.0063   0.4149   1.0000
   5.000   0.6329   0.01152   0.00493  -0.0054   0.3928   1.0000
   5.250   0.6556   0.01177   0.00517  -0.0044   0.3691   1.0000
   5.500   0.6780   0.01201   0.00540  -0.0034   0.3372   1.0000
   5.750   0.6996   0.01233   0.00562  -0.0024   0.2951   1.0000
   6.000   0.7211   0.01277   0.00594  -0.0014   0.2502   1.0000
   6.250   0.7414   0.01343   0.00642  -0.0004   0.1899   1.0000
   6.500   0.7529   0.01553   0.00764   0.0014   0.0604   1.0000
   6.750   0.7678   0.01729   0.00934   0.0034   0.0389   1.0000
   7.000   0.7864   0.01845   0.01062   0.0050   0.0346   1.0000
   7.250   0.8048   0.01958   0.01181   0.0063   0.0305   1.0000
   7.500   0.8187   0.02164   0.01389   0.0082   0.0273   1.0000
   7.750   0.8386   0.02302   0.01541   0.0095   0.0262   1.0000
   8.000   0.8585   0.02475   0.01728   0.0109   0.0251   1.0000
   8.250   0.8785   0.02682   0.01959   0.0122   0.0244   1.0000
   8.500   0.8977   0.02913   0.02216   0.0135   0.0240   1.0000
   8.750   0.9151   0.03120   0.02448   0.0147   0.0228   1.0000
   9.000   0.9310   0.03287   0.02630   0.0157   0.0210   1.0000
   9.250   0.9438   0.03571   0.02947   0.0172   0.0208   1.0000
   9.500   0.9512   0.03942   0.03365   0.0192   0.0214   1.0000
   9.750   0.9475   0.04412   0.03893   0.0217   0.0227   1.0000
  10.000   0.9387   0.04872   0.04395   0.0238   0.0238   1.0000
  10.250   0.9263   0.05268   0.04818   0.0257   0.0248   1.0000
  10.500   0.9087   0.05625   0.05195   0.0274   0.0254   1.0000
  10.750   0.8890   0.06035   0.05623   0.0273   0.0257   1.0000
  11.000   0.8708   0.06503   0.06104   0.0254   0.0261   1.0000
  11.250   0.8515   0.07064   0.06678   0.0220   0.0265   1.0000
  11.500   0.8300   0.07782   0.07405   0.0178   0.0271   1.0000
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