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EH 1.0/9.0 (eh1090-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EH 1.0/9.0 (eh1090-il)
Reynolds number: 50,000
Max Cl/Cd: 29.17 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1090-il-50000.txt
Download as CSV file: xf-eh1090-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5506   0.10074   0.09424   0.0176   1.0000   0.2938
  -8.250  -0.5472   0.09741   0.09096   0.0179   1.0000   0.3105
  -8.000  -0.5593   0.09514   0.08880   0.0176   1.0000   0.3286
  -7.750  -0.5390   0.09089   0.08450   0.0197   1.0000   0.3540
  -7.250  -0.5276   0.08370   0.07740   0.0211   1.0000   0.3891
  -7.000  -0.6258   0.06597   0.05938  -0.0143   1.0000   0.1633
  -6.750  -0.6135   0.05812   0.05122  -0.0161   1.0000   0.1297
  -6.500  -0.6071   0.05226   0.04484  -0.0166   1.0000   0.1170
  -6.250  -0.5993   0.04714   0.03875  -0.0159   1.0000   0.1077
  -6.000  -0.5844   0.04332   0.03443  -0.0147   1.0000   0.1079
  -5.750  -0.5656   0.04000   0.03117  -0.0139   1.0000   0.1137
  -5.500  -0.5464   0.03666   0.02735  -0.0126   1.0000   0.1155
  -5.250  -0.5246   0.03344   0.02357  -0.0112   1.0000   0.1179
  -5.000  -0.5017   0.03067   0.02045  -0.0100   1.0000   0.1258
  -4.750  -0.4785   0.02843   0.01798  -0.0088   1.0000   0.1418
  -4.500  -0.4527   0.02610   0.01537  -0.0075   1.0000   0.1621
  -4.250  -0.4281   0.02391   0.01343  -0.0063   1.0000   0.1999
  -4.000  -0.4042   0.02178   0.01167  -0.0050   1.0000   0.2637
  -3.750  -0.3851   0.01932   0.01012  -0.0030   1.0000   0.3805
  -3.500  -0.3800   0.01724   0.01041   0.0063   1.0000   0.7399
  -3.250  -0.1273   0.01908   0.01009  -0.0239   1.0000   1.0000
  -3.000  -0.1127   0.01847   0.00925  -0.0234   1.0000   1.0000
  -2.750  -0.0973   0.01798   0.00859  -0.0228   1.0000   1.0000
  -2.500  -0.0815   0.01757   0.00804  -0.0220   1.0000   1.0000
  -2.250  -0.0658   0.01723   0.00757  -0.0210   1.0000   1.0000
  -2.000  -0.0505   0.01695   0.00720  -0.0199   1.0000   1.0000
  -1.750  -0.0360   0.01673   0.00690  -0.0185   1.0000   1.0000
  -1.500  -0.0229   0.01657   0.00665  -0.0169   1.0000   1.0000
  -1.250  -0.0125   0.01647   0.00653  -0.0148   1.0000   1.0000
  -1.000  -0.0061   0.01646   0.00650  -0.0120   1.0000   1.0000
  -0.750  -0.0046   0.01655   0.00658  -0.0086   1.0000   1.0000
  -0.500  -0.0083   0.01676   0.00677  -0.0046   1.0000   1.0000
  -0.250  -0.0130   0.01707   0.00702  -0.0006   1.0000   1.0000
   0.000  -0.0153   0.01745   0.00733   0.0027   1.0000   1.0000
   0.250  -0.0144   0.01790   0.00770   0.0054   1.0000   1.0000
   0.500  -0.0102   0.01841   0.00812   0.0073   1.0000   1.0000
   0.750   0.0371   0.01925   0.00893   0.0013   0.9855   1.0000
   1.000   0.1118   0.02008   0.00979  -0.0091   0.9597   1.0000
   1.250   0.1847   0.02066   0.01045  -0.0186   0.9339   1.0000
   1.500   0.2629   0.02095   0.01088  -0.0285   0.9088   1.0000
   1.750   0.3271   0.02105   0.01114  -0.0349   0.8830   1.0000
   2.000   0.3759   0.02115   0.01136  -0.0381   0.8570   1.0000
   2.250   0.4111   0.02136   0.01165  -0.0387   0.8309   1.0000
   2.500   0.4371   0.02170   0.01210  -0.0376   0.8050   1.0000
   2.750   0.4600   0.02210   0.01255  -0.0359   0.7800   1.0000
   3.000   0.4823   0.02248   0.01299  -0.0340   0.7565   1.0000
   3.250   0.5040   0.02284   0.01340  -0.0318   0.7338   1.0000
   3.500   0.5230   0.02338   0.01406  -0.0296   0.7098   1.0000
   3.750   0.5432   0.02383   0.01458  -0.0273   0.6873   1.0000
   4.000   0.5637   0.02426   0.01508  -0.0249   0.6648   1.0000
   4.250   0.5830   0.02485   0.01577  -0.0228   0.6410   1.0000
   4.500   0.6036   0.02532   0.01631  -0.0205   0.6182   1.0000
   4.750   0.6246   0.02573   0.01680  -0.0181   0.5947   1.0000
   5.000   0.6444   0.02631   0.01757  -0.0159   0.5693   1.0000
   5.250   0.6649   0.02680   0.01817  -0.0136   0.5433   1.0000
   5.500   0.6857   0.02721   0.01868  -0.0112   0.5160   1.0000
   5.750   0.7067   0.02755   0.01912  -0.0088   0.4871   1.0000
   6.000   0.7283   0.02784   0.01948  -0.0062   0.4567   1.0000
   6.250   0.7481   0.02795   0.01972  -0.0034   0.4188   1.0000
   6.500   0.7672   0.02731   0.01889   0.0002   0.3699   1.0000
   6.750   0.7798   0.02673   0.01802   0.0041   0.3012   1.0000
   7.000   0.7845   0.02781   0.01838   0.0081   0.2074   1.0000
   7.250   0.7957   0.03046   0.02057   0.0108   0.1486   1.0000
   7.500   0.8125   0.03289   0.02287   0.0125   0.1205   1.0000
   7.750   0.8345   0.03581   0.02589   0.0139   0.1071   1.0000
   8.000   0.8555   0.03895   0.02911   0.0150   0.0986   1.0000
   8.250   0.8716   0.04183   0.03242   0.0164   0.0913   1.0000
   8.500   0.8881   0.04545   0.03615   0.0173   0.0865   1.0000
   8.750   0.8971   0.04949   0.04086   0.0188   0.0856   1.0000
   9.000   0.9017   0.05383   0.04571   0.0201   0.0855   1.0000
   9.250   0.9027   0.05841   0.05068   0.0211   0.0858   1.0000
   9.500   0.9017   0.06313   0.05569   0.0219   0.0865   1.0000
   9.750   0.9013   0.06821   0.06099   0.0223   0.0872   1.0000
  10.000   0.8467   0.07284   0.06617   0.0217   0.0903   1.0000
  10.250   0.8051   0.07979   0.07313   0.0171   0.0934   1.0000
  10.500   0.7783   0.08794   0.08126   0.0110   0.0964   1.0000
  10.750   0.7667   0.09516   0.08844   0.0071   0.0985   1.0000
  11.000   0.7671   0.10095   0.09424   0.0054   0.1004   1.0000
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