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EPPLER E851 AIRFOIL (e851-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E851 AIRFOIL (e851-il)
Reynolds number: 100,000
Max Cl/Cd: 49.03 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e851-il-100000.txt
Download as CSV file: xf-e851-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E851 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4785   0.09749   0.09277  -0.0410   1.0000   0.0800
  -8.500  -0.4985   0.09461   0.09001  -0.0433   1.0000   0.0804
  -8.250  -0.5218   0.09171   0.08724  -0.0446   1.0000   0.0805
  -8.000  -0.4860   0.08764   0.08312  -0.0377   1.0000   0.0848
  -7.750  -0.4953   0.08509   0.08066  -0.0364   1.0000   0.0870
  -7.500  -0.5125   0.08267   0.07835  -0.0351   1.0000   0.0885
  -7.250  -0.5291   0.07914   0.07490  -0.0367   1.0000   0.0902
  -7.000  -0.5585   0.07482   0.07045  -0.0426   1.0000   0.0930
  -6.750  -0.5644   0.07038   0.06602  -0.0424   1.0000   0.0952
  -6.500  -0.5568   0.06767   0.06340  -0.0394   1.0000   0.0981
  -6.000  -0.5577   0.06007   0.05560  -0.0409   1.0000   0.1107
  -5.750  -0.5538   0.05630   0.05158  -0.0425   1.0000   0.1221
  -5.500  -0.5424   0.05303   0.04831  -0.0415   1.0000   0.1282
  -5.250  -0.5310   0.04954   0.04465  -0.0420   1.0000   0.1420
  -5.000  -0.4812   0.03622   0.02939  -0.0453   1.0000   0.0351
  -4.750  -0.4531   0.03353   0.02578  -0.0443   1.0000   0.0301
  -4.500  -0.4286   0.03079   0.02260  -0.0440   1.0000   0.0296
  -4.250  -0.4048   0.02740   0.01899  -0.0446   1.0000   0.0343
  -4.000  -0.3796   0.02596   0.01705  -0.0439   1.0000   0.0400
  -3.750  -0.3531   0.02405   0.01481  -0.0429   1.0000   0.0422
  -3.500  -0.3281   0.02173   0.01255  -0.0423   1.0000   0.0477
  -3.250  -0.3027   0.02004   0.01098  -0.0418   1.0000   0.0755
  -3.000  -0.2723   0.01606   0.00928  -0.0433   1.0000   0.4902
  -2.750  -0.2564   0.01636   0.00975  -0.0399   1.0000   0.6148
  -2.500  -0.2408   0.01663   0.01002  -0.0367   1.0000   0.6677
  -2.250  -0.2251   0.01681   0.01016  -0.0338   1.0000   0.7060
  -2.000  -0.2098   0.01691   0.01023  -0.0308   1.0000   0.7390
  -1.750  -0.1965   0.01698   0.01029  -0.0273   1.0000   0.7724
  -1.500  -0.1831   0.01696   0.01026  -0.0241   1.0000   0.8020
  -1.250  -0.1656   0.01690   0.01003  -0.0222   1.0000   0.8249
  -1.000  -0.1464   0.01684   0.00990  -0.0208   1.0000   0.8459
  -0.750  -0.1278   0.01676   0.00978  -0.0194   1.0000   0.8700
  -0.500  -0.1084   0.01666   0.00968  -0.0181   1.0000   0.9019
  -0.250  -0.0737   0.01645   0.00948  -0.0205   1.0000   0.9628
   0.000  -0.0417   0.01647   0.00934  -0.0235   1.0000   1.0000
   0.250  -0.0098   0.01676   0.00941  -0.0261   1.0000   1.0000
   0.500   0.0187   0.01711   0.00960  -0.0277   1.0000   1.0000
   0.750   0.0450   0.01748   0.00985  -0.0287   1.0000   1.0000
   1.000   0.0697   0.01789   0.01016  -0.0293   1.0000   1.0000
   1.250   0.0933   0.01832   0.01052  -0.0297   1.0000   1.0000
   1.500   0.1282   0.01902   0.01118  -0.0323   0.9954   1.0000
   1.750   0.1689   0.01978   0.01192  -0.0359   0.9870   1.0000
   2.000   0.2111   0.02062   0.01276  -0.0397   0.9791   1.0000
   2.250   0.2513   0.02129   0.01346  -0.0430   0.9701   1.0000
   2.500   0.2886   0.02182   0.01404  -0.0457   0.9598   1.0000
   2.750   0.3275   0.02236   0.01475  -0.0486   0.9492   1.0000
   3.000   0.3686   0.02284   0.01536  -0.0518   0.9379   1.0000
   3.250   0.4113   0.02324   0.01590  -0.0550   0.9258   1.0000
   3.500   0.4554   0.02347   0.01632  -0.0582   0.9120   1.0000
   3.750   0.5060   0.02338   0.01649  -0.0620   0.8955   1.0000
   4.000   0.5750   0.02221   0.01580  -0.0674   0.8708   1.0000
   4.250   0.6796   0.01851   0.01278  -0.0758   0.8374   1.0000
   4.500   0.7379   0.01505   0.00974  -0.0749   0.7585   1.0000
   4.750   0.7450   0.01758   0.00888  -0.0676   0.1726   1.0000
   5.000   0.7524   0.02028   0.01087  -0.0638   0.0985   1.0000
   5.250   0.7694   0.02244   0.01282  -0.0615   0.0644   1.0000
   5.500   0.7934   0.02479   0.01511  -0.0605   0.0429   1.0000
   5.750   0.8295   0.02900   0.01927  -0.0613   0.0363   1.0000
   6.000   0.8602   0.03196   0.02284  -0.0604   0.0349   1.0000
   6.250   0.8843   0.03444   0.02582  -0.0588   0.0319   1.0000
   6.500   0.9024   0.03664   0.02811  -0.0579   0.0263   1.0000
   6.750   0.9181   0.04142   0.03332  -0.0561   0.0253   1.0000
   7.000   0.9364   0.04376   0.03613  -0.0537   0.0262   1.0000
   7.250   0.9460   0.04873   0.04203  -0.0492   0.0310   1.0000
   7.500   0.9527   0.05347   0.04712  -0.0465   0.0340   1.0000
   9.500   0.8657   0.09476   0.09067  -0.0265   0.0870   1.0000
   9.750   0.8388   0.09881   0.09481  -0.0259   0.0870   1.0000
  10.000   0.8129   0.10362   0.09971  -0.0271   0.0869   1.0000
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