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EPPLER E850 AIRFOIL (e850-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER E850 AIRFOIL (e850-il)
Reynolds number: 100,000
Max Cl/Cd: 41.53 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e850-il-100000-n5.txt
Download as CSV file: xf-e850-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E850 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5036   0.09771   0.09261  -0.0310   1.0000   0.0137
  -9.000  -0.5075   0.09323   0.08821  -0.0325   1.0000   0.0131
  -8.500  -0.5177   0.08353   0.07866  -0.0364   1.0000   0.0119
  -8.000  -0.4530   0.06609   0.06148  -0.0395   1.0000   0.0108
  -7.750  -0.4675   0.06162   0.05705  -0.0404   1.0000   0.0107
  -7.500  -0.4872   0.05820   0.05364  -0.0403   1.0000   0.0106
  -7.250  -0.5524   0.06495   0.06002  -0.0410   1.0000   0.0108
  -7.000  -0.5595   0.06104   0.05603  -0.0409   1.0000   0.0106
  -6.750  -0.5629   0.05711   0.05196  -0.0406   1.0000   0.0104
  -6.500  -0.5629   0.05300   0.04765  -0.0404   1.0000   0.0102
  -6.250  -0.5584   0.04899   0.04339  -0.0402   1.0000   0.0100
  -6.000  -0.5499   0.04509   0.03918  -0.0400   1.0000   0.0098
  -5.750  -0.5378   0.04122   0.03494  -0.0397   1.0000   0.0096
  -5.500  -0.5221   0.03753   0.03082  -0.0394   1.0000   0.0094
  -5.250  -0.5035   0.03420   0.02703  -0.0390   1.0000   0.0092
  -5.000  -0.4823   0.03114   0.02348  -0.0385   1.0000   0.0091
  -4.750  -0.4594   0.02841   0.02028  -0.0380   1.0000   0.0089
  -4.500  -0.4351   0.02597   0.01726  -0.0373   1.0000   0.0089
  -4.250  -0.4105   0.02398   0.01494  -0.0366   1.0000   0.0089
  -4.000  -0.3860   0.02223   0.01294  -0.0358   1.0000   0.0090
  -3.750  -0.3617   0.02073   0.01124  -0.0350   1.0000   0.0093
  -3.500  -0.3374   0.01951   0.00981  -0.0344   1.0000   0.0098
  -3.250  -0.3124   0.01857   0.00859  -0.0340   1.0000   0.0106
  -3.000  -0.2872   0.01769   0.00756  -0.0338   1.0000   0.0178
  -2.750  -0.2605   0.01460   0.00669  -0.0355   1.0000   0.4351
  -2.500  -0.2421   0.01446   0.00703  -0.0330   1.0000   0.6047
  -2.250  -0.2214   0.01449   0.00701  -0.0312   1.0000   0.6494
  -2.000  -0.2005   0.01451   0.00696  -0.0296   1.0000   0.6800
  -1.750  -0.1779   0.01449   0.00670  -0.0287   1.0000   0.6957
  -1.500  -0.1508   0.01452   0.00661  -0.0288   0.9984   0.7093
  -1.250  -0.1201   0.01462   0.00661  -0.0296   0.9953   0.7232
  -1.000  -0.0891   0.01472   0.00664  -0.0305   0.9923   0.7376
  -0.750  -0.0593   0.01478   0.00666  -0.0312   0.9888   0.7526
  -0.500  -0.0285   0.01489   0.00668  -0.0321   0.9852   0.7691
  -0.250   0.0033   0.01502   0.00682  -0.0332   0.9819   0.7871
   0.000   0.0321   0.01506   0.00689  -0.0336   0.9777   0.8079
   0.250   0.0613   0.01509   0.00699  -0.0341   0.9732   0.8329
   0.500   0.0937   0.01516   0.00716  -0.0352   0.9694   0.8664
   0.750   0.1284   0.01510   0.00723  -0.0368   0.9648   0.9239
   1.000   0.1666   0.01513   0.00729  -0.0396   0.9594   1.0000
   1.250   0.2046   0.01536   0.00751  -0.0422   0.9554   1.0000
   1.500   0.2347   0.01546   0.00765  -0.0432   0.9482   1.0000
   1.750   0.2735   0.01564   0.00789  -0.0458   0.9434   1.0000
   2.000   0.3045   0.01573   0.00806  -0.0467   0.9351   1.0000
   2.250   0.3432   0.01581   0.00828  -0.0491   0.9282   1.0000
   2.500   0.3809   0.01580   0.00863  -0.0511   0.9187   1.0000
   2.750   0.4204   0.01562   0.00868  -0.0531   0.9061   1.0000
   3.000   0.4657   0.01495   0.00830  -0.0554   0.8834   1.0000
   3.250   0.5110   0.01375   0.00740  -0.0565   0.8402   1.0000
   3.500   0.5819   0.01401   0.00601  -0.0611   0.3723   1.0000
   3.750   0.5937   0.01643   0.00692  -0.0589   0.1009   1.0000
   4.000   0.6129   0.01794   0.00811  -0.0573   0.0458   1.0000
   4.250   0.6331   0.01929   0.00953  -0.0556   0.0304   1.0000
   4.500   0.6534   0.02076   0.01111  -0.0540   0.0220   1.0000
   4.750   0.6756   0.02222   0.01263  -0.0530   0.0168   1.0000
   5.000   0.7007   0.02522   0.01575  -0.0521   0.0151   1.0000
   5.250   0.7284   0.02750   0.01833  -0.0515   0.0145   1.0000
   5.500   0.7546   0.03018   0.02141  -0.0505   0.0140   1.0000
   5.750   0.7777   0.03321   0.02494  -0.0489   0.0137   1.0000
   6.000   0.7970   0.03655   0.02879  -0.0469   0.0135   1.0000
   6.250   0.8128   0.04010   0.03284  -0.0446   0.0134   1.0000
   6.500   0.8251   0.04394   0.03715  -0.0420   0.0134   1.0000
   6.750   0.8359   0.04768   0.04133  -0.0394   0.0130   1.0000
   7.000   0.8444   0.05153   0.04558  -0.0369   0.0122   1.0000
   7.250   0.8502   0.05550   0.04990  -0.0345   0.0113   1.0000
   7.500   0.8528   0.05955   0.05425  -0.0323   0.0105   1.0000
   7.750   0.8523   0.06363   0.05858  -0.0303   0.0099   1.0000
   8.000   0.8490   0.06753   0.06267  -0.0286   0.0091   1.0000
   8.250   0.8408   0.07191   0.06723  -0.0270   0.0096   1.0000
   8.500   0.8283   0.07581   0.07124  -0.0253   0.0095   1.0000
   8.750   0.8143   0.08012   0.07564  -0.0249   0.0097   1.0000
   9.000   0.8001   0.08518   0.08076  -0.0262   0.0103   1.0000
   9.250   0.7885   0.09090   0.08651  -0.0295   0.0106   1.0000
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