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EPPLER 636 AIRFOIL (e636-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 636 AIRFOIL (e636-il)
Reynolds number: 50,000
Max Cl/Cd: 14.74 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e636-il-50000.txt
Download as CSV file: xf-e636-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 636 AIRFOIL                              
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4619   0.12786   0.12156   0.0054   1.0000   0.1823
 -10.000  -0.4306   0.12130   0.11499   0.0077   1.0000   0.1927
  -9.750  -0.4537   0.12073   0.11457   0.0039   1.0000   0.1977
  -9.500  -0.4245   0.11477   0.10860   0.0060   1.0000   0.2093
  -9.250  -0.4251   0.11146   0.10537   0.0048   1.0000   0.2165
  -9.000  -0.4299   0.10920   0.10320   0.0035   1.0000   0.2276
  -8.750  -0.4133   0.10481   0.09885   0.0044   1.0000   0.2398
  -8.500  -0.4074   0.10130   0.09542   0.0044   1.0000   0.2520
  -8.250  -0.4045   0.09810   0.09231   0.0042   1.0000   0.2658
  -8.000  -0.3995   0.09493   0.08923   0.0046   1.0000   0.2826
  -7.750  -0.3993   0.09211   0.08651   0.0049   1.0000   0.3021
  -7.500  -0.4112   0.08996   0.08451   0.0048   1.0000   0.3203
  -7.250  -0.3917   0.08639   0.08100   0.0072   1.0000   0.3496
  -7.000  -0.3900   0.08394   0.07868   0.0094   1.0000   0.3817
  -6.750  -0.3497   0.08015   0.07489   0.0136   1.0000   0.4366
  -6.500  -0.3228   0.07724   0.07204   0.0174   1.0000   0.4973
  -6.250  -0.3016   0.07502   0.06990   0.0217   1.0000   0.5660
  -6.000  -0.2412   0.07060   0.06546   0.0231   1.0000   0.6550
  -5.750  -0.1894   0.06644   0.06134   0.0222   1.0000   0.7380
  -5.000  -0.2462   0.06069   0.05633   0.0219   1.0000   0.6579
  -4.750  -0.2970   0.06053   0.05650   0.0222   0.9842   0.6166
  -4.500  -0.3525   0.04843   0.04291  -0.0213   0.9289   0.2834
  -4.250  -0.2923   0.04288   0.03559  -0.0282   0.9120   0.1624
  -4.000  -0.2534   0.03967   0.03164  -0.0294   0.8930   0.1409
  -3.750  -0.2164   0.03728   0.02865  -0.0300   0.8744   0.1309
  -3.500  -0.1823   0.03520   0.02604  -0.0300   0.8566   0.1261
  -3.250  -0.1500   0.03361   0.02407  -0.0298   0.8395   0.1270
  -3.000  -0.1180   0.03224   0.02236  -0.0294   0.8230   0.1290
  -2.750  -0.0834   0.03095   0.02075  -0.0294   0.8072   0.1307
  -2.500   0.1732   0.02296   0.01550  -0.0602   0.7851   1.0000
  -2.250   0.1935   0.02328   0.01530  -0.0589   0.7692   1.0000
  -2.000   0.2135   0.02363   0.01527  -0.0577   0.7544   1.0000
  -1.750   0.2334   0.02401   0.01534  -0.0566   0.7411   1.0000
  -1.500   0.2539   0.02440   0.01542  -0.0555   0.7287   1.0000
  -1.250   0.2747   0.02485   0.01564  -0.0546   0.7165   1.0000
  -1.000   0.2955   0.02539   0.01598  -0.0541   0.7044   1.0000
  -0.750   0.3161   0.02595   0.01635  -0.0533   0.6937   1.0000
  -0.500   0.3361   0.02641   0.01658  -0.0519   0.6846   1.0000
  -0.250   0.3569   0.02718   0.01725  -0.0517   0.6736   1.0000
   0.000   0.3771   0.02789   0.01782  -0.0510   0.6646   1.0000
   0.250   0.3971   0.02856   0.01835  -0.0501   0.6559   1.0000
   0.500   0.4165   0.02954   0.01927  -0.0498   0.6469   1.0000
   0.750   0.4363   0.03018   0.01976  -0.0485   0.6395   1.0000
   1.000   0.4541   0.03145   0.02102  -0.0485   0.6306   1.0000
   1.250   0.4738   0.03214   0.02159  -0.0471   0.6240   1.0000
   1.500   0.4889   0.03373   0.02320  -0.0472   0.6157   1.0000
   1.750   0.5080   0.03453   0.02390  -0.0459   0.6092   1.0000
   2.000   0.5200   0.03639   0.02579  -0.0458   0.6019   1.0000
   2.250   0.5355   0.03768   0.02705  -0.0448   0.5956   1.0000
   2.500   0.5478   0.03933   0.02868  -0.0439   0.5894   1.0000
   2.750   0.5537   0.04152   0.03089  -0.0432   0.5828   1.0000
   3.000   0.5735   0.04243   0.03174  -0.0418   0.5778   1.0000
   3.250   0.5652   0.04577   0.03513  -0.0410   0.5727   1.0000
   3.500   0.5607   0.04851   0.03787  -0.0397   0.5680   1.0000
   3.750   0.5754   0.04999   0.03933  -0.0384   0.5635   1.0000
   4.000   0.5741   0.05264   0.04196  -0.0370   0.5607   1.0000
   4.250   0.5538   0.05623   0.04553  -0.0350   0.5600   1.0000
   4.500   0.5385   0.05929   0.04854  -0.0327   0.5603   1.0000
   4.750   0.5329   0.06201   0.05123  -0.0309   0.5621   1.0000
   6.250   0.4138   0.08055   0.06965  -0.0246   0.6721   1.0000
   6.500   0.4209   0.08229   0.07137  -0.0235   0.6599   1.0000
   6.750   0.4348   0.08466   0.07373  -0.0232   0.6501   1.0000
   7.000   0.4643   0.08796   0.07706  -0.0244   0.6387   1.0000
   7.250   0.4658   0.08917   0.07826  -0.0227   0.6251   1.0000
   7.500   0.4678   0.09078   0.07989  -0.0214   0.6134   1.0000
   7.750   0.4796   0.09323   0.08235  -0.0211   0.6038   1.0000
   8.250   0.5056   0.09785   0.08703  -0.0204   0.5799   1.0000
   8.500   0.5078   0.09979   0.08898  -0.0194   0.5694   1.0000
   8.750   0.5318   0.10341   0.09268  -0.0202   0.5608   1.0000
   9.250   0.5376   0.10701   0.09634  -0.0184   0.5366   1.0000
   9.500   0.5482   0.10985   0.09923  -0.0183   0.5278   1.0000
   9.750   0.5732   0.11348   0.10295  -0.0188   0.5167   1.0000
  10.000   0.5661   0.11469   0.10417  -0.0177   0.5048   1.0000
  10.250   0.5711   0.11737   0.10692  -0.0175   0.4964   1.0000
  10.500   0.6025   0.12198   0.11165  -0.0183   0.4858   1.0000
  10.750   0.5906   0.12269   0.11236  -0.0172   0.4741   1.0000
  11.000   0.5936   0.12541   0.11513  -0.0171   0.4659   1.0000
  11.250   0.6209   0.12987   0.11971  -0.0177   0.4559   1.0000
  11.500   0.6097   0.13087   0.12072  -0.0171   0.4448   1.0000
  11.750   0.6157   0.13393   0.12384  -0.0172   0.4364   1.0000
  12.000   0.6342   0.13763   0.12768  -0.0175   0.4266   1.0000
  12.250   0.6264   0.13938   0.12945  -0.0175   0.4175   1.0000
  12.500   0.6442   0.14359   0.13376  -0.0179   0.4091   1.0000
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