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EPPLER 1211 AIRFOIL (e1211-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 1211 AIRFOIL (e1211-il)
Reynolds number: 100,000
Max Cl/Cd: 33.87 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1211-il-100000.txt
Download as CSV file: xf-e1211-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1211 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.2189   0.11711   0.11199  -0.0316   1.0000   0.1739
 -10.500  -0.2137   0.11515   0.11008  -0.0315   1.0000   0.1823
 -10.250  -0.2658   0.11414   0.10920  -0.0360   1.0000   0.1851
 -10.000  -0.2036   0.10946   0.10449  -0.0312   1.0000   0.1917
  -9.750  -0.2265   0.10847   0.10361  -0.0327   1.0000   0.1998
  -9.500  -0.2433   0.10535   0.10061  -0.0335   1.0000   0.2020
  -9.250  -0.2032   0.10238   0.09765  -0.0304   1.0000   0.2069
  -9.000  -0.2073   0.10110   0.09650  -0.0290   1.0000   0.2130
  -8.500  -0.1879   0.09485   0.09038  -0.0359   0.9814   0.2232
  -8.250  -0.2082   0.09182   0.08734  -0.0506   0.9512   0.2361
  -8.000  -0.1405   0.08618   0.08163  -0.0494   0.9416   0.2397
  -7.750  -0.0943   0.08256   0.07792  -0.0549   0.9262   0.2503
  -7.500  -0.0750   0.07760   0.07286  -0.0641   0.9069   0.2585
  -7.250  -0.0203   0.07425   0.06934  -0.0687   0.8861   0.2683
  -7.000  -0.0989   0.06190   0.05679  -0.0860   0.8595   0.2374
  -6.750  -0.0760   0.05915   0.05390  -0.0869   0.8390   0.2387
  -6.500  -0.0765   0.05383   0.04841  -0.0906   0.8198   0.2309
  -6.250  -0.0980   0.04623   0.04050  -0.0966   0.8031   0.2301
  -6.000  -0.1162   0.03819   0.03165  -0.1022   0.7896   0.2354
  -5.750  -0.0920   0.03734   0.03078  -0.1012   0.7749   0.2393
  -5.500  -0.0681   0.03741   0.03088  -0.0996   0.7590   0.2435
  -5.250  -0.0500   0.03560   0.02879  -0.0998   0.7461   0.2489
  -5.000  -0.0348   0.03295   0.02547  -0.1009   0.7335   0.2553
  -4.750  -0.0120   0.03217   0.02478  -0.0999   0.7203   0.2591
  -4.500   0.0140   0.03175   0.02427  -0.0991   0.7089   0.2636
  -4.250   0.0354   0.03085   0.02320  -0.0987   0.6958   0.2689
  -4.000   0.0605   0.02958   0.02138  -0.0992   0.6861   0.2750
  -3.750   0.0824   0.02878   0.02062  -0.0983   0.6733   0.2792
  -3.500   0.1098   0.02841   0.02014  -0.0978   0.6639   0.2840
  -3.250   0.1328   0.02788   0.01952  -0.0972   0.6515   0.2893
  -3.000   0.1609   0.02723   0.01842  -0.0974   0.6429   0.2954
  -2.750   0.1837   0.02665   0.01788  -0.0967   0.6310   0.2999
  -2.500   0.2115   0.02629   0.01740  -0.0963   0.6224   0.3051
  -2.250   0.2357   0.02604   0.01709  -0.0957   0.6115   0.3108
  -2.000   0.2640   0.02570   0.01640  -0.0957   0.6028   0.3171
  -1.750   0.2886   0.02532   0.01607  -0.0951   0.5932   0.3219
  -1.500   0.3151   0.02510   0.01579  -0.0946   0.5840   0.3280
  -1.250   0.3429   0.02496   0.01547  -0.0944   0.5761   0.3350
  -1.000   0.3680   0.02477   0.01522  -0.0939   0.5664   0.3409
  -0.750   0.3966   0.02453   0.01490  -0.0937   0.5593   0.3475
  -0.500   0.4207   0.02458   0.01496  -0.0930   0.5500   0.3551
  -0.250   0.4482   0.02443   0.01468  -0.0927   0.5421   0.3628
   0.000   0.4760   0.02436   0.01459  -0.0924   0.5355   0.3706
   0.250   0.4998   0.02447   0.01473  -0.0917   0.5265   0.3799
   0.500   0.5279   0.02431   0.01453  -0.0914   0.5197   0.3895
   0.750   0.5535   0.02446   0.01468  -0.0909   0.5125   0.4007
   1.000   0.5778   0.02451   0.01483  -0.0902   0.5046   0.4117
   1.250   0.6068   0.02450   0.01470  -0.0901   0.4987   0.4275
   1.500   0.6306   0.02469   0.01504  -0.0893   0.4919   0.4433
   1.750   0.6543   0.02480   0.01529  -0.0885   0.4844   0.4625
   2.000   0.6826   0.02474   0.01523  -0.0882   0.4788   0.4898
   2.250   0.7058   0.02494   0.01565  -0.0873   0.4729   0.5239
   2.500   0.7264   0.02511   0.01616  -0.0860   0.4659   0.5757
   2.750   0.7489   0.02484   0.01628  -0.0844   0.4605   0.6766
   3.000   0.7825   0.02433   0.01625  -0.0837   0.4555   1.0000
   3.250   0.8015   0.02512   0.01703  -0.0828   0.4483   1.0000
   3.500   0.8275   0.02558   0.01735  -0.0826   0.4426   1.0000
   3.750   0.8580   0.02588   0.01740  -0.0828   0.4382   1.0000
   4.000   0.8773   0.02675   0.01829  -0.0818   0.4323   1.0000
   4.250   0.8977   0.02746   0.01902  -0.0808   0.4262   1.0000
   4.500   0.9250   0.02784   0.01927  -0.0806   0.4214   1.0000
   4.750   0.9566   0.02824   0.01944  -0.0810   0.4177   1.0000
   5.000   0.9670   0.02954   0.02094  -0.0790   0.4117   1.0000
   5.250   0.9868   0.03033   0.02175  -0.0780   0.4062   1.0000
   5.500   1.0146   0.03068   0.02198  -0.0778   0.4019   1.0000
   5.750   1.0477   0.03102   0.02212  -0.0783   0.3984   1.0000
   6.000   1.0497   0.03283   0.02421  -0.0756   0.3928   1.0000
   6.250   1.0640   0.03401   0.02546  -0.0741   0.3879   1.0000
   6.500   1.0897   0.03451   0.02590  -0.0738   0.3839   1.0000
   6.750   1.1238   0.03467   0.02589  -0.0743   0.3805   1.0000
   7.000   1.1247   0.03670   0.02812  -0.0716   0.3757   1.0000
   7.250   1.1189   0.03902   0.03065  -0.0684   0.3705   1.0000
   7.500   1.1369   0.04001   0.03164  -0.0675   0.3667   1.0000
   7.750   1.1675   0.04030   0.03186  -0.0676   0.3637   1.0000
   8.000   1.2099   0.04028   0.03164  -0.0689   0.3609   1.0000
   8.250   0.8507   0.07446   0.06689  -0.0580   0.3430   1.0000
   8.500   0.9099   0.06987   0.06222  -0.0556   0.3428   1.0000
   8.750   0.9747   0.06486   0.05712  -0.0533   0.3427   1.0000
   9.000   1.0655   0.05822   0.05033  -0.0522   0.3429   1.0000
   9.500   0.6062   0.12825   0.12124  -0.0782   0.3688   1.0000
   9.750   0.6253   0.13057   0.12353  -0.0783   0.3650   1.0000
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