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EPPLER 333 AIRFOIL (e333-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 333 AIRFOIL (e333-il)
Reynolds number: 50,000
Max Cl/Cd: 17.44 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e333-il-50000.txt
Download as CSV file: xf-e333-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 333 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3913   0.11512   0.10879  -0.0163   1.0000   0.1733
  -9.000  -0.3595   0.10893   0.10257  -0.0134   1.0000   0.1835
  -8.750  -0.3832   0.10814   0.10195  -0.0170   1.0000   0.1886
  -8.500  -0.3561   0.10284   0.09666  -0.0146   1.0000   0.2000
  -8.250  -0.3610   0.09990   0.09384  -0.0157   1.0000   0.2073
  -8.000  -0.3715   0.09781   0.09189  -0.0167   1.0000   0.2189
  -7.750  -0.3568   0.09387   0.08801  -0.0152   1.0000   0.2322
  -7.500  -0.3520   0.09064   0.08488  -0.0144   1.0000   0.2457
  -7.250  -0.3516   0.08780   0.08216  -0.0136   1.0000   0.2607
  -7.000  -0.3566   0.08521   0.07972  -0.0126   1.0000   0.2776
  -6.750  -0.3675   0.08301   0.07769  -0.0110   1.0000   0.2946
  -6.500  -0.3574   0.08009   0.07487  -0.0078   1.0000   0.3236
  -6.250  -0.3636   0.07810   0.07307  -0.0041   1.0000   0.3564
  -6.000  -0.3206   0.07478   0.06974   0.0020   1.0000   0.4297
  -5.750  -0.2666   0.07154   0.06647   0.0085   1.0000   0.5361
  -5.500  -0.1659   0.06663   0.06139   0.0122   1.0000   0.7243
  -4.750  -0.0965   0.05696   0.05205   0.0083   1.0000   0.8061
  -4.500  -0.1631   0.05824   0.05377   0.0163   1.0000   0.7474
  -4.250  -0.2067   0.06036   0.05616   0.0248   1.0000   0.7584
  -4.000  -0.2609   0.05945   0.05541   0.0220   0.9772   0.6830
  -3.750  -0.3138   0.05612   0.05216   0.0151   0.9430   0.6193
  -3.500  -0.2448   0.04601   0.03875  -0.0344   0.9116   0.1889
  -3.250  -0.1885   0.04173   0.03372  -0.0379   0.8993   0.1395
  -3.000  -0.1323   0.03828   0.02956  -0.0414   0.8880   0.1202
  -2.750  -0.0891   0.03629   0.02688  -0.0427   0.8738   0.1135
  -2.500  -0.0476   0.03421   0.02450  -0.0443   0.8597   0.1155
  -2.250  -0.0047   0.03246   0.02244  -0.0457   0.8460   0.1176
  -2.000   0.2739   0.02276   0.01492  -0.0803   0.8443   1.0000
  -1.750   0.2990   0.02300   0.01468  -0.0795   0.8269   1.0000
  -1.500   0.3197   0.02342   0.01477  -0.0784   0.8095   1.0000
  -1.250   0.3415   0.02383   0.01488  -0.0774   0.7941   1.0000
  -1.000   0.3631   0.02426   0.01504  -0.0764   0.7799   1.0000
  -0.750   0.3840   0.02472   0.01524  -0.0752   0.7664   1.0000
  -0.500   0.4026   0.02532   0.01567  -0.0739   0.7530   1.0000
  -0.250   0.4207   0.02601   0.01619  -0.0727   0.7406   1.0000
   0.000   0.4392   0.02670   0.01672  -0.0714   0.7290   1.0000
   0.250   0.4605   0.02720   0.01703  -0.0701   0.7194   1.0000
   0.500   0.4766   0.02809   0.01784  -0.0689   0.7081   1.0000
   0.750   0.4924   0.02903   0.01869  -0.0676   0.6979   1.0000
   1.000   0.5147   0.02951   0.01901  -0.0663   0.6896   1.0000
   1.250   0.5258   0.03082   0.02030  -0.0649   0.6790   1.0000
   1.500   0.5416   0.03182   0.02123  -0.0635   0.6705   1.0000
   1.750   0.5576   0.03279   0.02214  -0.0620   0.6616   1.0000
   2.000   0.5666   0.03429   0.02362  -0.0604   0.6531   1.0000
   2.250   0.5849   0.03515   0.02441  -0.0590   0.6455   1.0000
   2.500   0.5867   0.03709   0.02637  -0.0568   0.6369   1.0000
   2.750   0.6054   0.03797   0.02719  -0.0555   0.6300   1.0000
   3.000   0.5984   0.04042   0.02966  -0.0527   0.6218   1.0000
   3.250   0.6154   0.04143   0.03065  -0.0512   0.6150   1.0000
   3.500   0.6008   0.04428   0.03350  -0.0479   0.6080   1.0000
   3.750   0.6025   0.04617   0.03538  -0.0456   0.6013   1.0000
   4.000   0.6028   0.04821   0.03740  -0.0432   0.5953   1.0000
   4.250   0.5680   0.05179   0.04094  -0.0386   0.5905   1.0000
   4.500   0.5967   0.05266   0.04181  -0.0384   0.5838   1.0000
   4.750   0.5661   0.05594   0.04502  -0.0339   0.5806   1.0000
   5.000   0.5401   0.05909   0.04812  -0.0305   0.5786   1.0000
   5.250   0.5286   0.06190   0.05091  -0.0286   0.5773   1.0000
   5.500   0.5222   0.06463   0.05362  -0.0273   0.5765   1.0000
   5.750   0.5200   0.06748   0.05646  -0.0267   0.5787   1.0000
   6.000   0.5240   0.07050   0.05951  -0.0269   0.5839   1.0000
   6.250   0.4216   0.07792   0.06699  -0.0280   0.7067   1.0000
   6.500   0.4282   0.07955   0.06861  -0.0269   0.6940   1.0000
   6.750   0.4401   0.08160   0.07066  -0.0265   0.6804   1.0000
   7.000   0.4530   0.08382   0.07288  -0.0263   0.6680   1.0000
   7.250   0.4780   0.08697   0.07606  -0.0274   0.6578   1.0000
   7.500   0.4981   0.08957   0.07871  -0.0278   0.6439   1.0000
   7.750   0.5013   0.09107   0.08023  -0.0266   0.6302   1.0000
   8.000   0.5047   0.09291   0.08208  -0.0256   0.6179   1.0000
   8.250   0.5138   0.09525   0.08444  -0.0253   0.6062   1.0000
   8.500   0.5300   0.09808   0.08732  -0.0256   0.5949   1.0000
   8.750   0.5561   0.10149   0.09084  -0.0265   0.5831   1.0000
   9.000   0.5551   0.10302   0.09238  -0.0254   0.5699   1.0000
   9.250   0.5582   0.10513   0.09453  -0.0248   0.5578   1.0000
   9.500   0.5679   0.10793   0.09738  -0.0249   0.5478   1.0000
   9.750   0.5985   0.11215   0.10169  -0.0261   0.5370   1.0000
  10.000   0.5980   0.11375   0.10334  -0.0253   0.5237   1.0000
  10.250   0.5954   0.11575   0.10538  -0.0248   0.5125   1.0000
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