Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

EPPLER 433 AIRFOIL (e433-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 433 AIRFOIL (e433-il)
Reynolds number: 50,000
Max Cl/Cd: 29.16 at α=11.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e433-il-50000.txt
Download as CSV file: xf-e433-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 433 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4373   0.11953   0.11397  -0.0175   1.0000   0.2475
  -7.500  -0.4281   0.11615   0.11062  -0.0152   1.0000   0.2583
  -7.250  -0.4668   0.11608   0.11070  -0.0136   1.0000   0.2643
  -7.000  -0.4560   0.11276   0.10741  -0.0111   1.0000   0.2786
  -6.750  -0.4518   0.10990   0.10459  -0.0088   1.0000   0.2908
  -6.500  -0.4600   0.10764   0.10241  -0.0063   1.0000   0.3033
  -6.250  -0.5035   0.10764   0.10257  -0.0029   1.0000   0.3119
  -6.000  -0.5093   0.10539   0.10038   0.0003   1.0000   0.3277
  -5.750  -0.4841   0.10150   0.09649   0.0027   1.0000   0.3509
  -5.500  -0.4904   0.09964   0.09469   0.0062   1.0000   0.3705
  -5.250  -0.5149   0.09851   0.09367   0.0108   1.0000   0.3906
  -5.000  -0.5034   0.09598   0.09118   0.0145   1.0000   0.4210
  -3.750  -0.4622   0.05616   0.04825  -0.0388   1.0000   0.1236
  -3.500  -0.4426   0.05235   0.04447  -0.0388   1.0000   0.1170
  -3.250  -0.4116   0.04897   0.04023  -0.0399   1.0000   0.1064
  -3.000  -0.3863   0.04642   0.03727  -0.0403   1.0000   0.1049
  -2.750  -0.3604   0.04443   0.03480  -0.0404   1.0000   0.1059
  -2.500  -0.3345   0.04277   0.03264  -0.0403   1.0000   0.1074
  -2.250  -0.3094   0.04127   0.03076  -0.0398   1.0000   0.1089
  -2.000  -0.2875   0.03987   0.02940  -0.0392   1.0000   0.1159
  -1.750  -0.2648   0.03888   0.02824  -0.0380   1.0000   0.1270
  -1.500  -0.2446   0.03812   0.02753  -0.0364   1.0000   0.1440
  -1.250  -0.2233   0.03734   0.02687  -0.0349   1.0000   0.1765
  -1.000  -0.1982   0.03415   0.02693  -0.0334   1.0000   0.5753
  -0.750  -0.2301   0.03450   0.02773  -0.0165   1.0000   0.8706
  -0.500  -0.2002   0.03433   0.02703  -0.0146   1.0000   1.0000
  -0.250  -0.1868   0.03449   0.02681  -0.0138   1.0000   1.0000
   0.000  -0.1710   0.03481   0.02678  -0.0135   1.0000   1.0000
   0.250  -0.1533   0.03526   0.02692  -0.0136   1.0000   1.0000
   0.500  -0.1345   0.03583   0.02721  -0.0139   1.0000   1.0000
   0.750  -0.1149   0.03650   0.02763  -0.0145   1.0000   1.0000
   1.000  -0.0949   0.03725   0.02814  -0.0152   1.0000   1.0000
   1.250  -0.0745   0.03809   0.02877  -0.0159   1.0000   1.0000
   1.500  -0.0541   0.03900   0.02948  -0.0168   1.0000   1.0000
   1.750  -0.0338   0.03997   0.03029  -0.0176   1.0000   1.0000
   2.000  -0.0137   0.04101   0.03115  -0.0185   1.0000   1.0000
   2.250   0.0063   0.04211   0.03211  -0.0194   1.0000   1.0000
   2.500   0.0260   0.04327   0.03315  -0.0203   1.0000   1.0000
   2.750   0.0454   0.04449   0.03426  -0.0212   1.0000   1.0000
   3.000   0.0659   0.04585   0.03551  -0.0224   0.9992   1.0000
   3.250   0.1017   0.04833   0.03786  -0.0266   0.9903   1.0000
   3.500   0.1363   0.05077   0.04019  -0.0305   0.9792   1.0000
   3.750   0.1684   0.05302   0.04236  -0.0339   0.9671   1.0000
   4.000   0.1997   0.05517   0.04446  -0.0372   0.9530   1.0000
   4.250   0.2298   0.05727   0.04652  -0.0401   0.9383   1.0000
   4.500   0.2600   0.05936   0.04857  -0.0430   0.9222   1.0000
   4.750   0.2904   0.06149   0.05068  -0.0458   0.9056   1.0000
   5.000   0.3331   0.06447   0.05363  -0.0502   0.8860   1.0000
   5.250   0.3523   0.06580   0.05499  -0.0511   0.8689   1.0000
   5.500   0.4545   0.06456   0.05369  -0.0573   0.7699   1.0000
   5.750   0.4908   0.06572   0.05488  -0.0593   0.7519   1.0000
   6.000   0.5302   0.06681   0.05600  -0.0614   0.7350   1.0000
   6.250   0.5516   0.06794   0.05718  -0.0617   0.7187   1.0000
   6.500   0.5739   0.06909   0.05843  -0.0621   0.7026   1.0000
   6.750   0.5976   0.07026   0.05967  -0.0626   0.6870   1.0000
   7.000   0.6214   0.07144   0.06093  -0.0631   0.6714   1.0000
   7.250   0.6451   0.07260   0.06217  -0.0635   0.6558   1.0000
   7.500   0.6687   0.07374   0.06343  -0.0639   0.6403   1.0000
   7.750   0.6920   0.07486   0.06466  -0.0641   0.6247   1.0000
   8.000   0.7149   0.07601   0.06593  -0.0643   0.6093   1.0000
   8.250   0.7376   0.07709   0.06715  -0.0644   0.5937   1.0000
   8.500   0.7597   0.07822   0.06841  -0.0644   0.5783   1.0000
   8.750   0.7819   0.07927   0.06959  -0.0642   0.5625   1.0000
   9.000   0.8033   0.08038   0.07087  -0.0641   0.5470   1.0000
   9.250   0.8257   0.08128   0.07192  -0.0637   0.5311   1.0000
   9.500   0.8469   0.08231   0.07310  -0.0634   0.5155   1.0000
   9.750   0.8703   0.08298   0.07395  -0.0629   0.4995   1.0000
  10.000   0.8926   0.08373   0.07489  -0.0622   0.4838   1.0000
  10.250   0.9185   0.08393   0.07528  -0.0614   0.4677   1.0000
  10.500   0.9462   0.08373   0.07529  -0.0604   0.4518   1.0000
  10.750   0.9817   0.08228   0.07411  -0.0590   0.4358   1.0000
  11.000   1.0619   0.07364   0.06589  -0.0558   0.4204   1.0000
  11.250   1.3453   0.04613   0.03867  -0.0599   0.3542   1.0000
  11.500   1.3626   0.04690   0.03933  -0.0580   0.3246   1.0000
  11.750   1.3787   0.04798   0.04027  -0.0561   0.2959   1.0000
  12.000   1.3722   0.05023   0.04263  -0.0526   0.2768   1.0000
  12.250   1.3878   0.05164   0.04383  -0.0510   0.2504   1.0000
  12.500   1.3837   0.05410   0.04639  -0.0480   0.2332   1.0000
  12.750   1.3821   0.05657   0.04893  -0.0454   0.2163   1.0000
  13.000   1.3830   0.05915   0.05157  -0.0432   0.2002   1.0000
  13.250   1.3830   0.06189   0.05437  -0.0411   0.1858   1.0000
  13.500   1.3832   0.06489   0.05745  -0.0393   0.1730   1.0000
  13.750   1.3847   0.06789   0.06054  -0.0377   0.1611   1.0000
  14.000   1.0513   0.11797   0.11155  -0.0521   0.2407   1.0000
  14.250   1.0285   0.12764   0.12117  -0.0558   0.2347   1.0000
<< Back to EPPLER 433 AIRFOIL (e433-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 433 AIRFOIL (e433-il)