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

EPPLER 638 AIRFOIL (e638-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 638 AIRFOIL (e638-il)
Reynolds number: 100,000
Max Cl/Cd: 32.15 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e638-il-100000.txt
Download as CSV file: xf-e638-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 638 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3194   0.10543   0.10105  -0.0269   1.0000   0.0686
  -9.000  -0.3287   0.10318   0.09890  -0.0311   1.0000   0.0690
  -8.750  -0.3417   0.10064   0.09646  -0.0350   1.0000   0.0692
  -8.500  -0.3032   0.09378   0.08960  -0.0277   1.0000   0.0722
  -8.250  -0.2953   0.09081   0.08670  -0.0272   1.0000   0.0764
  -8.000  -0.2970   0.08792   0.08392  -0.0286   1.0000   0.0788
  -7.750  -0.3082   0.08527   0.08141  -0.0311   1.0000   0.0811
  -7.500  -0.3234   0.08316   0.07944  -0.0323   1.0000   0.0821
  -7.250  -0.3414   0.08211   0.07844  -0.0368   0.9877   0.0833
  -7.000  -0.3091   0.07482   0.07125  -0.0405   0.9655   0.0861
  -6.750  -0.2717   0.06993   0.06631  -0.0458   0.9375   0.0928
  -6.500  -0.2496   0.06483   0.06077  -0.0564   0.9041   0.1001
  -6.250  -0.2203   0.06046   0.05640  -0.0571   0.8759   0.1056
  -6.000  -0.2161   0.05906   0.05437  -0.0590   0.8472   0.1151
  -5.750  -0.1980   0.05463   0.05007  -0.0578   0.8242   0.1190
  -5.500  -0.1906   0.05341   0.04834  -0.0573   0.8041   0.1310
  -5.250  -0.1741   0.04973   0.04478  -0.0560   0.7850   0.1362
  -5.000  -0.1628   0.04756   0.04234  -0.0550   0.7683   0.1492
  -4.750  -0.1499   0.04535   0.03995  -0.0536   0.7531   0.1650
  -4.500  -0.1357   0.04321   0.03772  -0.0521   0.7387   0.1837
  -4.250  -0.1235   0.04138   0.03575  -0.0503   0.7258   0.2130
  -4.000  -0.1105   0.03942   0.03376  -0.0481   0.7135   0.2464
  -3.750  -0.0993   0.03801   0.03224  -0.0457   0.7008   0.2893
  -3.500  -0.0851   0.03573   0.03006  -0.0431   0.6894   0.3238
  -3.250  -0.0189   0.03211   0.02399  -0.0450   0.6821   0.1134
  -3.000   0.0077   0.02912   0.02065  -0.0437   0.6712   0.0939
  -2.750   0.0342   0.02729   0.01830  -0.0422   0.6616   0.0872
  -2.500   0.0623   0.02629   0.01681  -0.0408   0.6522   0.0827
  -2.250   0.0894   0.02487   0.01519  -0.0401   0.6421   0.0818
  -2.000   0.1174   0.02390   0.01390  -0.0393   0.6343   0.0845
  -1.750   0.1449   0.02270   0.01272  -0.0390   0.6246   0.0874
  -1.500   0.1723   0.02191   0.01186  -0.0384   0.6168   0.0907
  -1.250   0.1982   0.02141   0.01130  -0.0376   0.6081   0.0978
  -1.000   0.3766   0.01734   0.00971  -0.0635   0.5931   1.0000
  -0.750   0.3993   0.01751   0.00958  -0.0626   0.5866   1.0000
  -0.500   0.4214   0.01771   0.00962  -0.0617   0.5785   1.0000
  -0.250   0.4444   0.01790   0.00954  -0.0608   0.5730   1.0000
   0.000   0.4660   0.01821   0.00979  -0.0600   0.5654   1.0000
   0.250   0.4887   0.01843   0.00983  -0.0591   0.5597   1.0000
   0.500   0.5106   0.01877   0.01008  -0.0582   0.5534   1.0000
   0.750   0.5326   0.01907   0.01028  -0.0573   0.5470   1.0000
   1.000   0.5559   0.01933   0.01034  -0.0564   0.5425   1.0000
   1.250   0.5765   0.01982   0.01087  -0.0556   0.5360   1.0000
   1.500   0.5987   0.02015   0.01111  -0.0547   0.5306   1.0000
   1.750   0.6220   0.02047   0.01127  -0.0538   0.5264   1.0000
   2.000   0.6413   0.02106   0.01194  -0.0528   0.5199   1.0000
   2.250   0.6635   0.02144   0.01226  -0.0519   0.5151   1.0000
   2.500   0.6871   0.02179   0.01247  -0.0511   0.5115   1.0000
   2.750   0.7046   0.02257   0.01339  -0.0499   0.5052   1.0000
   3.000   0.7263   0.02300   0.01377  -0.0489   0.5004   1.0000
   3.250   0.7500   0.02333   0.01401  -0.0481   0.4967   1.0000
   3.500   0.7661   0.02429   0.01511  -0.0468   0.4912   1.0000
   3.750   0.7858   0.02493   0.01577  -0.0458   0.4864   1.0000
   4.000   0.8094   0.02527   0.01603  -0.0449   0.4826   1.0000
   4.250   0.8266   0.02616   0.01702  -0.0437   0.4778   1.0000
   4.500   0.8421   0.02717   0.01814  -0.0423   0.4726   1.0000
   4.750   0.8641   0.02767   0.01862  -0.0414   0.4688   1.0000
   5.000   0.8898   0.02796   0.01882  -0.0408   0.4657   1.0000
   5.250   0.8946   0.02977   0.02090  -0.0385   0.4591   1.0000
   5.500   0.9134   0.03053   0.02171  -0.0374   0.4548   1.0000
   5.750   0.9394   0.03078   0.02191  -0.0368   0.4518   1.0000
   6.000   0.9439   0.03271   0.02404  -0.0346   0.4461   1.0000
   6.250   0.9546   0.03408   0.02552  -0.0329   0.4407   1.0000
   6.500   0.9801   0.03432   0.02577  -0.0322   0.4374   1.0000
   6.750   0.9961   0.03541   0.02692  -0.0309   0.4334   1.0000
   7.000   0.9816   0.03872   0.03048  -0.0275   0.4259   1.0000
   7.250   1.0108   0.03859   0.03035  -0.0271   0.4226   1.0000
   7.500   1.0508   0.03784   0.02954  -0.0275   0.4203   1.0000
   7.750   0.9781   0.04569   0.03774  -0.0209   0.4093   1.0000
   8.000   1.0267   0.04393   0.03599  -0.0213   0.4070   1.0000
   8.250   1.0751   0.04242   0.03445  -0.0220   0.4046   1.0000
   8.500   0.9058   0.05925   0.05140  -0.0144   0.3893   1.0000
   8.750   1.0315   0.05004   0.04233  -0.0155   0.3908   1.0000
  12.250   0.8249   0.11439   0.10717  -0.0180   0.2802   1.0000
  12.500   0.8614   0.11251   0.10541  -0.0160   0.2750   1.0000
<< Back to EPPLER 638 AIRFOIL (e638-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 638 AIRFOIL (e638-il)