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EPPLER 637 AIRFOIL (e637-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 637 AIRFOIL (e637-il)
Reynolds number: 50,000
Max Cl/Cd: 24.93 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e637-il-50000-n5.txt
Download as CSV file: xf-e637-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 637 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3572   0.10100   0.09495  -0.0152   1.0000   0.1177
  -8.250  -0.3641   0.09825   0.09233  -0.0181   1.0000   0.1226
  -8.000  -0.3940   0.09651   0.09075  -0.0234   1.0000   0.1248
  -7.750  -0.3682   0.09116   0.08548  -0.0214   1.0000   0.1283
  -7.500  -0.3617   0.08789   0.08230  -0.0213   1.0000   0.1320
  -7.250  -0.3671   0.08501   0.07954  -0.0220   1.0000   0.1359
  -7.000  -0.3897   0.08270   0.07732  -0.0248   1.0000   0.1426
  -6.750  -0.3762   0.07887   0.07362  -0.0236   1.0000   0.1454
  -6.500  -0.3739   0.07601   0.07089  -0.0231   1.0000   0.1486
  -6.250  -0.3632   0.06682   0.06099  -0.0324   0.9494   0.0667
  -6.000  -0.3364   0.06094   0.05469  -0.0362   0.9116   0.0546
  -5.750  -0.3075   0.05680   0.05038  -0.0388   0.8811   0.0518
  -5.500  -0.2871   0.05312   0.04637  -0.0397   0.8529   0.0498
  -5.250  -0.2706   0.04989   0.04272  -0.0392   0.8279   0.0480
  -5.000  -0.2551   0.04682   0.03913  -0.0378   0.8064   0.0461
  -4.750  -0.2386   0.04413   0.03559  -0.0355   0.7879   0.0441
  -4.500  -0.2206   0.04182   0.03292  -0.0339   0.7698   0.0439
  -4.250  -0.2022   0.03965   0.03032  -0.0322   0.7531   0.0444
  -4.000  -0.1828   0.03779   0.02829  -0.0311   0.7369   0.0460
  -3.750  -0.1616   0.03609   0.02622  -0.0297   0.7224   0.0473
  -3.500  -0.1387   0.03440   0.02413  -0.0284   0.7081   0.0484
  -3.250  -0.1138   0.03271   0.02205  -0.0272   0.6945   0.0489
  -3.000  -0.0872   0.03118   0.02012  -0.0263   0.6815   0.0497
  -2.750  -0.0584   0.02978   0.01835  -0.0257   0.6697   0.0509
  -2.500  -0.0275   0.02861   0.01678  -0.0253   0.6585   0.0538
  -2.250   0.0044   0.02748   0.01561  -0.0257   0.6461   0.0591
  -2.000   0.0380   0.02657   0.01441  -0.0260   0.6350   0.0642
  -1.750   0.0666   0.02572   0.01342  -0.0257   0.6255   0.0719
  -1.250   0.2562   0.02124   0.01137  -0.0475   0.5984   1.0000
  -1.000   0.2787   0.02135   0.01110  -0.0465   0.5901   1.0000
  -0.750   0.3010   0.02149   0.01096  -0.0457   0.5809   1.0000
  -0.500   0.3233   0.02166   0.01085  -0.0448   0.5731   1.0000
  -0.250   0.3457   0.02185   0.01083  -0.0440   0.5647   1.0000
   0.000   0.3684   0.02206   0.01080  -0.0432   0.5576   1.0000
   0.250   0.3909   0.02233   0.01090  -0.0425   0.5497   1.0000
   0.500   0.4135   0.02254   0.01087  -0.0415   0.5437   1.0000
   0.750   0.4356   0.02290   0.01114  -0.0409   0.5356   1.0000
   1.000   0.4580   0.02315   0.01121  -0.0399   0.5297   1.0000
   1.250   0.4798   0.02355   0.01153  -0.0392   0.5228   1.0000
   1.500   0.5017   0.02390   0.01176  -0.0383   0.5165   1.0000
   1.750   0.5241   0.02420   0.01191  -0.0373   0.5114   1.0000
   2.000   0.5449   0.02474   0.01244  -0.0366   0.5042   1.0000
   2.250   0.5669   0.02510   0.01271  -0.0356   0.4990   1.0000
   2.500   0.5883   0.02557   0.01311  -0.0347   0.4936   1.0000
   2.750   0.6086   0.02615   0.01370  -0.0338   0.4872   1.0000
   3.000   0.6305   0.02653   0.01400  -0.0328   0.4825   1.0000
   3.250   0.6508   0.02714   0.01461  -0.0319   0.4772   1.0000
   3.500   0.6702   0.02782   0.01533  -0.0309   0.4712   1.0000
   3.750   0.6918   0.02826   0.01571  -0.0299   0.4668   1.0000
   4.000   0.7114   0.02893   0.01640  -0.0289   0.4618   1.0000
   4.250   0.7289   0.02980   0.01737  -0.0279   0.4560   1.0000
   4.500   0.7499   0.03031   0.01785  -0.0268   0.4515   1.0000
   4.750   0.7701   0.03093   0.01847  -0.0258   0.4473   1.0000
   5.000   0.7839   0.03213   0.01982  -0.0245   0.4410   1.0000
   5.250   0.8034   0.03278   0.02052  -0.0234   0.4366   1.0000
   5.500   0.8264   0.03315   0.02085  -0.0225   0.4333   1.0000
   5.750   0.8337   0.03489   0.02280  -0.0209   0.4265   1.0000
   6.000   0.8507   0.03575   0.02372  -0.0196   0.4217   1.0000
   6.250   0.8732   0.03616   0.02416  -0.0187   0.4183   1.0000
   6.500   0.8759   0.03823   0.02643  -0.0168   0.4117   1.0000
   6.750   0.8886   0.03943   0.02772  -0.0154   0.4067   1.0000
   7.000   0.9111   0.03982   0.02813  -0.0144   0.4032   1.0000
   7.250   0.9041   0.04261   0.03111  -0.0121   0.3964   1.0000
   7.500   0.9121   0.04412   0.03274  -0.0104   0.3911   1.0000
   7.750   0.9358   0.04440   0.03307  -0.0094   0.3878   1.0000
   8.000   0.9014   0.04927   0.03809  -0.0064   0.3787   1.0000
   8.250   0.9145   0.05035   0.03924  -0.0050   0.3744   1.0000
   8.500   0.9454   0.05001   0.03898  -0.0042   0.3716   1.0000
   9.000   0.8299   0.06549   0.05443  -0.0021   0.3486   1.0000
   9.250   0.8418   0.06685   0.05587  -0.0013   0.3437   1.0000
   9.500   0.8532   0.06834   0.05742  -0.0005   0.3392   1.0000
  10.000   0.8551   0.07370   0.06292   0.0001   0.3264   1.0000
  10.250   0.8230   0.08040   0.06960  -0.0014   0.3156   1.0000
  10.500   0.8463   0.08050   0.06985  -0.0003   0.3120   1.0000
  10.750   0.8225   0.08632   0.07568  -0.0016   0.3021   1.0000
  11.000   0.8432   0.08674   0.07621  -0.0006   0.2978   1.0000
  11.500   0.8434   0.09268   0.08230  -0.0010   0.2838   1.0000
  11.750   0.8297   0.09755   0.08721  -0.0022   0.2750   1.0000
  12.000   0.8455   0.09850   0.08831  -0.0015   0.2698   1.0000
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