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

EPPLER 550 AIRFOIL (e550-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 550 AIRFOIL (e550-il)
Reynolds number: 100,000
Max Cl/Cd: 33.56 at α=12°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e550-il-100000.txt
Download as CSV file: xf-e550-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 550 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
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.4043   0.11463   0.10995  -0.0435   1.0000   0.1423
 -11.500  -0.4904   0.08336   0.07859  -0.0620   1.0000   0.0678
 -11.250  -0.5503   0.07482   0.06988  -0.0665   1.0000   0.0655
 -11.000  -0.5772   0.06961   0.06450  -0.0671   1.0000   0.0602
 -10.750  -0.6002   0.06547   0.06029  -0.0665   1.0000   0.0583
 -10.500  -0.6384   0.06174   0.05638  -0.0641   1.0000   0.0563
 -10.250  -0.7109   0.06049   0.05465  -0.0559   1.0000   0.0538
 -10.000  -0.6970   0.05677   0.05099  -0.0549   1.0000   0.0519
  -9.750  -0.7078   0.05452   0.04872  -0.0514   1.0000   0.0510
  -9.500  -0.7374   0.05426   0.04851  -0.0447   1.0000   0.0508
  -9.250  -0.7267   0.04976   0.04360  -0.0465   0.9902   0.0487
  -9.000  -0.7047   0.04410   0.03682  -0.0495   0.9740   0.0457
  -8.750  -0.6664   0.04018   0.03253  -0.0529   0.9621   0.0460
  -8.500  -0.6225   0.03758   0.02991  -0.0569   0.9515   0.0494
  -8.250  -0.5753   0.03464   0.02654  -0.0607   0.9432   0.0521
  -8.000  -0.5243   0.03161   0.02336  -0.0639   0.9340   0.0550
  -7.750  -0.4702   0.02953   0.02116  -0.0681   0.9256   0.0627
  -7.500  -0.4189   0.02743   0.01925  -0.0712   0.9141   0.0746
  -7.250  -0.3792   0.02565   0.01774  -0.0733   0.8991   0.0957
  -7.000  -0.3581   0.02398   0.01637  -0.0732   0.8810   0.1319
  -6.750  -0.3589   0.02199   0.01490  -0.0708   0.8619   0.1948
  -6.500  -0.3786   0.01953   0.01361  -0.0662   0.8441   0.3527
  -6.250  -0.3597   0.02227   0.01687  -0.0610   0.8289   0.5529
  -6.000  -0.3230   0.02550   0.01990  -0.0577   0.8155   0.5857
  -5.750  -0.2648   0.02971   0.02390  -0.0552   0.8046   0.6045
  -5.500  -0.2062   0.03280   0.02669  -0.0540   0.7948   0.6211
  -5.250  -0.1659   0.03435   0.02804  -0.0527   0.7835   0.6369
  -5.000  -0.1212   0.03559   0.02902  -0.0524   0.7740   0.6532
  -4.750  -0.0633   0.03665   0.02983  -0.0540   0.7641   0.6704
  -4.500  -0.0326   0.03691   0.02991  -0.0536   0.7552   0.6862
  -4.250   0.0146   0.03700   0.02979  -0.0555   0.7471   0.6985
  -4.000   0.0520   0.03683   0.02946  -0.0565   0.7396   0.7097
  -3.750   0.0573   0.03644   0.02900  -0.0539   0.7321   0.7201
  -3.500   0.0954   0.03619   0.02856  -0.0552   0.7260   0.7297
  -3.250   0.1070   0.03579   0.02813  -0.0535   0.7184   0.7376
  -3.000   0.1366   0.03544   0.02764  -0.0540   0.7126   0.7450
  -2.750   0.1312   0.03507   0.02724  -0.0502   0.7075   0.7494
  -2.500   0.1634   0.03468   0.02679  -0.0513   0.7012   0.7554
  -2.250   0.1846   0.03430   0.02632  -0.0509   0.6962   0.7599
  -2.000   0.1895   0.03398   0.02594  -0.0484   0.6916   0.7627
  -1.750   0.1767   0.03377   0.02578  -0.0433   0.6860   0.7650
  -1.500   0.1669   0.03339   0.02539  -0.0388   0.6814   0.7676
  -1.250   0.2049   0.03307   0.02494  -0.0405   0.6777   0.7706
  -1.000   0.2146   0.03297   0.02487  -0.0387   0.6730   0.7727
  -0.750   0.2212   0.03284   0.02476  -0.0365   0.6681   0.7751
  -0.500   0.2297   0.03255   0.02443  -0.0345   0.6639   0.7771
  -0.250   0.2405   0.03222   0.02402  -0.0330   0.6606   0.7796
   0.000   0.2351   0.03227   0.02413  -0.0297   0.6557   0.7818
   0.250   0.2397   0.03227   0.02416  -0.0279   0.6512   0.7840
   0.500   0.2622   0.03220   0.02406  -0.0275   0.6474   0.7861
   0.750   0.2881   0.03200   0.02378  -0.0278   0.6444   0.7880
   1.000   0.2916   0.03247   0.02433  -0.0257   0.6392   0.7899
   1.250   0.2982   0.03295   0.02487  -0.0240   0.6345   0.7927
   1.500   0.3181   0.03301   0.02491  -0.0240   0.6308   0.7954
   1.750   0.3477   0.03281   0.02463  -0.0253   0.6280   0.7974
   2.000   0.3471   0.03393   0.02584  -0.0239   0.6221   0.7998
   2.250   0.3502   0.03491   0.02691  -0.0217   0.6168   0.8023
   2.500   0.3759   0.03499   0.02697  -0.0218   0.6134   0.8048
   2.750   0.4114   0.03475   0.02670  -0.0231   0.6110   0.8072
   3.000   0.3635   0.03850   0.03062  -0.0179   0.6009   0.8105
   3.250   0.3957   0.03854   0.03066  -0.0193   0.5974   0.8130
   3.500   0.4396   0.03812   0.03019  -0.0218   0.5952   0.8158
   3.750   0.3501   0.04407   0.03630  -0.0135   0.5826   0.8198
   4.000   0.3907   0.04359   0.03583  -0.0146   0.5798   0.8227
   4.250   0.4488   0.04246   0.03471  -0.0173   0.5783   0.8254
   4.500   0.3753   0.04867   0.04096  -0.0129   0.5658   0.8299
   4.750   0.4206   0.04812   0.04044  -0.0145   0.5627   0.8328
   5.000   0.4768   0.04661   0.03896  -0.0158   0.5610   0.8356
   5.250   0.4139   0.05238   0.04479  -0.0123   0.5486   0.8406
   5.500   0.4662   0.05135   0.04379  -0.0138   0.5455   0.8449
   5.750   0.4326   0.05547   0.04795  -0.0121   0.5348   0.8496
   6.000   0.4713   0.05503   0.04757  -0.0124   0.5305   0.8538
   6.250   0.4618   0.05778   0.05037  -0.0117   0.5212   0.8588
   6.500   0.4919   0.05814   0.05077  -0.0125   0.5154   0.8639
   6.750   0.5445   0.05668   0.04940  -0.0129   0.5125   0.8690
   7.000   0.5167   0.06062   0.05339  -0.0117   0.4999   0.8756
   7.250   0.5658   0.05940   0.05226  -0.0122   0.4967   0.8826
   7.500   0.5451   0.06287   0.05579  -0.0111   0.4843   0.8907
   7.750   0.5923   0.06155   0.05458  -0.0113   0.4808   0.8998
   8.000   0.5771   0.06483   0.05794  -0.0107   0.4683   0.9106
   8.250   0.6252   0.06326   0.05651  -0.0107   0.4649   0.9239
   8.500   0.6183   0.06655   0.05993  -0.0116   0.4520   0.9427
  11.000   1.0945   0.04701   0.04147  -0.0198   0.4003   1.0000
  11.250   1.0547   0.05278   0.04723  -0.0188   0.3846   1.0000
  11.500   1.0265   0.05809   0.05254  -0.0186   0.3685   1.0000
  11.750   1.1266   0.04850   0.04318  -0.0173   0.3667   1.0000
  12.000   1.2591   0.03752   0.03226  -0.0185   0.3541   1.0000
  12.250   1.2630   0.03828   0.03303  -0.0169   0.3388   1.0000
  12.500   1.2628   0.03961   0.03436  -0.0155   0.3225   1.0000
  12.750   1.2616   0.04126   0.03597  -0.0142   0.3051   1.0000
  13.000   1.2623   0.04290   0.03754  -0.0131   0.2869   1.0000
  13.250   1.2646   0.04452   0.03905  -0.0121   0.2685   1.0000
  13.500   1.2696   0.04601   0.04032  -0.0112   0.2497   1.0000
  13.750   1.2632   0.04876   0.04301  -0.0105   0.2333   1.0000
  14.000   1.2576   0.05156   0.04573  -0.0100   0.2176   1.0000
  14.250   1.2530   0.05437   0.04846  -0.0096   0.2026   1.0000
  14.500   1.2493   0.05720   0.05120  -0.0093   0.1882   1.0000
  14.750   1.2470   0.05998   0.05388  -0.0092   0.1746   1.0000
  15.000   1.2461   0.06268   0.05647  -0.0091   0.1615   1.0000
  15.250   1.2370   0.06656   0.06047  -0.0096   0.1510   1.0000
  15.500   1.2340   0.06981   0.06372  -0.0099   0.1401   1.0000
  15.750   1.2344   0.07269   0.06653  -0.0101   0.1299   1.0000
  16.000   1.2402   0.07492   0.06860  -0.0100   0.1194   1.0000
  16.250   1.2311   0.07928   0.07317  -0.0111   0.1124   1.0000
  16.500   1.2344   0.08200   0.07583  -0.0114   0.1042   1.0000
  16.750   1.2320   0.08553   0.07943  -0.0122   0.0975   1.0000
  17.000   1.2328   0.08879   0.08273  -0.0128   0.0911   1.0000
  17.250   1.2316   0.09228   0.08627  -0.0138   0.0853   1.0000
  17.500   1.2337   0.09546   0.08949  -0.0144   0.0800   1.0000
  17.750   1.2254   0.10014   0.09436  -0.0161   0.0761   1.0000
  18.000   1.2381   0.10178   0.09586  -0.0160   0.0704   1.0000
  18.250   1.2234   0.10761   0.10200  -0.0185   0.0681   1.0000
  18.500   1.2125   0.11295   0.10755  -0.0208   0.0654   1.0000
  18.750   1.2297   0.11386   0.10826  -0.0204   0.0605   1.0000
  19.000   1.2092   0.12086   0.11559  -0.0240   0.0595   1.0000
  19.250   1.1876   0.12837   0.12338  -0.0280   0.0586   1.0000
<< Back to EPPLER 550 AIRFOIL (e550-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 550 AIRFOIL (e550-il)