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EPPLER 558 AIRFOIL (e558-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 558 AIRFOIL (e558-il)
Reynolds number: 50,000
Max Cl/Cd: 9.53 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e558-il-50000.txt
Download as CSV file: xf-e558-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 558 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.3037   0.12048   0.11330  -0.0285   1.0000   0.2933
 -10.000  -0.2850   0.11575   0.10854  -0.0275   1.0000   0.3062
  -9.750  -0.3086   0.11523   0.10816  -0.0277   1.0000   0.3111
  -9.500  -0.2917   0.11101   0.10391  -0.0265   1.0000   0.3271
  -9.250  -0.3872   0.09534   0.08860  -0.0399   1.0000   0.2114
  -9.000  -0.3455   0.09907   0.09221  -0.0330   1.0000   0.2595
  -8.750  -0.3490   0.09493   0.08812  -0.0329   1.0000   0.2583
  -8.500  -0.3599   0.09107   0.08435  -0.0326   1.0000   0.2575
  -8.250  -0.3773   0.08747   0.08087  -0.0319   1.0000   0.2574
  -8.000  -0.3995   0.08408   0.07763  -0.0305   1.0000   0.2571
  -7.750  -0.4272   0.08107   0.07479  -0.0282   1.0000   0.2563
  -7.500  -0.4597   0.07794   0.07185  -0.0260   1.0000   0.2554
  -7.250  -0.4937   0.07374   0.06782  -0.0255   1.0000   0.2545
  -7.000  -0.5312   0.06759   0.06178  -0.0276   1.0000   0.2537
  -6.750  -0.5686   0.05817   0.05227  -0.0339   1.0000   0.2570
  -6.500  -0.5582   0.05967   0.05380  -0.0290   1.0000   0.2705
  -6.250  -0.5600   0.05695   0.05100  -0.0290   1.0000   0.2827
  -6.000  -0.5588   0.05344   0.04730  -0.0308   1.0000   0.2970
  -5.750  -0.5513   0.05107   0.04474  -0.0317   1.0000   0.3126
  -5.500  -0.5415   0.05295   0.04673  -0.0262   1.0000   0.3229
  -5.250  -0.5324   0.05168   0.04537  -0.0256   1.0000   0.3359
  -5.000  -0.5211   0.04984   0.04336  -0.0264   1.0000   0.3506
  -4.750  -0.5068   0.04792   0.04122  -0.0279   1.0000   0.3660
  -4.500  -0.4610   0.04807   0.04122  -0.0316   0.9865   0.3857
  -4.250  -0.4177   0.04764   0.04062  -0.0355   0.9733   0.4049
  -4.000  -0.3762   0.04680   0.03958  -0.0396   0.9602   0.4231
  -3.500  -0.2977   0.04493   0.03728  -0.0474   0.9335   0.4578
  -3.250  -0.2593   0.04400   0.03611  -0.0513   0.9201   0.4744
  -3.000  -0.2271   0.04387   0.03595  -0.0520   0.9070   0.4873
  -2.750  -0.1868   0.04325   0.03518  -0.0551   0.8951   0.5022
  -2.500  -0.1498   0.04258   0.03435  -0.0579   0.8823   0.5172
  -2.250  -0.1191   0.04197   0.03356  -0.0603   0.8691   0.5311
  -2.000  -0.0850   0.04163   0.03308  -0.0623   0.8571   0.5454
  -1.750  -0.0435   0.04120   0.03258  -0.0647   0.8461   0.5598
  -1.500  -0.0221   0.04108   0.03239  -0.0648   0.8329   0.5716
  -1.250   0.0108   0.04084   0.03200  -0.0668   0.8215   0.5861
  -1.000   0.0527   0.04044   0.03149  -0.0695   0.8110   0.6025
  -0.750   0.0673   0.04074   0.03179  -0.0680   0.7984   0.6130
  -0.500   0.1041   0.04054   0.03153  -0.0697   0.7887   0.6286
  -0.250   0.1286   0.04066   0.03160  -0.0700   0.7773   0.6429
   0.000   0.1502   0.04102   0.03188  -0.0702   0.7661   0.6576
   0.250   0.1894   0.04076   0.03159  -0.0717   0.7573   0.6757
   0.500   0.1962   0.04164   0.03248  -0.0698   0.7456   0.6885
   0.750   0.2449   0.04113   0.03190  -0.0723   0.7382   0.7105
   1.000   0.2409   0.04251   0.03333  -0.0693   0.7262   0.7231
   1.250   0.2909   0.04185   0.03264  -0.0713   0.7197   0.7482
   1.500   0.2802   0.04370   0.03452  -0.0683   0.7076   0.7631
   1.750   0.3225   0.04312   0.03399  -0.0688   0.7014   0.7916
   2.000   0.3060   0.04527   0.03621  -0.0654   0.6904   0.8106
   2.250   0.3453   0.04468   0.03570  -0.0653   0.6841   0.8511
   2.500   0.3265   0.04691   0.03810  -0.0621   0.6744   0.8864
   2.750   0.4118   0.04634   0.03755  -0.0716   0.6662   1.0000
   3.000   0.4087   0.04955   0.04061  -0.0741   0.6558   1.0000
   3.250   0.4736   0.04970   0.04049  -0.0799   0.6485   1.0000
   3.500   0.4540   0.05337   0.04406  -0.0789   0.6405   1.0000
   3.750   0.4832   0.05474   0.04528  -0.0804   0.6333   1.0000
   4.000   0.4949   0.05696   0.04740  -0.0807   0.6268   1.0000
   4.250   0.4899   0.05996   0.05034  -0.0801   0.6216   1.0000
   4.500   0.5146   0.06159   0.05190  -0.0810   0.6154   1.0000
   4.750   0.5266   0.06389   0.05414  -0.0811   0.6096   1.0000
   5.000   0.5199   0.06716   0.05738  -0.0807   0.6073   1.0000
   5.250   0.5197   0.07017   0.06037  -0.0807   0.6056   1.0000
   5.500   0.5214   0.07322   0.06341  -0.0809   0.6055   1.0000
   5.750   0.5264   0.07642   0.06661  -0.0815   0.6081   1.0000
   6.000   0.5441   0.07961   0.06979  -0.0829   0.6115   1.0000
   6.250   0.4411   0.08763   0.07798  -0.0834   0.7153   1.0000
   6.500   0.4643   0.09013   0.08045  -0.0845   0.7042   1.0000
   6.750   0.4688   0.09201   0.08233  -0.0839   0.6942   1.0000
   7.000   0.5062   0.09571   0.08602  -0.0863   0.6844   1.0000
   7.250   0.5007   0.09674   0.08705  -0.0847   0.6724   1.0000
   7.500   0.5382   0.10108   0.09140  -0.0873   0.6655   1.0000
   7.750   0.5349   0.10192   0.09225  -0.0858   0.6517   1.0000
   8.000   0.5429   0.10435   0.09470  -0.0858   0.6422   1.0000
   8.250   0.5699   0.10754   0.09791  -0.0871   0.6320   1.0000
   8.500   0.5682   0.10931   0.09970  -0.0863   0.6208   1.0000
   8.750   0.6103   0.11409   0.10454  -0.0887   0.6125   1.0000
   9.000   0.5974   0.11467   0.10513  -0.0871   0.5997   1.0000
   9.250   0.6185   0.11847   0.10897  -0.0882   0.5928   1.0000
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