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EPPLER 582 AIRFOIL (e582-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 582 AIRFOIL (e582-il)
Reynolds number: 50,000
Max Cl/Cd: 7.92 at α=10.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e582-il-50000.txt
Download as CSV file: xf-e582-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 582 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
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.4099   0.11996   0.11538  -0.0123   1.0000   0.2531
  -7.000  -0.4541   0.11445   0.10941  -0.0110   1.0000   0.2584
  -6.750  -0.4808   0.11355   0.10863  -0.0089   1.0000   0.2673
  -6.500  -0.4742   0.11061   0.10571  -0.0064   1.0000   0.2797
  -6.250  -0.4829   0.10840   0.10357  -0.0038   1.0000   0.2924
  -6.000  -0.5240   0.10776   0.10306  -0.0006   1.0000   0.2993
  -5.750  -0.5289   0.10536   0.10071   0.0022   1.0000   0.3148
  -5.500  -0.5452   0.10321   0.09864   0.0046   1.0000   0.3315
  -5.250  -0.5278   0.10037   0.09583   0.0090   1.0000   0.3549
  -5.000  -0.5385   0.09844   0.09397   0.0122   1.0000   0.3767
  -3.000  -0.4276   0.05265   0.04427  -0.0333   1.0000   0.1113
  -2.750  -0.4052   0.04992   0.04131  -0.0335   1.0000   0.1072
  -2.500  -0.3799   0.04763   0.03854  -0.0336   1.0000   0.1050
  -2.250  -0.3557   0.04598   0.03646  -0.0335   1.0000   0.1056
  -2.000  -0.3312   0.04454   0.03461  -0.0332   1.0000   0.1086
  -1.750  -0.3071   0.04338   0.03305  -0.0326   1.0000   0.1106
  -1.500  -0.2846   0.04228   0.03180  -0.0320   1.0000   0.1162
  -1.250  -0.2627   0.04159   0.03094  -0.0308   1.0000   0.1286
  -1.000  -0.2435   0.04100   0.03043  -0.0290   1.0000   0.1460
  -0.750  -0.2224   0.04044   0.03001  -0.0277   1.0000   0.1792
  -0.500  -0.2036   0.03754   0.03045  -0.0246   1.0000   0.5984
  -0.250  -0.2016   0.03759   0.03048  -0.0131   1.0000   1.0000
   0.000  -0.1887   0.03790   0.03033  -0.0120   1.0000   1.0000
   0.250  -0.1745   0.03832   0.03038  -0.0113   1.0000   1.0000
   0.500  -0.1591   0.03884   0.03057  -0.0109   1.0000   1.0000
   0.750  -0.1427   0.03946   0.03091  -0.0107   1.0000   1.0000
   1.000  -0.1256   0.04017   0.03136  -0.0107   1.0000   1.0000
   1.250  -0.1081   0.04096   0.03190  -0.0109   1.0000   1.0000
   1.500  -0.0902   0.04182   0.03254  -0.0112   1.0000   1.0000
   1.750  -0.0720   0.04275   0.03327  -0.0115   1.0000   1.0000
   2.000  -0.0537   0.04375   0.03409  -0.0120   1.0000   1.0000
   2.250  -0.0354   0.04481   0.03499  -0.0125   1.0000   1.0000
   2.500  -0.0051   0.04683   0.03680  -0.0154   0.9953   1.0000
   2.750   0.0269   0.04895   0.03875  -0.0186   0.9869   1.0000
   3.000   0.0566   0.05094   0.04060  -0.0215   0.9778   1.0000
   3.250   0.0872   0.05314   0.04265  -0.0244   0.9689   1.0000
   3.500   0.1236   0.05598   0.04534  -0.0284   0.9585   1.0000
   3.750   0.1498   0.05760   0.04690  -0.0305   0.9459   1.0000
   4.000   0.1737   0.05914   0.04837  -0.0322   0.9332   1.0000
   4.250   0.1965   0.06078   0.04995  -0.0338   0.9207   1.0000
   4.500   0.2205   0.06263   0.05176  -0.0355   0.9083   1.0000
   4.750   0.2471   0.06485   0.05393  -0.0377   0.8970   1.0000
   5.000   0.2800   0.06761   0.05664  -0.0408   0.8846   1.0000
   5.250   0.3047   0.06951   0.05854  -0.0425   0.8705   1.0000
   5.500   0.3239   0.07104   0.06007  -0.0434   0.8563   1.0000
   5.750   0.3420   0.07267   0.06171  -0.0441   0.8426   1.0000
   6.000   0.3612   0.07451   0.06357  -0.0451   0.8288   1.0000
   6.250   0.3808   0.07648   0.06557  -0.0461   0.8153   1.0000
   6.500   0.4012   0.07861   0.06773  -0.0473   0.8023   1.0000
   6.750   0.4238   0.08094   0.07010  -0.0488   0.7894   1.0000
   7.000   0.4478   0.08336   0.07256  -0.0504   0.7757   1.0000
   7.250   0.4705   0.08567   0.07492  -0.0517   0.7611   1.0000
   7.500   0.4934   0.08802   0.07732  -0.0531   0.7465   1.0000
   7.750   0.5130   0.09018   0.07957  -0.0540   0.7313   1.0000
   8.000   0.5318   0.09234   0.08181  -0.0548   0.7159   1.0000
   8.250   0.5489   0.09441   0.08395  -0.0554   0.7003   1.0000
   8.500   0.5648   0.09649   0.08610  -0.0559   0.6846   1.0000
   8.750   0.5755   0.09798   0.08767  -0.0556   0.6643   1.0000
   9.000   0.6768   0.09246   0.08228  -0.0550   0.5779   1.0000
   9.250   0.6965   0.09406   0.08401  -0.0553   0.5631   1.0000
   9.500   0.7166   0.09567   0.08573  -0.0555   0.5485   1.0000
   9.750   0.7370   0.09726   0.08745  -0.0557   0.5340   1.0000
  10.000   0.7575   0.09879   0.08911  -0.0558   0.5196   1.0000
  10.250   0.7803   0.10017   0.09065  -0.0560   0.5056   1.0000
  10.500   0.8033   0.10139   0.09203  -0.0560   0.4913   1.0000
  10.750   0.8146   0.10356   0.09433  -0.0559   0.4777   1.0000
  11.000   0.8191   0.10630   0.09719  -0.0559   0.4640   1.0000
  11.250   0.8213   0.10946   0.10048  -0.0560   0.4510   1.0000
  11.500   0.8286   0.11220   0.10335  -0.0561   0.4378   1.0000
  11.750   0.8386   0.11478   0.10606  -0.0563   0.4248   1.0000
  12.000   0.8514   0.11709   0.10851  -0.0564   0.4115   1.0000
  12.250   0.8663   0.11912   0.11069  -0.0563   0.3978   1.0000
  12.500   0.8838   0.12085   0.11260  -0.0561   0.3843   1.0000
  12.750   0.9016   0.12236   0.11429  -0.0557   0.3703   1.0000
  13.000   0.9235   0.12332   0.11544  -0.0551   0.3564   1.0000
  13.250   0.8863   0.13234   0.12442  -0.0582   0.3493   1.0000
  13.500   0.8964   0.13515   0.12737  -0.0585   0.3370   1.0000
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