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EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER EA 6(-1)-012 AIRFOIL (ea61012-il)
Reynolds number: 50,000
Max Cl/Cd: 25.62 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ea61012-il-50000.txt
Download as CSV file: xf-ea61012-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER EA 6(-1)-012 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6871   0.08506   0.07754  -0.0134   1.0000   0.2131
  -9.250  -0.7071   0.07654   0.06901  -0.0173   1.0000   0.1993
  -9.000  -0.7409   0.06972   0.06211  -0.0191   1.0000   0.1910
  -8.750  -0.8066   0.06145   0.05321  -0.0197   1.0000   0.1798
  -8.500  -0.8032   0.05725   0.04880  -0.0192   1.0000   0.1790
  -8.250  -0.8002   0.05337   0.04461  -0.0184   1.0000   0.1787
  -8.000  -0.7943   0.04977   0.04066  -0.0174   1.0000   0.1786
  -7.750  -0.7843   0.04647   0.03699  -0.0164   1.0000   0.1786
  -7.500  -0.7708   0.04345   0.03362  -0.0154   1.0000   0.1787
  -7.250  -0.7565   0.04085   0.03057  -0.0142   1.0000   0.1799
  -7.000  -0.7365   0.03836   0.02805  -0.0136   1.0000   0.1834
  -6.750  -0.7158   0.03635   0.02594  -0.0128   1.0000   0.1873
  -6.500  -0.6956   0.03439   0.02374  -0.0118   1.0000   0.1907
  -6.250  -0.6754   0.03261   0.02164  -0.0107   1.0000   0.1943
  -6.000  -0.6542   0.03093   0.01994  -0.0097   1.0000   0.2001
  -5.750  -0.6338   0.02959   0.01858  -0.0085   1.0000   0.2079
  -5.500  -0.6126   0.02824   0.01719  -0.0072   1.0000   0.2157
  -5.250  -0.5913   0.02705   0.01604  -0.0059   1.0000   0.2256
  -5.000  -0.5716   0.02595   0.01509  -0.0043   1.0000   0.2392
  -4.750  -0.5549   0.02492   0.01420  -0.0023   1.0000   0.2592
  -4.500  -0.5419   0.02375   0.01339   0.0002   1.0000   0.2877
  -4.250  -0.5357   0.02217   0.01247   0.0036   1.0000   0.3501
  -4.000  -0.5435   0.02121   0.01351   0.0127   1.0000   0.5819
  -3.750  -0.5422   0.02286   0.01524   0.0215   1.0000   0.6841
  -3.500  -0.5336   0.02414   0.01641   0.0283   1.0000   0.7331
  -3.250  -0.5250   0.02486   0.01701   0.0341   1.0000   0.7727
  -3.000  -0.5023   0.02578   0.01780   0.0384   1.0000   0.8089
  -2.750  -0.2099   0.02956   0.02072   0.0073   1.0000   0.9147
  -2.500  -0.1312   0.02848   0.01940  -0.0023   1.0000   0.9454
  -2.250  -0.0787   0.02754   0.01833  -0.0085   1.0000   0.9684
  -2.000  -0.0191   0.02629   0.01697  -0.0163   1.0000   0.9869
  -1.750   0.0244   0.02542   0.01603  -0.0216   1.0000   1.0000
  -1.500   0.0193   0.02558   0.01618  -0.0184   1.0000   1.0000
  -1.250   0.0143   0.02572   0.01632  -0.0151   1.0000   1.0000
  -1.000   0.0102   0.02584   0.01642  -0.0119   1.0000   1.0000
  -0.750   0.0070   0.02592   0.01649  -0.0088   1.0000   1.0000
  -0.500   0.0043   0.02597   0.01654  -0.0059   1.0000   1.0000
  -0.250   0.0021   0.02600   0.01657  -0.0029   1.0000   1.0000
   0.000   0.0000   0.02601   0.01657   0.0000   1.0000   1.0000
   0.250  -0.0021   0.02600   0.01656   0.0029   1.0000   1.0000
   0.500  -0.0043   0.02597   0.01653   0.0059   1.0000   1.0000
   0.750  -0.0070   0.02591   0.01649   0.0088   1.0000   1.0000
   1.000  -0.0102   0.02583   0.01641   0.0119   1.0000   1.0000
   1.250  -0.0143   0.02572   0.01631   0.0151   1.0000   1.0000
   1.500  -0.0193   0.02557   0.01617   0.0184   1.0000   1.0000
   1.750  -0.0244   0.02541   0.01602   0.0216   1.0000   1.0000
   2.000   0.0189   0.02627   0.01695   0.0164   0.9869   1.0000
   2.250   0.0781   0.02751   0.01831   0.0086   0.9685   1.0000
   2.500   0.1311   0.02846   0.01939   0.0023   0.9455   1.0000
   2.750   0.2105   0.02955   0.02071  -0.0074   0.9144   1.0000
   3.000   0.5022   0.02578   0.01780  -0.0384   0.8090   1.0000
   3.250   0.5250   0.02486   0.01701  -0.0340   0.7727   1.0000
   3.500   0.5335   0.02413   0.01640  -0.0282   0.7330   1.0000
   3.750   0.5422   0.02287   0.01524  -0.0216   0.6843   1.0000
   4.000   0.5436   0.02122   0.01352  -0.0128   0.5833   1.0000
   4.250   0.5356   0.02217   0.01247  -0.0036   0.3501   1.0000
   4.500   0.5419   0.02375   0.01339  -0.0002   0.2878   1.0000
   4.750   0.5549   0.02491   0.01419   0.0023   0.2593   1.0000
   5.000   0.5715   0.02595   0.01509   0.0043   0.2392   1.0000
   5.250   0.5912   0.02705   0.01604   0.0059   0.2256   1.0000
   5.500   0.6126   0.02824   0.01719   0.0072   0.2156   1.0000
   5.750   0.6337   0.02959   0.01858   0.0085   0.2079   1.0000
   6.000   0.6542   0.03093   0.01994   0.0097   0.2002   1.0000
   6.250   0.6754   0.03261   0.02164   0.0107   0.1943   1.0000
   6.500   0.6956   0.03439   0.02374   0.0118   0.1907   1.0000
   6.750   0.7158   0.03635   0.02594   0.0128   0.1874   1.0000
   7.000   0.7365   0.03836   0.02805   0.0136   0.1834   1.0000
   7.250   0.7566   0.04085   0.03057   0.0142   0.1799   1.0000
   7.500   0.7708   0.04346   0.03363   0.0154   0.1787   1.0000
   7.750   0.7844   0.04646   0.03699   0.0164   0.1786   1.0000
   8.000   0.7942   0.04979   0.04068   0.0174   0.1786   1.0000
   8.250   0.8002   0.05337   0.04462   0.0184   0.1787   1.0000
   8.500   0.8031   0.05727   0.04882   0.0192   0.1789   1.0000
   8.750   0.8065   0.06146   0.05322   0.0197   0.1798   1.0000
   9.000   0.7410   0.06974   0.06213   0.0190   0.1909   1.0000
   9.250   0.7050   0.07672   0.06919   0.0171   0.1994   1.0000
   9.500   0.4267   0.10811   0.10085  -0.0112   0.3935   1.0000
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