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E61 (5.64%) (e61-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: E61 (5.64%) (e61-il)
Reynolds number: 500,000
Max Cl/Cd: 170.99 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e61-il-500000.txt
Download as CSV file: xf-e61-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E61  (5.64%)                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.3687   0.15397   0.15170  -0.0207   1.0000   0.0070
 -11.500  -0.3660   0.15166   0.14940  -0.0207   1.0000   0.0070
 -11.250  -0.3637   0.14934   0.14711  -0.0206   1.0000   0.0071
  -8.500  -0.2178   0.10911   0.10685  -0.0544   0.9839   0.0075
  -8.250  -0.2112   0.10610   0.10385  -0.0547   0.9789   0.0077
  -8.000  -0.1975   0.10292   0.10067  -0.0573   0.9763   0.0079
  -7.750  -0.1817   0.09965   0.09740  -0.0606   0.9744   0.0082
  -7.500  -0.1641   0.09627   0.09401  -0.0645   0.9729   0.0085
  -7.250  -0.1450   0.09289   0.09064  -0.0688   0.9718   0.0090
  -7.000  -0.1416   0.09049   0.08825  -0.0687   0.9643   0.0096
  -6.750  -0.1247   0.08732   0.08508  -0.0723   0.9618   0.0100
  -4.750   0.0635   0.05667   0.05435  -0.1177   0.9277   0.0121
  -4.500   0.0937   0.05295   0.05061  -0.1231   0.9258   0.0126
  -4.250   0.1309   0.04923   0.04685  -0.1300   0.9245   0.0132
  -4.000   0.1764   0.04509   0.04265  -0.1392   0.9236   0.0142
  -3.750   0.2311   0.04034   0.03780  -0.1506   0.9231   0.0161
  -1.500   0.6839   0.01128   0.00611  -0.2100   0.9134   0.0279
  -1.250   0.7253   0.01013   0.00481  -0.2129   0.9122   0.0257
  -1.000   0.7673   0.00934   0.00394  -0.2158   0.9107   0.0254
  -0.750   0.8109   0.00878   0.00331  -0.2192   0.9092   0.0275
  -0.500   0.8392   0.00854   0.00302  -0.2192   0.9012   0.0293
  -0.250   0.8808   0.00820   0.00265  -0.2221   0.8975   0.0380
   0.000   0.9158   0.00781   0.00254  -0.2239   0.8903   0.1671
   0.250   0.9560   0.00733   0.00253  -0.2269   0.8843   0.3926
   0.500   0.9793   0.00637   0.00268  -0.2259   0.8748   1.0000
   0.750   1.0124   0.00637   0.00260  -0.2270   0.8657   1.0000
   1.000   1.0449   0.00640   0.00253  -0.2279   0.8559   1.0000
   1.250   1.0728   0.00647   0.00254  -0.2279   0.8441   1.0000
   1.500   1.1007   0.00656   0.00256  -0.2278   0.8322   1.0000
   1.750   1.1282   0.00667   0.00261  -0.2277   0.8202   1.0000
   2.000   1.1552   0.00679   0.00270  -0.2274   0.8079   1.0000
   2.250   1.1815   0.00692   0.00278  -0.2270   0.7946   1.0000
   2.500   1.2072   0.00706   0.00287  -0.2265   0.7806   1.0000
   2.750   1.2325   0.00722   0.00299  -0.2259   0.7663   1.0000
   3.000   1.2572   0.00739   0.00315  -0.2252   0.7510   1.0000
   3.250   1.2813   0.00759   0.00329  -0.2243   0.7339   1.0000
   3.500   1.3045   0.00779   0.00345  -0.2232   0.7144   1.0000
   3.750   1.3270   0.00803   0.00364  -0.2220   0.6929   1.0000
   4.000   1.3487   0.00829   0.00384  -0.2207   0.6680   1.0000
   4.250   1.3684   0.00864   0.00407  -0.2188   0.6358   1.0000
   4.500   1.3809   0.00931   0.00445  -0.2155   0.5688   1.0000
   4.750   1.3864   0.01052   0.00505  -0.2110   0.4627   1.0000
   5.000   1.3959   0.01177   0.00575  -0.2075   0.3668   1.0000
   5.250   1.4106   0.01276   0.00637  -0.2052   0.2990   1.0000
   5.500   1.4254   0.01379   0.00703  -0.2029   0.2339   1.0000
   5.750   1.4375   0.01504   0.00783  -0.2002   0.1574   1.0000
   6.000   1.4466   0.01656   0.00882  -0.1970   0.0772   1.0000
   6.250   1.4538   0.01828   0.01010  -0.1933   0.0137   1.0000
   6.500   1.4707   0.01907   0.01094  -0.1911   0.0090   1.0000
   6.750   1.4869   0.01995   0.01194  -0.1888   0.0073   1.0000
   7.000   1.4996   0.02120   0.01339  -0.1858   0.0066   1.0000
   7.250   1.5126   0.02232   0.01465  -0.1830   0.0065   1.0000
   7.500   1.5240   0.02361   0.01607  -0.1800   0.0064   1.0000
   7.750   1.5336   0.02509   0.01770  -0.1767   0.0063   1.0000
   8.000   1.5415   0.02680   0.01956  -0.1732   0.0064   1.0000
   8.250   1.5491   0.02885   0.02179  -0.1698   0.0064   1.0000
   8.500   1.5582   0.03151   0.02462  -0.1667   0.0065   1.0000
   8.750   1.5739   0.03356   0.02680  -0.1647   0.0067   1.0000
   9.000   1.5898   0.03490   0.02826  -0.1628   0.0069   1.0000
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