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AG13 (ag13-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: AG13 (ag13-il)
Reynolds number: 1,000,000
Max Cl/Cd: 92.95 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag13-il-1000000.txt
Download as CSV file: xf-ag13-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG13                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5682   0.09690   0.09535   0.0184   1.0000   0.0067
  -8.000  -0.5638   0.09318   0.09164   0.0163   1.0000   0.0068
  -7.750  -0.5595   0.08940   0.08787   0.0140   1.0000   0.0069
  -7.500  -0.5549   0.08550   0.08399   0.0112   1.0000   0.0071
  -7.250  -0.5459   0.08103   0.07954   0.0066   1.0000   0.0072
  -7.000  -0.5332   0.07604   0.07456   0.0006   1.0000   0.0074
  -6.750  -0.5171   0.07054   0.06905  -0.0065   1.0000   0.0077
  -6.500  -0.4972   0.06441   0.06289  -0.0146   1.0000   0.0080
  -6.250  -0.4733   0.05750   0.05592  -0.0233   1.0000   0.0084
  -6.000  -0.4454   0.05003   0.04834  -0.0315   1.0000   0.0089
  -5.750  -0.4156   0.04465   0.04284  -0.0360   1.0000   0.0096
  -5.250  -0.3619   0.01712   0.01342  -0.0477   1.0000   0.0059
  -5.000  -0.3353   0.01361   0.00940  -0.0480   1.0000   0.0065
  -4.750  -0.3087   0.01307   0.00880  -0.0479   1.0000   0.0072
  -4.500  -0.2818   0.01219   0.00779  -0.0477   1.0000   0.0077
  -4.250  -0.2549   0.01115   0.00659  -0.0474   1.0000   0.0081
  -4.000  -0.2280   0.01015   0.00546  -0.0471   1.0000   0.0083
  -3.750  -0.2014   0.00968   0.00491  -0.0467   1.0000   0.0090
  -3.500  -0.1748   0.00906   0.00421  -0.0463   1.0000   0.0094
  -3.250  -0.1480   0.00829   0.00334  -0.0459   1.0000   0.0095
  -3.000  -0.1211   0.00768   0.00264  -0.0456   1.0000   0.0099
  -2.750  -0.0936   0.00700   0.00184  -0.0453   1.0000   0.0141
  -2.500  -0.0592   0.00663   0.00146  -0.0466   0.9970   0.0255
  -2.250  -0.0225   0.00627   0.00125  -0.0485   0.9918   0.0623
  -2.000   0.0136   0.00595   0.00110  -0.0504   0.9827   0.1085
  -1.750   0.0480   0.00562   0.00098  -0.0518   0.9679   0.1747
  -1.500   0.0782   0.00534   0.00087  -0.0522   0.9427   0.2437
  -1.250   0.1040   0.00515   0.00080  -0.0515   0.9090   0.3119
  -1.000   0.1290   0.00504   0.00075  -0.0507   0.8701   0.3789
  -0.750   0.1547   0.00498   0.00071  -0.0500   0.8290   0.4453
  -0.500   0.1810   0.00490   0.00069  -0.0496   0.7886   0.5224
   0.000   0.2332   0.00455   0.00070  -0.0487   0.7110   0.7469
   0.250   0.2573   0.00409   0.00068  -0.0474   0.6740   1.0000
   0.500   0.2848   0.00427   0.00069  -0.0472   0.6367   1.0000
   0.750   0.3124   0.00444   0.00070  -0.0471   0.6003   1.0000
   1.000   0.3401   0.00461   0.00072  -0.0469   0.5656   1.0000
   1.250   0.3677   0.00479   0.00075  -0.0468   0.5316   1.0000
   1.500   0.3954   0.00497   0.00079  -0.0467   0.4987   1.0000
   1.750   0.4230   0.00516   0.00085  -0.0466   0.4644   1.0000
   2.000   0.4506   0.00536   0.00092  -0.0465   0.4319   1.0000
   2.250   0.4782   0.00555   0.00099  -0.0464   0.4002   1.0000
   2.500   0.5057   0.00577   0.00107  -0.0463   0.3681   1.0000
   2.750   0.5332   0.00600   0.00117  -0.0462   0.3359   1.0000
   3.000   0.5607   0.00622   0.00130  -0.0461   0.3066   1.0000
   3.250   0.5880   0.00646   0.00143  -0.0460   0.2761   1.0000
   3.500   0.6154   0.00670   0.00156  -0.0459   0.2482   1.0000
   3.750   0.6427   0.00695   0.00172  -0.0458   0.2235   1.0000
   4.000   0.6699   0.00723   0.00190  -0.0457   0.1956   1.0000
   4.250   0.6971   0.00750   0.00209  -0.0456   0.1729   1.0000
   4.500   0.7241   0.00781   0.00229  -0.0455   0.1478   1.0000
   4.750   0.7511   0.00809   0.00251  -0.0454   0.1269   1.0000
   5.000   0.7780   0.00842   0.00275  -0.0452   0.1058   1.0000
   5.250   0.8048   0.00874   0.00303  -0.0451   0.0890   1.0000
   5.500   0.8315   0.00910   0.00331  -0.0449   0.0722   1.0000
   5.750   0.8581   0.00947   0.00362  -0.0447   0.0568   1.0000
   6.000   0.8846   0.00984   0.00395  -0.0445   0.0447   1.0000
   6.250   0.9109   0.01024   0.00431  -0.0443   0.0336   1.0000
   6.500   0.9371   0.01068   0.00471  -0.0441   0.0241   1.0000
   6.750   0.9633   0.01110   0.00514  -0.0439   0.0182   1.0000
   7.000   0.9892   0.01161   0.00568  -0.0436   0.0128   1.0000
   7.250   1.0142   0.01243   0.00656  -0.0431   0.0085   1.0000
   7.500   1.0399   0.01294   0.00713  -0.0427   0.0071   1.0000
   7.750   1.0645   0.01374   0.00800  -0.0423   0.0056   1.0000
   8.000   1.0875   0.01498   0.00944  -0.0416   0.0049   1.0000
   8.250   1.1118   0.01572   0.01029  -0.0411   0.0047   1.0000
   8.500   1.1354   0.01664   0.01136  -0.0405   0.0044   1.0000
   8.750   1.1584   0.01763   0.01249  -0.0399   0.0042   1.0000
   9.000   1.1810   0.01868   0.01368  -0.0393   0.0040   1.0000
   9.250   1.2032   0.01971   0.01484  -0.0386   0.0037   1.0000
   9.500   1.2252   0.02072   0.01597  -0.0380   0.0034   1.0000
   9.750   1.2463   0.02182   0.01718  -0.0374   0.0032   1.0000
  10.000   1.2607   0.02428   0.01991  -0.0362   0.0029   1.0000
  10.250   1.2530   0.03097   0.02731  -0.0333   0.0026   1.0000
  10.500   1.2668   0.03285   0.02941  -0.0324   0.0026   1.0000
  10.750   1.2732   0.03571   0.03254  -0.0311   0.0025   1.0000
  11.000   1.2720   0.03925   0.03638  -0.0297   0.0025   1.0000
  11.250   1.2605   0.04312   0.04051  -0.0280   0.0025   1.0000
  11.500   1.2430   0.04747   0.04507  -0.0281   0.0025   1.0000
  11.750   1.2247   0.05414   0.05196  -0.0331   0.0025   1.0000
  12.000   1.2051   0.06385   0.06188  -0.0421   0.0026   1.0000
  12.250   1.1777   0.07624   0.07443  -0.0519   0.0026   1.0000
  14.000   1.0775   0.13330   0.13184  -0.0806   0.0028   1.0000
  14.250   1.0645   0.14177   0.14034  -0.0844   0.0029   1.0000
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