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AG17 (ag17-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AG17 (ag17-il)
Reynolds number: 50,000
Max Cl/Cd: 34.56 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag17-il-50000.txt
Download as CSV file: xf-ag17-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG17                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5525   0.11587   0.10926   0.0146   1.0000   0.1833
  -8.250  -0.5338   0.11018   0.10355   0.0163   1.0000   0.1916
  -8.000  -0.5478   0.10947   0.10299   0.0118   1.0000   0.1973
  -7.750  -0.5321   0.10450   0.09802   0.0133   1.0000   0.2097
  -7.500  -0.5217   0.10040   0.09396   0.0138   1.0000   0.2212
  -7.250  -0.5152   0.09681   0.09043   0.0135   1.0000   0.2323
  -7.000  -0.5113   0.09353   0.08723   0.0128   1.0000   0.2437
  -6.750  -0.5065   0.09023   0.08399   0.0116   1.0000   0.2563
  -6.500  -0.4994   0.08676   0.08059   0.0103   1.0000   0.2696
  -6.250  -0.4904   0.08319   0.07707   0.0096   1.0000   0.2840
  -6.000  -0.4804   0.07961   0.07354   0.0091   1.0000   0.2997
  -5.500  -0.4617   0.07270   0.06673   0.0067   1.0000   0.3376
  -5.250  -0.4485   0.06906   0.06307   0.0073   1.0000   0.3561
  -5.000  -0.3695   0.04711   0.03965  -0.0438   1.0000   0.1285
  -4.750  -0.3395   0.04090   0.03285  -0.0472   1.0000   0.1143
  -4.500  -0.3118   0.03639   0.02793  -0.0487   1.0000   0.1107
  -4.250  -0.2812   0.03254   0.02329  -0.0502   1.0000   0.1144
  -4.000  -0.2536   0.02936   0.01974  -0.0505   1.0000   0.1201
  -3.750  -0.2236   0.02645   0.01627  -0.0504   1.0000   0.1263
  -3.500  -0.1962   0.02428   0.01393  -0.0501   1.0000   0.1472
  -3.250  -0.1683   0.02209   0.01163  -0.0494   1.0000   0.1750
  -3.000  -0.1420   0.02029   0.00999  -0.0486   1.0000   0.2395
  -2.750  -0.1172   0.01861   0.00894  -0.0474   1.0000   0.3640
  -2.500  -0.0966   0.01714   0.00824  -0.0449   1.0000   0.5228
  -2.250  -0.0820   0.01582   0.00757  -0.0405   1.0000   0.6752
  -2.000  -0.0582   0.01408   0.00647  -0.0370   1.0000   1.0000
  -1.750  -0.0244   0.01402   0.00568  -0.0395   1.0000   1.0000
  -1.500   0.0034   0.01403   0.00524  -0.0400   1.0000   1.0000
  -1.250   0.0299   0.01407   0.00493  -0.0401   1.0000   1.0000
  -1.000   0.0558   0.01413   0.00467  -0.0401   1.0000   1.0000
  -0.750   0.0813   0.01421   0.00453  -0.0400   1.0000   1.0000
  -0.500   0.1064   0.01433   0.00446  -0.0398   1.0000   1.0000
  -0.250   0.1311   0.01447   0.00446  -0.0397   1.0000   1.0000
   0.000   0.1555   0.01466   0.00450  -0.0396   1.0000   1.0000
   0.250   0.1796   0.01488   0.00464  -0.0395   1.0000   1.0000
   0.500   0.2036   0.01515   0.00485  -0.0395   1.0000   1.0000
   0.750   0.2272   0.01549   0.00514  -0.0397   1.0000   1.0000
   1.000   0.2503   0.01589   0.00553  -0.0399   1.0000   1.0000
   1.250   0.2727   0.01638   0.00602  -0.0402   1.0000   1.0000
   1.500   0.2944   0.01698   0.00663  -0.0408   1.0000   1.0000
   1.750   0.3152   0.01771   0.00740  -0.0415   1.0000   1.0000
   2.000   0.3350   0.01861   0.00835  -0.0425   1.0000   1.0000
   2.250   0.4128   0.01953   0.00952  -0.0537   0.9686   1.0000
   2.500   0.4950   0.01987   0.01016  -0.0640   0.9270   1.0000
   2.750   0.5564   0.01983   0.01039  -0.0687   0.8837   1.0000
   3.000   0.5975   0.01973   0.01055  -0.0687   0.8395   1.0000
   3.250   0.6271   0.01970   0.01061  -0.0663   0.7950   1.0000
   3.500   0.6510   0.01976   0.01068  -0.0628   0.7504   1.0000
   3.750   0.6727   0.02000   0.01089  -0.0594   0.7031   1.0000
   4.000   0.6943   0.02032   0.01122  -0.0561   0.6550   1.0000
   4.250   0.7159   0.02080   0.01160  -0.0531   0.6052   1.0000
   4.500   0.7379   0.02135   0.01199  -0.0504   0.5566   1.0000
   4.750   0.7600   0.02210   0.01263  -0.0481   0.5055   1.0000
   5.000   0.7824   0.02296   0.01335  -0.0461   0.4563   1.0000
   5.250   0.8048   0.02395   0.01426  -0.0442   0.4081   1.0000
   5.500   0.8273   0.02511   0.01528  -0.0426   0.3616   1.0000
   5.750   0.8497   0.02644   0.01654  -0.0411   0.3167   1.0000
   6.000   0.8722   0.02799   0.01801  -0.0397   0.2758   1.0000
   6.250   0.8949   0.02975   0.01966  -0.0384   0.2394   1.0000
   6.500   0.9166   0.03157   0.02154  -0.0372   0.2048   1.0000
   6.750   0.9393   0.03420   0.02437  -0.0360   0.1795   1.0000
   7.000   0.9598   0.03650   0.02689  -0.0349   0.1543   1.0000
   7.250   0.9806   0.04012   0.03093  -0.0337   0.1398   1.0000
   7.500   0.9993   0.04308   0.03412  -0.0327   0.1237   1.0000
   7.750   1.0136   0.04761   0.03924  -0.0317   0.1154   1.0000
   8.000   1.0246   0.05283   0.04506  -0.0309   0.1116   1.0000
   8.250   1.0363   0.05742   0.04984  -0.0302   0.1058   1.0000
   8.500   1.0334   0.06355   0.05664  -0.0304   0.1044   1.0000
   8.750   1.0268   0.06987   0.06338  -0.0312   0.1041   1.0000
   9.000   1.0173   0.07624   0.07003  -0.0325   0.1048   1.0000
   9.250   1.0074   0.08253   0.07646  -0.0341   0.1058   1.0000
   9.500   0.9960   0.08892   0.08298  -0.0363   0.1080   1.0000
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