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AG16 (ag16-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AG16 (ag16-il)
Reynolds number: 50,000
Max Cl/Cd: 33.51 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag16-il-50000.txt
Download as CSV file: xf-ag16-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG16                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5356   0.10901   0.10241   0.0167   1.0000   0.2187
  -8.000  -0.5313   0.10539   0.09885   0.0159   1.0000   0.2270
  -7.750  -0.5430   0.10417   0.09776   0.0126   1.0000   0.2360
  -7.500  -0.5243   0.09897   0.09255   0.0143   1.0000   0.2473
  -7.250  -0.5168   0.09505   0.08868   0.0140   1.0000   0.2558
  -6.750  -0.5159   0.08906   0.08287   0.0101   1.0000   0.2786
  -6.500  -0.5094   0.08559   0.07946   0.0083   1.0000   0.2920
  -6.250  -0.5005   0.08187   0.07578   0.0072   1.0000   0.3058
  -6.000  -0.4904   0.07802   0.07197   0.0062   1.0000   0.3197
  -5.750  -0.4584   0.06258   0.05615  -0.0263   1.0000   0.1796
  -5.500  -0.4220   0.04699   0.03929  -0.0447   1.0000   0.1133
  -5.250  -0.3990   0.04222   0.03436  -0.0457   1.0000   0.1100
  -5.000  -0.3721   0.03733   0.02885  -0.0479   1.0000   0.1099
  -4.750  -0.3442   0.03325   0.02407  -0.0492   1.0000   0.1150
  -4.500  -0.3170   0.03009   0.02059  -0.0493   1.0000   0.1209
  -4.250  -0.2884   0.02707   0.01708  -0.0493   1.0000   0.1312
  -4.000  -0.2609   0.02489   0.01464  -0.0490   1.0000   0.1555
  -3.750  -0.2342   0.02292   0.01260  -0.0483   1.0000   0.1954
  -3.500  -0.2091   0.02122   0.01120  -0.0474   1.0000   0.2628
  -3.250  -0.1861   0.01988   0.01033  -0.0458   1.0000   0.3545
  -3.000  -0.1642   0.01885   0.00965  -0.0438   1.0000   0.4505
  -2.750  -0.1430   0.01800   0.00907  -0.0414   1.0000   0.5411
  -2.500  -0.1240   0.01716   0.00849  -0.0385   1.0000   0.6254
  -2.250  -0.1082   0.01624   0.00785  -0.0346   1.0000   0.7131
  -2.000  -0.0985   0.01517   0.00720  -0.0290   1.0000   0.8214
  -1.750  -0.0471   0.01423   0.00595  -0.0347   1.0000   1.0000
  -1.500  -0.0165   0.01421   0.00544  -0.0364   1.0000   1.0000
  -1.250   0.0109   0.01424   0.00513  -0.0370   1.0000   1.0000
  -1.000   0.0371   0.01431   0.00489  -0.0372   1.0000   1.0000
  -0.750   0.0626   0.01440   0.00477  -0.0373   1.0000   1.0000
  -0.500   0.0876   0.01454   0.00473  -0.0373   1.0000   1.0000
  -0.250   0.1121   0.01470   0.00476  -0.0373   1.0000   1.0000
   0.000   0.1364   0.01492   0.00484  -0.0373   1.0000   1.0000
   0.250   0.1604   0.01518   0.00502  -0.0374   1.0000   1.0000
   0.500   0.1840   0.01550   0.00528  -0.0376   1.0000   1.0000
   0.750   0.2071   0.01590   0.00564  -0.0378   1.0000   1.0000
   1.000   0.2295   0.01638   0.00611  -0.0382   1.0000   1.0000
   1.250   0.2511   0.01697   0.00669  -0.0388   1.0000   1.0000
   1.500   0.2720   0.01768   0.00742  -0.0396   1.0000   1.0000
   1.750   0.2918   0.01854   0.00830  -0.0405   1.0000   1.0000
   2.000   0.3542   0.01950   0.00940  -0.0492   0.9764   1.0000
   2.250   0.4348   0.02008   0.01024  -0.0597   0.9405   1.0000
   2.500   0.5022   0.02026   0.01068  -0.0665   0.9024   1.0000
   2.750   0.5515   0.02025   0.01094  -0.0688   0.8639   1.0000
   3.000   0.5883   0.02023   0.01105  -0.0681   0.8255   1.0000
   3.250   0.6173   0.02024   0.01115  -0.0658   0.7873   1.0000
   3.500   0.6417   0.02035   0.01129  -0.0628   0.7482   1.0000
   3.750   0.6641   0.02058   0.01159  -0.0596   0.7073   1.0000
   4.000   0.6862   0.02090   0.01188  -0.0566   0.6657   1.0000
   4.250   0.7082   0.02132   0.01225  -0.0538   0.6226   1.0000
   4.500   0.7303   0.02186   0.01275  -0.0512   0.5777   1.0000
   4.750   0.7527   0.02246   0.01328  -0.0489   0.5326   1.0000
   5.000   0.7752   0.02314   0.01390  -0.0466   0.4877   1.0000
   5.250   0.7976   0.02397   0.01464  -0.0446   0.4415   1.0000
   5.500   0.8198   0.02490   0.01548  -0.0427   0.3946   1.0000
   5.750   0.8420   0.02593   0.01630  -0.0409   0.3490   1.0000
   6.000   0.8636   0.02724   0.01757  -0.0393   0.3021   1.0000
   6.250   0.8849   0.02869   0.01889  -0.0378   0.2573   1.0000
   6.500   0.9066   0.03060   0.02075  -0.0363   0.2181   1.0000
   6.750   0.9285   0.03269   0.02271  -0.0351   0.1849   1.0000
   7.000   0.9497   0.03510   0.02516  -0.0339   0.1573   1.0000
   7.250   0.9702   0.03839   0.02899  -0.0326   0.1388   1.0000
   7.500   0.9892   0.04152   0.03241  -0.0315   0.1228   1.0000
   7.750   1.0055   0.04547   0.03682  -0.0304   0.1124   1.0000
   8.000   1.0221   0.04973   0.04139  -0.0294   0.1065   1.0000
   8.250   1.0299   0.05498   0.04721  -0.0286   0.1026   1.0000
   8.500   1.0295   0.06066   0.05356  -0.0284   0.1002   1.0000
   8.750   1.0245   0.06662   0.06000  -0.0287   0.0992   1.0000
   9.000   1.0139   0.07301   0.06673  -0.0298   0.1002   1.0000
   9.250   1.0010   0.07940   0.07331  -0.0314   0.1018   1.0000
   9.500   0.9886   0.08552   0.07953  -0.0332   0.1031   1.0000
   9.750   0.9804   0.09144   0.08548  -0.0347   0.1042   1.0000
  10.000   0.9339   0.10847   0.10249  -0.0504   0.1307   1.0000
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