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AG16 (ag16-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AG16 (ag16-il)
Reynolds number: 50,000
Max Cl/Cd: 35.97 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag16-il-50000-n5.txt
Download as CSV file: xf-ag16-il-50000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5440   0.09615   0.08944   0.0012   1.0000   0.0408
  -8.000  -0.5420   0.09176   0.08512  -0.0015   1.0000   0.0406
  -7.750  -0.5401   0.08724   0.08067  -0.0048   1.0000   0.0404
  -7.500  -0.5351   0.08211   0.07561  -0.0096   1.0000   0.0403
  -7.250  -0.5279   0.07641   0.06994  -0.0154   1.0000   0.0402
  -7.000  -0.5183   0.06991   0.06342  -0.0223   1.0000   0.0400
  -6.750  -0.5069   0.06293   0.05634  -0.0291   1.0000   0.0398
  -6.500  -0.4936   0.05646   0.04965  -0.0349   1.0000   0.0399
  -6.250  -0.4771   0.05038   0.04321  -0.0397   1.0000   0.0402
  -6.000  -0.4575   0.04469   0.03703  -0.0433   1.0000   0.0405
  -5.750  -0.4350   0.03968   0.03144  -0.0457   1.0000   0.0409
  -5.500  -0.4105   0.03547   0.02653  -0.0471   1.0000   0.0420
  -5.250  -0.3846   0.03193   0.02237  -0.0477   1.0000   0.0439
  -5.000  -0.3581   0.02915   0.01900  -0.0478   1.0000   0.0486
  -4.750  -0.3331   0.02721   0.01689  -0.0477   1.0000   0.0553
  -4.500  -0.3066   0.02504   0.01430  -0.0470   1.0000   0.0609
  -4.250  -0.2813   0.02346   0.01262  -0.0466   1.0000   0.0743
  -4.000  -0.2558   0.02185   0.01101  -0.0462   1.0000   0.0918
  -3.750  -0.2298   0.02048   0.00970  -0.0461   1.0000   0.1295
  -3.500  -0.2033   0.01924   0.00863  -0.0462   1.0000   0.1925
  -3.250  -0.1774   0.01835   0.00800  -0.0461   1.0000   0.2703
  -3.000  -0.1521   0.01771   0.00752  -0.0457   1.0000   0.3461
  -2.750  -0.1274   0.01719   0.00714  -0.0450   1.0000   0.4159
  -2.500  -0.1034   0.01669   0.00680  -0.0441   1.0000   0.4817
  -2.250  -0.0801   0.01621   0.00643  -0.0430   1.0000   0.5479
  -2.000  -0.0582   0.01571   0.00613  -0.0414   1.0000   0.6183
  -1.750  -0.0385   0.01516   0.00585  -0.0391   1.0000   0.7008
  -1.500  -0.0194   0.01451   0.00553  -0.0362   1.0000   0.8318
  -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   0.1624   0.01518   0.00502  -0.0378   0.9982   1.0000
   0.500   0.2110   0.01545   0.00521  -0.0424   0.9774   1.0000
   0.750   0.2596   0.01568   0.00541  -0.0468   0.9565   1.0000
   1.000   0.3047   0.01585   0.00557  -0.0503   0.9331   1.0000
   1.250   0.3466   0.01598   0.00572  -0.0528   0.9077   1.0000
   1.500   0.3848   0.01609   0.00585  -0.0543   0.8799   1.0000
   1.750   0.4193   0.01621   0.00602  -0.0550   0.8501   1.0000
   2.000   0.4509   0.01634   0.00614  -0.0548   0.8191   1.0000
   2.250   0.4795   0.01649   0.00628  -0.0541   0.7867   1.0000
   2.500   0.5062   0.01669   0.00645  -0.0530   0.7529   1.0000
   2.750   0.5321   0.01691   0.00669  -0.0517   0.7195   1.0000
   3.000   0.5570   0.01719   0.00693  -0.0504   0.6841   1.0000
   3.250   0.5818   0.01750   0.00719  -0.0490   0.6492   1.0000
   3.500   0.6063   0.01785   0.00750  -0.0477   0.6131   1.0000
   3.750   0.6308   0.01824   0.00792  -0.0465   0.5765   1.0000
   4.000   0.6551   0.01867   0.00830  -0.0453   0.5400   1.0000
   4.250   0.6793   0.01915   0.00876  -0.0442   0.5022   1.0000
   4.500   0.7034   0.01967   0.00927  -0.0432   0.4641   1.0000
   4.750   0.7277   0.02025   0.00983  -0.0423   0.4259   1.0000
   5.000   0.7518   0.02090   0.01051  -0.0414   0.3877   1.0000
   5.250   0.7757   0.02162   0.01121  -0.0406   0.3492   1.0000
   5.500   0.7993   0.02240   0.01200  -0.0399   0.3103   1.0000
   5.750   0.8225   0.02328   0.01288  -0.0392   0.2720   1.0000
   6.000   0.8451   0.02428   0.01384  -0.0385   0.2347   1.0000
   6.250   0.8671   0.02541   0.01494  -0.0378   0.1992   1.0000
   6.500   0.8887   0.02669   0.01624  -0.0371   0.1644   1.0000
   6.750   0.9096   0.02817   0.01779  -0.0364   0.1347   1.0000
   7.000   0.9301   0.02989   0.01955  -0.0355   0.1105   1.0000
   7.250   0.9493   0.03171   0.02131  -0.0347   0.0912   1.0000
   7.500   0.9691   0.03381   0.02359  -0.0337   0.0767   1.0000
   7.750   0.9876   0.03599   0.02589  -0.0327   0.0652   1.0000
   8.000   1.0061   0.03829   0.02847  -0.0317   0.0559   1.0000
   8.250   1.0242   0.04151   0.03205  -0.0305   0.0509   1.0000
   8.500   1.0394   0.04424   0.03513  -0.0296   0.0450   1.0000
   8.750   1.0514   0.04746   0.03858  -0.0287   0.0410   1.0000
   9.000   1.0599   0.05184   0.04361  -0.0276   0.0391   1.0000
   9.250   1.0626   0.05653   0.04894  -0.0268   0.0378   1.0000
   9.500   1.0593   0.06136   0.05423  -0.0263   0.0370   1.0000
   9.750   1.0500   0.06628   0.05952  -0.0263   0.0365   1.0000
  10.000   1.0337   0.07131   0.06483  -0.0268   0.0365   1.0000
  10.250   1.0131   0.07721   0.07093  -0.0294   0.0370   1.0000
  10.500   0.9914   0.08455   0.07832  -0.0346   0.0378   1.0000
  10.750   0.9706   0.09328   0.08713  -0.0414   0.0386   1.0000
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