Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

AG17 (ag17-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AG17 (ag17-il)
Reynolds number: 50,000
Max Cl/Cd: 36.33 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag17-il-50000-n5.txt
Download as CSV file: xf-ag17-il-50000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5347   0.10167   0.09485   0.0046   1.0000   0.0466
  -8.000  -0.5298   0.09764   0.09086   0.0030   1.0000   0.0448
  -7.750  -0.5262   0.09347   0.08676   0.0007   1.0000   0.0431
  -7.500  -0.5221   0.08900   0.08236  -0.0028   1.0000   0.0417
  -7.000  -0.5016   0.07459   0.06793  -0.0218   1.0000   0.0369
  -6.750  -0.4892   0.06909   0.06237  -0.0262   1.0000   0.0367
  -6.500  -0.4750   0.06332   0.05647  -0.0312   1.0000   0.0368
  -6.000  -0.4452   0.05499   0.04802  -0.0362   1.0000   0.0399
  -5.750  -0.4248   0.04972   0.04246  -0.0403   1.0000   0.0405
  -5.500  -0.4018   0.04438   0.03668  -0.0438   1.0000   0.0407
  -5.250  -0.3766   0.03944   0.03118  -0.0465   1.0000   0.0407
  -5.000  -0.3504   0.03530   0.02646  -0.0480   1.0000   0.0414
  -4.750  -0.3234   0.03182   0.02241  -0.0488   1.0000   0.0428
  -4.500  -0.2959   0.02895   0.01896  -0.0491   1.0000   0.0459
  -4.250  -0.2692   0.02669   0.01630  -0.0490   1.0000   0.0521
  -4.000  -0.2427   0.02475   0.01409  -0.0486   1.0000   0.0581
  -3.750  -0.2166   0.02291   0.01197  -0.0479   1.0000   0.0658
  -3.500  -0.1906   0.02156   0.01051  -0.0476   1.0000   0.0836
  -3.250  -0.1641   0.02015   0.00907  -0.0472   1.0000   0.1111
  -3.000  -0.1368   0.01878   0.00790  -0.0473   1.0000   0.1704
  -2.750  -0.1104   0.01763   0.00722  -0.0474   1.0000   0.2718
  -2.500  -0.0854   0.01682   0.00679  -0.0469   1.0000   0.3848
  -2.250  -0.0620   0.01616   0.00640  -0.0459   1.0000   0.4908
  -2.000  -0.0407   0.01548   0.00607  -0.0441   1.0000   0.5973
  -1.750  -0.0235   0.01469   0.00575  -0.0410   1.0000   0.7303
  -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   0.2041   0.01515   0.00485  -0.0396   0.9996   1.0000
   0.750   0.2531   0.01543   0.00508  -0.0444   0.9775   1.0000
   1.000   0.3026   0.01566   0.00529  -0.0488   0.9556   1.0000
   1.250   0.3476   0.01581   0.00547  -0.0522   0.9304   1.0000
   1.500   0.3890   0.01594   0.00562  -0.0545   0.9029   1.0000
   1.750   0.4270   0.01604   0.00580  -0.0558   0.8731   1.0000
   2.000   0.4599   0.01616   0.00594  -0.0560   0.8398   1.0000
   2.250   0.4903   0.01629   0.00607  -0.0556   0.8055   1.0000
   2.500   0.5177   0.01647   0.00623  -0.0545   0.7690   1.0000
   2.750   0.5437   0.01668   0.00640  -0.0531   0.7316   1.0000
   3.000   0.5686   0.01694   0.00667  -0.0516   0.6930   1.0000
   3.250   0.5930   0.01725   0.00692  -0.0501   0.6528   1.0000
   3.500   0.6171   0.01761   0.00721  -0.0486   0.6116   1.0000
   3.750   0.6410   0.01803   0.00755  -0.0472   0.5694   1.0000
   4.000   0.6649   0.01849   0.00802  -0.0460   0.5265   1.0000
   4.250   0.6887   0.01902   0.00848  -0.0448   0.4837   1.0000
   4.500   0.7124   0.01961   0.00901  -0.0437   0.4414   1.0000
   4.750   0.7364   0.02027   0.00961  -0.0428   0.3998   1.0000
   5.000   0.7604   0.02100   0.01029  -0.0420   0.3590   1.0000
   5.250   0.7843   0.02181   0.01113  -0.0413   0.3198   1.0000
   5.500   0.8077   0.02269   0.01197  -0.0406   0.2824   1.0000
   5.750   0.8311   0.02365   0.01295  -0.0400   0.2454   1.0000
   6.000   0.8540   0.02472   0.01400  -0.0394   0.2116   1.0000
   6.250   0.8766   0.02590   0.01523  -0.0387   0.1797   1.0000
   6.500   0.8987   0.02725   0.01661  -0.0380   0.1521   1.0000
   6.750   0.9201   0.02869   0.01809  -0.0373   0.1261   1.0000
   7.000   0.9409   0.03043   0.01992  -0.0364   0.1076   1.0000
   7.250   0.9616   0.03218   0.02188  -0.0356   0.0882   1.0000
   7.500   0.9821   0.03444   0.02432  -0.0346   0.0763   1.0000
   7.750   1.0009   0.03659   0.02667  -0.0337   0.0645   1.0000
   8.000   1.0188   0.03900   0.02915  -0.0328   0.0567   1.0000
   8.250   1.0368   0.04250   0.03327  -0.0316   0.0512   1.0000
   8.500   1.0512   0.04488   0.03575  -0.0309   0.0451   1.0000
   8.750   1.0620   0.04899   0.04050  -0.0299   0.0417   1.0000
   9.000   1.0680   0.05368   0.04579  -0.0289   0.0399   1.0000
   9.250   1.0685   0.05856   0.05118  -0.0283   0.0387   1.0000
   9.500   1.0633   0.06353   0.05656  -0.0281   0.0379   1.0000
   9.750   1.0520   0.06867   0.06204  -0.0285   0.0375   1.0000
  10.000   1.0325   0.07417   0.06779  -0.0297   0.0379   1.0000
  10.250   1.0095   0.08102   0.07482  -0.0338   0.0387   1.0000
  10.500   0.9870   0.08955   0.08345  -0.0406   0.0397   1.0000
<< Back to AG17 (ag17-il)

Polar data table (+)

Polar graphs


<< Back to AG17 (ag17-il)