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AG10 (ag10-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG10 (ag10-il)
Reynolds number: 100,000
Max Cl/Cd: 41.22 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag10-il-100000.txt
Download as CSV file: xf-ag10-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG10                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.6246   0.12520   0.12018   0.0306   1.0000   0.0621
  -9.000  -0.6274   0.12331   0.11835   0.0258   1.0000   0.0626
  -8.750  -0.6291   0.12100   0.11610   0.0208   1.0000   0.0629
  -8.500  -0.6178   0.11390   0.10902   0.0233   1.0000   0.0639
  -8.250  -0.6034   0.10882   0.10392   0.0263   1.0000   0.0660
  -8.000  -0.5964   0.10524   0.10033   0.0257   1.0000   0.0681
  -7.750  -0.5914   0.10185   0.09698   0.0240   1.0000   0.0708
  -7.500  -0.5885   0.09867   0.09385   0.0204   1.0000   0.0737
  -7.250  -0.5788   0.09562   0.09080   0.0059   1.0000   0.0759
  -7.000  -0.5699   0.09003   0.08523   0.0031   1.0000   0.0772
  -6.750  -0.5620   0.08613   0.08138   0.0099   1.0000   0.0803
  -6.500  -0.5493   0.08233   0.07758   0.0076   1.0000   0.0841
  -6.250  -0.5167   0.07879   0.07364  -0.0128   1.0000   0.0904
  -6.000  -0.5119   0.07269   0.06779  -0.0075   1.0000   0.0921
  -5.750  -0.4995   0.06911   0.06425  -0.0057   1.0000   0.0953
  -5.500  -0.4646   0.06488   0.05960  -0.0181   1.0000   0.1054
  -5.250  -0.4546   0.06025   0.05515  -0.0151   1.0000   0.1077
  -5.000  -0.4353   0.05689   0.05176  -0.0158   1.0000   0.1142
  -4.750  -0.4092   0.05251   0.04720  -0.0199   1.0000   0.1232
  -4.500  -0.3830   0.04894   0.04343  -0.0229   1.0000   0.1368
  -4.250  -0.3598   0.04573   0.04014  -0.0241   1.0000   0.1522
  -4.000  -0.3357   0.04257   0.03686  -0.0251   1.0000   0.1677
  -3.750  -0.2954   0.02255   0.01735  -0.0239   1.0000   0.1854
  -3.500  -0.2519   0.02968   0.02234  -0.0317   1.0000   0.0866
  -3.250  -0.2208   0.02602   0.01815  -0.0322   1.0000   0.0828
  -3.000  -0.1887   0.02269   0.01407  -0.0323   1.0000   0.0794
  -2.750  -0.1600   0.02077   0.01190  -0.0320   1.0000   0.0833
  -2.500  -0.1305   0.01924   0.00992  -0.0316   1.0000   0.0895
  -2.250  -0.1029   0.01761   0.00823  -0.0313   1.0000   0.0987
  -2.000  -0.0752   0.01627   0.00683  -0.0308   1.0000   0.1107
  -1.750  -0.0480   0.01524   0.00582  -0.0303   1.0000   0.1304
  -1.500  -0.0214   0.01419   0.00494  -0.0299   1.0000   0.1592
  -1.250   0.0048   0.01320   0.00419  -0.0293   1.0000   0.2081
  -1.000   0.0294   0.01124   0.00356  -0.0289   1.0000   0.4582
  -0.750   0.0546   0.00978   0.00309  -0.0268   1.0000   1.0000
  -0.500   0.0811   0.00979   0.00289  -0.0263   1.0000   1.0000
  -0.250   0.1073   0.00982   0.00276  -0.0259   1.0000   1.0000
   0.000   0.1333   0.00985   0.00269  -0.0255   1.0000   1.0000
   0.250   0.1592   0.00990   0.00268  -0.0250   1.0000   1.0000
   0.500   0.1848   0.00997   0.00270  -0.0246   1.0000   1.0000
   0.750   0.2103   0.01005   0.00278  -0.0242   1.0000   1.0000
   1.000   0.2356   0.01016   0.00290  -0.0238   1.0000   1.0000
   1.250   0.2607   0.01029   0.00307  -0.0234   1.0000   1.0000
   1.500   0.2857   0.01045   0.00331  -0.0232   1.0000   1.0000
   1.750   0.3103   0.01067   0.00363  -0.0231   1.0000   1.0000
   2.000   0.3688   0.01075   0.00387  -0.0295   0.9667   1.0000
   2.250   0.4252   0.01070   0.00391  -0.0339   0.8818   1.0000
   2.500   0.4514   0.01095   0.00389  -0.0315   0.7681   1.0000
   2.750   0.4717   0.01152   0.00399  -0.0284   0.6594   1.0000
   3.000   0.4944   0.01221   0.00422  -0.0266   0.5725   1.0000
   3.250   0.5186   0.01291   0.00455  -0.0255   0.5062   1.0000
   3.500   0.5436   0.01359   0.00498  -0.0247   0.4522   1.0000
   3.750   0.5689   0.01428   0.00542  -0.0240   0.4055   1.0000
   4.000   0.5946   0.01494   0.00591  -0.0235   0.3621   1.0000
   4.250   0.6204   0.01560   0.00645  -0.0230   0.3215   1.0000
   4.500   0.6460   0.01632   0.00704  -0.0225   0.2841   1.0000
   4.750   0.6718   0.01702   0.00767  -0.0221   0.2466   1.0000
   5.000   0.6972   0.01784   0.00833  -0.0217   0.2120   1.0000
   5.250   0.7228   0.01869   0.00920  -0.0213   0.1763   1.0000
   5.500   0.7478   0.01974   0.01012  -0.0208   0.1440   1.0000
   5.750   0.7726   0.02104   0.01132  -0.0203   0.1158   1.0000
   6.000   0.7978   0.02258   0.01296  -0.0197   0.0938   1.0000
   6.250   0.8228   0.02435   0.01481  -0.0191   0.0793   1.0000
   6.500   0.8477   0.02644   0.01709  -0.0184   0.0692   1.0000
   6.750   0.8713   0.02929   0.01995  -0.0179   0.0630   1.0000
   7.000   0.8955   0.03154   0.02287  -0.0171   0.0577   1.0000
   7.250   0.9177   0.03464   0.02638  -0.0164   0.0552   1.0000
   7.500   0.9376   0.03851   0.03076  -0.0157   0.0542   1.0000
   7.750   0.9544   0.04279   0.03551  -0.0152   0.0534   1.0000
   8.000   0.9665   0.04811   0.04121  -0.0150   0.0523   1.0000
   8.250   0.9751   0.05373   0.04733  -0.0150   0.0522   1.0000
   8.500   0.9519   0.06657   0.06147  -0.0187   0.0604   1.0000
   8.750   0.9515   0.07250   0.06753  -0.0200   0.0620   1.0000
   9.000   0.8917   0.10386   0.09922  -0.0464   0.1292   1.0000
   9.250   0.8994   0.10943   0.10478  -0.0448   0.1276   1.0000
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