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AG08 (ag08-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG08 (ag08-il)
Reynolds number: 100,000
Max Cl/Cd: 45.62 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag08-il-100000.txt
Download as CSV file: xf-ag08-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG08                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5762   0.11263   0.10781   0.0154   1.0000   0.0737
  -8.500  -0.5820   0.11040   0.10567   0.0090   1.0000   0.0745
  -8.250  -0.5836   0.10728   0.10261   0.0006   1.0000   0.0748
  -8.000  -0.5624   0.10032   0.09561   0.0117   1.0000   0.0781
  -7.750  -0.5554   0.09679   0.09211   0.0111   1.0000   0.0810
  -7.500  -0.5520   0.09329   0.08865   0.0083   1.0000   0.0843
  -7.250  -0.5463   0.08915   0.08453  -0.0079   1.0000   0.0880
  -7.000  -0.5383   0.08371   0.07913  -0.0082   1.0000   0.0895
  -6.750  -0.5278   0.08032   0.07573  -0.0039   1.0000   0.0925
  -6.500  -0.5142   0.07623   0.07163  -0.0083   1.0000   0.0987
  -6.250  -0.4993   0.07068   0.06602  -0.0160   1.0000   0.1041
  -6.000  -0.4762   0.06633   0.06134  -0.0265   1.0000   0.1159
  -5.750  -0.4684   0.06292   0.05821  -0.0194   1.0000   0.1202
  -5.500  -0.4470   0.05806   0.05318  -0.0253   1.0000   0.1315
  -5.250  -0.4250   0.05399   0.04891  -0.0291   1.0000   0.1444
  -5.000  -0.4083   0.05067   0.04567  -0.0276   1.0000   0.1504
  -4.750  -0.3596   0.03572   0.02914  -0.0387   1.0000   0.0764
  -4.500  -0.3337   0.03118   0.02424  -0.0391   1.0000   0.0692
  -4.250  -0.3061   0.02765   0.02007  -0.0393   1.0000   0.0715
  -4.000  -0.2781   0.02440   0.01619  -0.0391   1.0000   0.0722
  -3.750  -0.2496   0.02199   0.01311  -0.0385   1.0000   0.0746
  -3.500  -0.2238   0.02015   0.01117  -0.0382   1.0000   0.0852
  -3.250  -0.1968   0.01826   0.00909  -0.0376   1.0000   0.0936
  -3.000  -0.1705   0.01710   0.00786  -0.0370   1.0000   0.1113
  -2.750  -0.1443   0.01581   0.00656  -0.0364   1.0000   0.1319
  -2.500  -0.1187   0.01479   0.00572  -0.0358   1.0000   0.1645
  -2.250  -0.0934   0.01381   0.00495  -0.0353   1.0000   0.2150
  -2.000  -0.0684   0.01266   0.00438  -0.0348   1.0000   0.3139
  -1.750  -0.0476   0.01090   0.00402  -0.0332   1.0000   0.6122
  -1.500  -0.0149   0.00997   0.00362  -0.0326   1.0000   1.0000
  -1.250   0.0104   0.00998   0.00335  -0.0321   1.0000   1.0000
  -1.000   0.0355   0.01001   0.00318  -0.0316   1.0000   1.0000
  -0.750   0.0606   0.01006   0.00307  -0.0311   1.0000   1.0000
  -0.500   0.0854   0.01012   0.00301  -0.0307   1.0000   1.0000
  -0.250   0.1102   0.01021   0.00299  -0.0302   1.0000   1.0000
   0.000   0.1348   0.01031   0.00301  -0.0298   1.0000   1.0000
   0.250   0.1593   0.01045   0.00309  -0.0294   1.0000   1.0000
   0.500   0.1837   0.01061   0.00323  -0.0292   1.0000   1.0000
   0.750   0.2078   0.01083   0.00343  -0.0290   1.0000   1.0000
   1.000   0.2315   0.01109   0.00371  -0.0289   1.0000   1.0000
   1.250   0.2547   0.01144   0.00408  -0.0290   1.0000   1.0000
   1.500   0.3192   0.01159   0.00431  -0.0367   0.9753   1.0000
   1.750   0.3797   0.01155   0.00441  -0.0429   0.9477   1.0000
   2.000   0.4279   0.01142   0.00438  -0.0461   0.9127   1.0000
   2.250   0.4621   0.01135   0.00434  -0.0460   0.8672   1.0000
   2.500   0.4887   0.01139   0.00432  -0.0442   0.8166   1.0000
   2.750   0.5119   0.01156   0.00439  -0.0418   0.7608   1.0000
   3.000   0.5345   0.01185   0.00450  -0.0396   0.7022   1.0000
   3.250   0.5574   0.01225   0.00468  -0.0378   0.6430   1.0000
   3.500   0.5807   0.01273   0.00492  -0.0362   0.5862   1.0000
   3.750   0.6044   0.01327   0.00526  -0.0350   0.5330   1.0000
   4.000   0.6285   0.01385   0.00563  -0.0339   0.4835   1.0000
   4.250   0.6527   0.01446   0.00606  -0.0330   0.4377   1.0000
   4.500   0.6772   0.01509   0.00655  -0.0322   0.3940   1.0000
   4.750   0.7016   0.01575   0.00708  -0.0314   0.3523   1.0000
   5.000   0.7259   0.01644   0.00770  -0.0307   0.3115   1.0000
   5.250   0.7500   0.01718   0.00833  -0.0300   0.2707   1.0000
   5.500   0.7738   0.01800   0.00904  -0.0294   0.2291   1.0000
   5.750   0.7972   0.01895   0.00990  -0.0287   0.1862   1.0000
   6.000   0.8201   0.02012   0.01094  -0.0279   0.1443   1.0000
   6.250   0.8421   0.02163   0.01228  -0.0272   0.1095   1.0000
   6.500   0.8649   0.02335   0.01399  -0.0262   0.0850   1.0000
   6.750   0.8875   0.02579   0.01637  -0.0253   0.0706   1.0000
   7.000   0.9112   0.02751   0.01835  -0.0244   0.0603   1.0000
   7.250   0.9341   0.03049   0.02167  -0.0233   0.0549   1.0000
   7.500   0.9561   0.03365   0.02533  -0.0221   0.0519   1.0000
   7.750   0.9761   0.03665   0.02866  -0.0211   0.0490   1.0000
   8.000   0.9900   0.04149   0.03380  -0.0205   0.0460   1.0000
   8.250   1.0026   0.04520   0.03823  -0.0192   0.0451   1.0000
   8.500   1.0105   0.05040   0.04395  -0.0182   0.0454   1.0000
   8.750   1.0149   0.05666   0.05054  -0.0177   0.0461   1.0000
   9.000   0.9974   0.06371   0.05873  -0.0178   0.0502   1.0000
   9.250   0.9760   0.07159   0.06697  -0.0201   0.0528   1.0000
   9.500   0.8752   0.06516   0.06075  -0.0118   0.0517   1.0000
   9.750   0.8411   0.07330   0.06898  -0.0167   0.0537   1.0000
  10.000   0.8198   0.08127   0.07697  -0.0208   0.0558   1.0000
  10.250   0.8065   0.08822   0.08389  -0.0232   0.0570   1.0000
  10.500   0.7977   0.09445   0.09011  -0.0245   0.0578   1.0000
  10.750   0.8597   0.13307   0.12839  -0.0635   0.1162   1.0000
  11.000   0.7308   0.12776   0.12335  -0.0419   0.1231   1.0000
  11.250   0.7287   0.13310   0.12867  -0.0428   0.1220   1.0000
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