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AG11 (ag11-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG11 (ag11-il)
Reynolds number: 100,000
Max Cl/Cd: 46.13 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag11-il-100000-n5.txt
Download as CSV file: xf-ag11-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG11                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5155   0.10226   0.09741   0.0123   1.0000   0.0349
  -7.750  -0.5105   0.09885   0.09404   0.0103   1.0000   0.0358
  -7.500  -0.5056   0.09544   0.09067   0.0081   1.0000   0.0366
  -7.250  -0.5001   0.09194   0.08722   0.0051   1.0000   0.0376
  -7.000  -0.4904   0.08810   0.08341   0.0002   1.0000   0.0388
  -6.750  -0.4744   0.08411   0.07940  -0.0085   1.0000   0.0403
  -6.500  -0.4540   0.08019   0.07536  -0.0167   1.0000   0.0409
  -6.250  -0.4350   0.07629   0.07129  -0.0216   1.0000   0.0412
  -6.000  -0.4222   0.07094   0.06595  -0.0235   1.0000   0.0417
  -5.750  -0.4134   0.06628   0.06138  -0.0227   1.0000   0.0427
  -5.500  -0.3990   0.06260   0.05771  -0.0233   1.0000   0.0439
  -5.250  -0.3810   0.05899   0.05404  -0.0250   1.0000   0.0455
  -5.000  -0.3601   0.05533   0.05028  -0.0272   1.0000   0.0474
  -4.750  -0.3239   0.05292   0.04730  -0.0321   1.0000   0.0541
  -4.500  -0.3023   0.04851   0.04270  -0.0336   1.0000   0.0550
  -4.250  -0.2835   0.04427   0.03847  -0.0341   1.0000   0.0560
  -4.000  -0.2557   0.03888   0.03281  -0.0340   1.0000   0.0266
  -3.750  -0.2296   0.03516   0.02877  -0.0346   1.0000   0.0236
  -3.500  -0.1977   0.03191   0.02488  -0.0342   1.0000   0.0204
  -3.000  -0.1496   0.02680   0.01934  -0.0344   1.0000   0.0229
  -2.750  -0.1239   0.02489   0.01715  -0.0339   1.0000   0.0244
  -2.500  -0.0974   0.02281   0.01470  -0.0332   1.0000   0.0241
  -2.250  -0.0711   0.02098   0.01250  -0.0324   1.0000   0.0239
  -2.000  -0.0451   0.01942   0.01068  -0.0316   1.0000   0.0241
  -1.750  -0.0196   0.01811   0.00917  -0.0307   1.0000   0.0247
  -1.500   0.0056   0.01698   0.00794  -0.0299   1.0000   0.0256
  -1.250   0.0304   0.01612   0.00701  -0.0292   1.0000   0.0279
  -1.000   0.0574   0.01523   0.00612  -0.0291   0.9982   0.0325
  -0.750   0.0995   0.01439   0.00519  -0.0319   0.9856   0.0392
  -0.500   0.1411   0.01359   0.00439  -0.0345   0.9714   0.0678
  -0.250   0.1780   0.01054   0.00401  -0.0358   0.9579   1.0000
   0.000   0.2169   0.01056   0.00374  -0.0378   0.9352   1.0000
   0.250   0.2536   0.01059   0.00355  -0.0392   0.9096   1.0000
   0.500   0.2865   0.01063   0.00341  -0.0397   0.8803   1.0000
   0.750   0.3162   0.01070   0.00329  -0.0395   0.8480   1.0000
   1.000   0.3436   0.01079   0.00321  -0.0387   0.8141   1.0000
   1.250   0.3696   0.01092   0.00316  -0.0377   0.7780   1.0000
   1.500   0.3950   0.01108   0.00313  -0.0367   0.7416   1.0000
   1.750   0.4204   0.01127   0.00314  -0.0356   0.7040   1.0000
   2.000   0.4460   0.01149   0.00317  -0.0348   0.6658   1.0000
   2.250   0.4715   0.01174   0.00324  -0.0339   0.6269   1.0000
   2.500   0.4969   0.01203   0.00335  -0.0331   0.5873   1.0000
   2.750   0.5224   0.01233   0.00347  -0.0324   0.5470   1.0000
   3.000   0.5479   0.01267   0.00362  -0.0318   0.5066   1.0000
   3.250   0.5734   0.01303   0.00381  -0.0312   0.4670   1.0000
   3.500   0.5988   0.01342   0.00403  -0.0307   0.4283   1.0000
   3.750   0.6243   0.01384   0.00432  -0.0302   0.3906   1.0000
   4.000   0.6497   0.01428   0.00461  -0.0298   0.3544   1.0000
   4.250   0.6751   0.01475   0.00494  -0.0294   0.3202   1.0000
   4.500   0.7004   0.01524   0.00531  -0.0291   0.2886   1.0000
   4.750   0.7258   0.01574   0.00577  -0.0288   0.2581   1.0000
   5.000   0.7510   0.01628   0.00622  -0.0285   0.2305   1.0000
   5.250   0.7760   0.01685   0.00672  -0.0282   0.2052   1.0000
   5.500   0.8011   0.01743   0.00727  -0.0278   0.1824   1.0000
   5.750   0.8258   0.01808   0.00790  -0.0275   0.1617   1.0000
   6.000   0.8505   0.01875   0.00858  -0.0272   0.1440   1.0000
   6.250   0.8748   0.01949   0.00934  -0.0268   0.1300   1.0000
   6.500   0.8988   0.02027   0.01014  -0.0263   0.1173   1.0000
   6.750   0.9225   0.02109   0.01102  -0.0259   0.1063   1.0000
   7.000   0.9458   0.02198   0.01196  -0.0254   0.0980   1.0000
   7.250   0.9692   0.02295   0.01307  -0.0249   0.0911   1.0000
   7.500   0.9917   0.02396   0.01410  -0.0243   0.0848   1.0000
   7.750   1.0148   0.02495   0.01530  -0.0238   0.0783   1.0000
   8.000   1.0370   0.02605   0.01652  -0.0232   0.0737   1.0000
   8.250   1.0586   0.02746   0.01805  -0.0225   0.0701   1.0000
   8.500   1.0806   0.02877   0.01964  -0.0219   0.0657   1.0000
   8.750   1.1015   0.02994   0.02092  -0.0213   0.0615   1.0000
   9.000   1.1213   0.03164   0.02280  -0.0206   0.0584   1.0000
   9.250   1.1409   0.03355   0.02515  -0.0198   0.0556   1.0000
   9.500   1.1593   0.03539   0.02729  -0.0190   0.0526   1.0000
   9.750   1.1769   0.03685   0.02887  -0.0183   0.0496   1.0000
  10.000   1.1916   0.03916   0.03143  -0.0174   0.0472   1.0000
  10.250   1.2028   0.04207   0.03488  -0.0162   0.0452   1.0000
  10.500   1.2105   0.04509   0.03838  -0.0151   0.0431   1.0000
  10.750   1.2174   0.04757   0.04119  -0.0140   0.0410   1.0000
  11.000   1.2262   0.04930   0.04304  -0.0131   0.0389   1.0000
  11.250   1.2292   0.05218   0.04602  -0.0121   0.0375   1.0000
  11.500   1.2173   0.05618   0.05044  -0.0107   0.0370   1.0000
  11.750   1.1988   0.06045   0.05504  -0.0099   0.0367   1.0000
  12.000   1.1777   0.06571   0.06058  -0.0115   0.0366   1.0000
  12.250   1.1547   0.07235   0.06745  -0.0156   0.0367   1.0000
  12.500   1.1299   0.08060   0.07587  -0.0217   0.0369   1.0000
  12.750   1.1027   0.09051   0.08590  -0.0292   0.0373   1.0000
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