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

AG11 (ag11-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: AG11 (ag11-il)
Reynolds number: 1,000,000
Max Cl/Cd: 90.99 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag11-il-1000000-n5.txt
Download as CSV file: xf-ag11-il-1000000-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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4254   0.08571   0.08422   0.0061   1.0000   0.0046
  -8.000  -0.4237   0.08206   0.08059   0.0050   1.0000   0.0047
  -7.750  -0.4222   0.07841   0.07695   0.0039   1.0000   0.0049
  -7.500  -0.4211   0.07468   0.07323   0.0027   1.0000   0.0051
  -7.250  -0.4210   0.07081   0.06937   0.0012   1.0000   0.0052
  -7.000  -0.4211   0.06704   0.06562  -0.0005   1.0000   0.0054
  -6.750  -0.4187   0.06274   0.06133  -0.0039   1.0000   0.0058
  -6.500  -0.4111   0.05772   0.05630  -0.0088   1.0000   0.0062
  -6.250  -0.3997   0.05228   0.05083  -0.0144   1.0000   0.0064
  -6.000  -0.3847   0.04688   0.04538  -0.0195   1.0000   0.0065
  -5.750  -0.3688   0.04197   0.04041  -0.0230   1.0000   0.0065
  -5.500  -0.3517   0.03726   0.03563  -0.0259   1.0000   0.0065
  -5.250  -0.3303   0.03235   0.03062  -0.0291   0.9948   0.0065
  -5.000  -0.3035   0.02733   0.02548  -0.0327   0.9809   0.0065
  -4.750  -0.2760   0.02272   0.02071  -0.0356   0.9633   0.0066
  -4.500  -0.2513   0.01874   0.01653  -0.0370   0.9431   0.0066
  -4.250  -0.2479   0.03197   0.02944  -0.0375   0.9730   0.0066
  -3.250  -0.1391   0.01542   0.01129  -0.0373   0.8784   0.0040
  -3.000  -0.1122   0.01225   0.00754  -0.0363   0.8514   0.0039
  -2.750  -0.0853   0.01049   0.00534  -0.0356   0.8219   0.0039
  -2.500  -0.0584   0.00955   0.00411  -0.0350   0.7906   0.0039
  -2.250  -0.0313   0.00899   0.00331  -0.0346   0.7592   0.0040
  -2.000  -0.0038   0.00865   0.00278  -0.0344   0.7294   0.0042
  -1.750   0.0237   0.00818   0.00214  -0.0342   0.6998   0.0048
  -1.500   0.0515   0.00806   0.00189  -0.0341   0.6706   0.0054
  -1.250   0.0794   0.00795   0.00165  -0.0340   0.6401   0.0060
  -1.000   0.1074   0.00788   0.00145  -0.0339   0.6111   0.0064
  -0.750   0.1354   0.00779   0.00125  -0.0339   0.5816   0.0076
  -0.500   0.1634   0.00777   0.00112  -0.0338   0.5526   0.0089
  -0.250   0.1915   0.00774   0.00098  -0.0338   0.5238   0.0122
   0.000   0.2195   0.00773   0.00090  -0.0338   0.4920   0.0273
   0.250   0.2474   0.00760   0.00088  -0.0339   0.4615   0.1035
   0.500   0.2750   0.00697   0.00088  -0.0343   0.4341   0.3846
   0.750   0.2984   0.00577   0.00095  -0.0338   0.4060   0.8382
   1.000   0.3254   0.00557   0.00093  -0.0332   0.3791   1.0000
   1.250   0.3533   0.00573   0.00096  -0.0332   0.3543   1.0000
   1.500   0.3811   0.00591   0.00099  -0.0332   0.3277   1.0000
   1.750   0.4089   0.00609   0.00104  -0.0332   0.3016   1.0000
   2.000   0.4367   0.00627   0.00110  -0.0332   0.2778   1.0000
   2.250   0.4645   0.00646   0.00117  -0.0331   0.2535   1.0000
   2.500   0.4922   0.00666   0.00126  -0.0331   0.2304   1.0000
   2.750   0.5199   0.00685   0.00135  -0.0331   0.2103   1.0000
   3.000   0.5475   0.00706   0.00145  -0.0331   0.1897   1.0000
   3.250   0.5751   0.00726   0.00156  -0.0331   0.1724   1.0000
   3.500   0.6028   0.00744   0.00168  -0.0331   0.1582   1.0000
   3.750   0.6303   0.00764   0.00182  -0.0330   0.1447   1.0000
   4.000   0.6578   0.00786   0.00196  -0.0330   0.1290   1.0000
   4.250   0.6852   0.00809   0.00212  -0.0330   0.1148   1.0000
   4.500   0.7126   0.00831   0.00228  -0.0329   0.1041   1.0000
   4.750   0.7399   0.00853   0.00248  -0.0329   0.0941   1.0000
   5.000   0.7671   0.00878   0.00267  -0.0328   0.0839   1.0000
   5.250   0.7943   0.00903   0.00288  -0.0328   0.0744   1.0000
   5.500   0.8214   0.00925   0.00308  -0.0327   0.0678   1.0000
   5.750   0.8484   0.00952   0.00333  -0.0326   0.0613   1.0000
   6.000   0.8755   0.00974   0.00356  -0.0326   0.0571   1.0000
   6.250   0.9024   0.01002   0.00381  -0.0325   0.0519   1.0000
   6.500   0.9292   0.01029   0.00410  -0.0324   0.0476   1.0000
   6.750   0.9560   0.01052   0.00436  -0.0323   0.0454   1.0000
   7.000   0.9827   0.01080   0.00465  -0.0322   0.0425   1.0000
   7.250   1.0091   0.01113   0.00498  -0.0320   0.0391   1.0000
   7.500   1.0355   0.01142   0.00531  -0.0319   0.0368   1.0000
   7.750   1.0620   0.01169   0.00563  -0.0318   0.0351   1.0000
   8.000   1.0881   0.01202   0.00598  -0.0316   0.0329   1.0000
   8.250   1.1139   0.01241   0.00637  -0.0314   0.0300   1.0000
   8.500   1.1396   0.01279   0.00679  -0.0312   0.0279   1.0000
   8.750   1.1655   0.01311   0.00719  -0.0311   0.0268   1.0000
   9.000   1.1910   0.01348   0.00759  -0.0309   0.0249   1.0000
   9.250   1.2162   0.01391   0.00806  -0.0306   0.0230   1.0000
   9.500   1.2408   0.01442   0.00860  -0.0303   0.0205   1.0000
   9.750   1.2660   0.01479   0.00902  -0.0301   0.0189   1.0000
  10.000   1.2903   0.01530   0.00953  -0.0299   0.0163   1.0000
  10.250   1.3142   0.01585   0.01014  -0.0295   0.0143   1.0000
  10.500   1.3383   0.01635   0.01071  -0.0292   0.0131   1.0000
  10.750   1.3616   0.01695   0.01134  -0.0288   0.0114   1.0000
  11.000   1.3845   0.01760   0.01205  -0.0284   0.0098   1.0000
  11.250   1.4071   0.01825   0.01277  -0.0279   0.0085   1.0000
  11.500   1.4285   0.01906   0.01363  -0.0274   0.0070   1.0000
  11.750   1.4501   0.01981   0.01446  -0.0268   0.0060   1.0000
  12.000   1.4702   0.02072   0.01547  -0.0261   0.0048   1.0000
  12.250   1.4899   0.02164   0.01649  -0.0254   0.0040   1.0000
  12.500   1.5085   0.02266   0.01761  -0.0246   0.0033   1.0000
  12.750   1.5256   0.02381   0.01887  -0.0237   0.0026   1.0000
  13.000   1.5424   0.02491   0.02009  -0.0227   0.0022   1.0000
  13.250   1.5570   0.02619   0.02150  -0.0216   0.0019   1.0000
  13.500   1.5691   0.02767   0.02311  -0.0203   0.0016   1.0000
  13.750   1.5805   0.02908   0.02467  -0.0190   0.0014   1.0000
  14.000   1.5898   0.03054   0.02628  -0.0175   0.0013   1.0000
  14.250   1.5939   0.03213   0.02802  -0.0154   0.0013   1.0000
  14.500   1.5924   0.03398   0.03002  -0.0131   0.0012   1.0000
  14.750   1.5884   0.03634   0.03253  -0.0116   0.0012   1.0000
  15.000   1.5817   0.03942   0.03578  -0.0113   0.0011   1.0000
  15.250   1.5730   0.04338   0.03990  -0.0124   0.0011   1.0000
  15.500   1.5600   0.04872   0.04542  -0.0152   0.0010   1.0000
  15.750   1.5420   0.05573   0.05263  -0.0197   0.0010   1.0000
  16.000   1.5160   0.06490   0.06200  -0.0258   0.0011   1.0000
  16.250   1.4765   0.07698   0.07428  -0.0332   0.0011   1.0000
  16.500   1.4238   0.09144   0.08894  -0.0411   0.0012   1.0000
  16.750   1.3580   0.10841   0.10607  -0.0496   0.0014   1.0000
<< Back to AG11 (ag11-il)

Polar data table (+)

Polar graphs


<< Back to AG11 (ag11-il)