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=100,000 Ncrit=9


Details Polar file
Airfoil: AG11 (ag11-il)
Reynolds number: 100,000
Max Cl/Cd: 46.65 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag11-il-100000.txt
Download as CSV file: xf-ag11-il-100000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5299   0.11135   0.10648   0.0139   1.0000   0.0528
  -8.250  -0.5290   0.10906   0.10425   0.0094   1.0000   0.0536
  -8.000  -0.5279   0.10663   0.10189   0.0038   1.0000   0.0540
  -7.750  -0.5197   0.10340   0.09867  -0.0042   1.0000   0.0543
  -7.500  -0.5120   0.09797   0.09328  -0.0061   1.0000   0.0548
  -7.250  -0.5046   0.09273   0.08809   0.0028   1.0000   0.0570
  -7.000  -0.4952   0.08904   0.08440   0.0019   1.0000   0.0591
  -6.750  -0.4848   0.08530   0.08068  -0.0011   1.0000   0.0614
  -6.500  -0.4721   0.08141   0.07679  -0.0058   1.0000   0.0642
  -6.250  -0.4465   0.07807   0.07323  -0.0194   1.0000   0.0678
  -6.000  -0.4335   0.07276   0.06792  -0.0218   1.0000   0.0690
  -5.750  -0.4253   0.06866   0.06396  -0.0186   1.0000   0.0714
  -5.500  -0.4094   0.06515   0.06042  -0.0198   1.0000   0.0755
  -5.250  -0.3762   0.06189   0.05668  -0.0294   1.0000   0.0827
  -5.000  -0.3667   0.05707   0.05212  -0.0272   1.0000   0.0854
  -4.750  -0.3465   0.05392   0.04890  -0.0284   1.0000   0.0925
  -4.500  -0.3239   0.05009   0.04492  -0.0308   1.0000   0.0999
  -4.250  -0.2986   0.04708   0.04163  -0.0332   1.0000   0.1123
  -4.000  -0.2756   0.04428   0.03866  -0.0343   1.0000   0.1262
  -3.750  -0.2550   0.04118   0.03552  -0.0346   1.0000   0.1411
  -3.500  -0.2353   0.03822   0.03261  -0.0343   1.0000   0.1580
  -3.250  -0.2157   0.03578   0.03017  -0.0340   1.0000   0.1883
  -3.000  -0.1985   0.03362   0.02807  -0.0330   1.0000   0.2344
  -2.750  -0.1840   0.03136   0.02598  -0.0310   1.0000   0.2892
  -2.500  -0.1664   0.02917   0.02387  -0.0293   1.0000   0.3421
  -2.250  -0.0841   0.02461   0.01693  -0.0348   1.0000   0.0986
  -2.000  -0.0521   0.02265   0.01425  -0.0329   1.0000   0.0687
  -1.750  -0.0263   0.02028   0.01179  -0.0322   1.0000   0.0637
  -1.500   0.0003   0.01864   0.00991  -0.0312   1.0000   0.0602
  -1.250   0.0264   0.01734   0.00846  -0.0302   1.0000   0.0595
  -1.000   0.0515   0.01644   0.00750  -0.0294   1.0000   0.0653
  -0.750   0.0770   0.01535   0.00645  -0.0287   1.0000   0.0703
  -0.500   0.1023   0.01449   0.00567  -0.0282   1.0000   0.0836
  -0.250   0.1310   0.01076   0.00465  -0.0273   1.0000   1.0000
   0.000   0.1549   0.01097   0.00447  -0.0268   1.0000   1.0000
   0.250   0.1781   0.01124   0.00454  -0.0266   1.0000   1.0000
   0.500   0.2006   0.01158   0.00475  -0.0265   1.0000   1.0000
   0.750   0.2543   0.01181   0.00483  -0.0323   0.9849   1.0000
   1.000   0.3146   0.01185   0.00478  -0.0388   0.9631   1.0000
   1.250   0.3694   0.01177   0.00467  -0.0438   0.9392   1.0000
   1.500   0.4117   0.01169   0.00456  -0.0458   0.9087   1.0000
   1.750   0.4441   0.01165   0.00447  -0.0455   0.8743   1.0000
   2.000   0.4698   0.01168   0.00442  -0.0438   0.8365   1.0000
   2.250   0.4932   0.01177   0.00439  -0.0417   0.7972   1.0000
   2.500   0.5159   0.01192   0.00441  -0.0395   0.7558   1.0000
   2.750   0.5388   0.01215   0.00446  -0.0376   0.7119   1.0000
   3.000   0.5621   0.01243   0.00456  -0.0359   0.6661   1.0000
   3.250   0.5858   0.01276   0.00469  -0.0344   0.6192   1.0000
   3.500   0.6095   0.01316   0.00485  -0.0331   0.5722   1.0000
   3.750   0.6337   0.01360   0.00512  -0.0320   0.5231   1.0000
   4.000   0.6578   0.01410   0.00538  -0.0310   0.4759   1.0000
   4.250   0.6820   0.01466   0.00570  -0.0302   0.4307   1.0000
   4.500   0.7063   0.01527   0.00609  -0.0294   0.3873   1.0000
   4.750   0.7305   0.01593   0.00661  -0.0288   0.3460   1.0000
   5.000   0.7549   0.01663   0.00716  -0.0282   0.3070   1.0000
   5.250   0.7791   0.01743   0.00778  -0.0277   0.2724   1.0000
   5.500   0.8033   0.01827   0.00851  -0.0271   0.2401   1.0000
   5.750   0.8275   0.01917   0.00934  -0.0266   0.2121   1.0000
   6.000   0.8516   0.02028   0.01027  -0.0261   0.1906   1.0000
   6.250   0.8762   0.02133   0.01139  -0.0255   0.1712   1.0000
   6.500   0.9005   0.02240   0.01245  -0.0250   0.1553   1.0000
   6.750   0.9252   0.02373   0.01387  -0.0245   0.1435   1.0000
   7.000   0.9499   0.02525   0.01539  -0.0240   0.1343   1.0000
   7.250   0.9742   0.02652   0.01673  -0.0236   0.1248   1.0000
   7.500   0.9979   0.02816   0.01869  -0.0230   0.1167   1.0000
   7.750   1.0219   0.02998   0.02053  -0.0226   0.1110   1.0000
   8.000   1.0436   0.03233   0.02344  -0.0217   0.1060   1.0000
   8.250   1.0657   0.03395   0.02524  -0.0211   0.0997   1.0000
   8.500   1.0849   0.03683   0.02843  -0.0205   0.0956   1.0000
   8.750   1.1002   0.04026   0.03257  -0.0193   0.0927   1.0000
   9.000   1.1129   0.04401   0.03688  -0.0183   0.0901   1.0000
   9.250   1.1332   0.04585   0.03869  -0.0178   0.0856   1.0000
   9.500   1.1387   0.05053   0.04385  -0.0169   0.0832   1.0000
   9.750   1.1367   0.05593   0.04988  -0.0159   0.0826   1.0000
  10.000   1.1294   0.06181   0.05623  -0.0153   0.0826   1.0000
  10.250   1.1178   0.06784   0.06260  -0.0152   0.0830   1.0000
  10.500   1.1043   0.07388   0.06886  -0.0155   0.0835   1.0000
  10.750   1.0918   0.07990   0.07500  -0.0160   0.0840   1.0000
<< Back to AG11 (ag11-il)

Polar data table (+)

Polar graphs


<< Back to AG11 (ag11-il)