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

AG35 (ag35-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AG35 (ag35-il)
Reynolds number: 50,000
Max Cl/Cd: 32.35 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag35-il-50000.txt
Download as CSV file: xf-ag35-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG35                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4836   0.10209   0.09555  -0.0002   1.0000   0.2452
  -9.250  -0.4785   0.09893   0.09244  -0.0002   1.0000   0.2583
  -9.000  -0.4715   0.09538   0.08893   0.0003   1.0000   0.2726
  -8.750  -0.4692   0.09230   0.08593   0.0005   1.0000   0.2875
  -8.500  -0.4712   0.08959   0.08330  -0.0005   1.0000   0.3030
  -8.000  -0.4441   0.08167   0.07545   0.0032   1.0000   0.3382
  -7.750  -0.4402   0.07876   0.07261   0.0037   1.0000   0.3592
  -7.500  -0.4354   0.07555   0.06945   0.0046   1.0000   0.3805
  -7.250  -0.4351   0.07305   0.06704   0.0055   1.0000   0.4078
  -7.000  -0.4266   0.07003   0.06408   0.0084   1.0000   0.4375
  -6.750  -0.4025   0.05486   0.04813  -0.0325   1.0000   0.2138
  -6.500  -0.3722   0.04451   0.03657  -0.0406   1.0000   0.1415
  -6.250  -0.3528   0.04081   0.03265  -0.0403   1.0000   0.1383
  -6.000  -0.3317   0.03736   0.02877  -0.0403   1.0000   0.1372
  -5.750  -0.3089   0.03414   0.02502  -0.0400   1.0000   0.1358
  -5.500  -0.2848   0.03130   0.02167  -0.0394   1.0000   0.1354
  -5.250  -0.2601   0.02909   0.01893  -0.0387   1.0000   0.1407
  -5.000  -0.2367   0.02717   0.01687  -0.0379   1.0000   0.1496
  -4.750  -0.2109   0.02534   0.01469  -0.0370   1.0000   0.1589
  -4.500  -0.1868   0.02391   0.01326  -0.0362   1.0000   0.1785
  -4.250  -0.1613   0.02246   0.01181  -0.0352   1.0000   0.2059
  -4.000  -0.1349   0.02096   0.01058  -0.0345   1.0000   0.2669
  -3.750  -0.0980   0.01634   0.00926  -0.0318   1.0000   1.0000
  -3.500  -0.0761   0.01642   0.00859  -0.0313   1.0000   1.0000
  -3.250  -0.0555   0.01656   0.00825  -0.0305   1.0000   1.0000
  -3.000  -0.0350   0.01674   0.00807  -0.0298   1.0000   1.0000
  -2.750  -0.0146   0.01697   0.00800  -0.0292   1.0000   1.0000
  -2.500   0.0057   0.01725   0.00803  -0.0287   1.0000   1.0000
  -2.250   0.0257   0.01761   0.00815  -0.0284   1.0000   1.0000
  -2.000   0.0451   0.01804   0.00841  -0.0281   1.0000   1.0000
  -1.750   0.0638   0.01858   0.00879  -0.0280   1.0000   1.0000
  -1.500   0.0816   0.01925   0.00932  -0.0280   1.0000   1.0000
  -1.250   0.0987   0.02002   0.00996  -0.0283   1.0000   1.0000
  -1.000   0.1265   0.02092   0.01074  -0.0305   0.9956   1.0000
  -0.750   0.1878   0.02189   0.01157  -0.0385   0.9754   1.0000
  -0.500   0.2485   0.02274   0.01232  -0.0460   0.9563   1.0000
  -0.250   0.2980   0.02343   0.01297  -0.0511   0.9350   1.0000
   0.000   0.3501   0.02400   0.01354  -0.0563   0.9151   1.0000
   0.250   0.4052   0.02441   0.01397  -0.0616   0.8965   1.0000
   0.500   0.4440   0.02487   0.01446  -0.0639   0.8752   1.0000
   0.750   0.4883   0.02510   0.01476  -0.0665   0.8559   1.0000
   1.000   0.5281   0.02528   0.01500  -0.0680   0.8364   1.0000
   1.250   0.5599   0.02555   0.01533  -0.0681   0.8151   1.0000
   1.500   0.5970   0.02549   0.01538  -0.0683   0.7960   1.0000
   1.750   0.6231   0.02580   0.01576  -0.0672   0.7738   1.0000
   2.000   0.6537   0.02580   0.01582  -0.0661   0.7534   1.0000
   2.250   0.6790   0.02605   0.01618  -0.0646   0.7309   1.0000
   2.500   0.7063   0.02606   0.01626  -0.0628   0.7092   1.0000
   2.750   0.7300   0.02632   0.01660  -0.0610   0.6851   1.0000
   3.000   0.7567   0.02623   0.01654  -0.0590   0.6624   1.0000
   3.250   0.7813   0.02629   0.01670  -0.0569   0.6373   1.0000
   3.500   0.8056   0.02646   0.01691  -0.0549   0.6110   1.0000
   3.750   0.8296   0.02682   0.01729  -0.0529   0.5836   1.0000
   4.000   0.8530   0.02735   0.01783  -0.0509   0.5547   1.0000
   4.250   0.8759   0.02798   0.01851  -0.0490   0.5248   1.0000
   4.500   0.8984   0.02863   0.01918  -0.0472   0.4945   1.0000
   4.750   0.9208   0.02922   0.01980  -0.0454   0.4641   1.0000
   5.000   0.9435   0.02967   0.02026  -0.0436   0.4340   1.0000
   5.250   0.9662   0.03007   0.02066  -0.0419   0.4038   1.0000
   5.500   0.9889   0.03057   0.02109  -0.0402   0.3738   1.0000
   5.750   1.0094   0.03143   0.02197  -0.0386   0.3426   1.0000
   6.000   1.0294   0.03245   0.02300  -0.0369   0.3115   1.0000
   6.250   1.0489   0.03370   0.02426  -0.0353   0.2811   1.0000
   6.500   1.0678   0.03520   0.02573  -0.0338   0.2519   1.0000
   6.750   1.0855   0.03693   0.02745  -0.0322   0.2239   1.0000
   7.000   1.1024   0.03888   0.02938  -0.0307   0.1979   1.0000
   7.250   1.1210   0.04095   0.03125  -0.0294   0.1742   1.0000
   7.500   1.1321   0.04384   0.03454  -0.0275   0.1562   1.0000
   7.750   1.1438   0.04670   0.03768  -0.0258   0.1408   1.0000
   8.000   1.1546   0.05006   0.04127  -0.0242   0.1297   1.0000
   8.250   1.1678   0.05312   0.04436  -0.0230   0.1191   1.0000
   8.500   1.1615   0.05788   0.04982  -0.0208   0.1160   1.0000
   8.750   1.1536   0.06256   0.05496  -0.0189   0.1132   1.0000
   9.000   1.1669   0.06580   0.05811  -0.0181   0.1062   1.0000
   9.250   1.1490   0.07076   0.06348  -0.0164   0.1058   1.0000
   9.500   1.1283   0.07573   0.06873  -0.0152   0.1059   1.0000
   9.750   1.1059   0.08077   0.07394  -0.0145   0.1063   1.0000
  10.000   1.0844   0.08646   0.07976  -0.0153   0.1068   1.0000
<< Back to AG35 (ag35-il)

Polar data table (+)

Polar graphs


<< Back to AG35 (ag35-il)