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

AG46ct -02f rot. (ag46ct02r-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG46ct -02f rot. (ag46ct02r-il)
Reynolds number: 100,000
Max Cl/Cd: 44.37 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag46ct02r-il-100000-n5.txt
Download as CSV file: xf-ag46ct02r-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG46ct  -02f rot.                               
  
 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  -0.5703   0.09893   0.09419   0.0169   1.0000   0.0423
  -7.750  -0.5669   0.09500   0.09031   0.0142   1.0000   0.0427
  -7.500  -0.5635   0.09074   0.08611   0.0075   1.0000   0.0446
  -7.000  -0.4751   0.06430   0.05986  -0.0087   1.0000   0.0229
  -6.750  -0.5300   0.07092   0.06618  -0.0096   1.0000   0.0226
  -6.500  -0.5182   0.06625   0.06148  -0.0124   1.0000   0.0219
  -6.250  -0.5029   0.06103   0.05613  -0.0163   1.0000   0.0213
  -6.000  -0.4845   0.05540   0.05032  -0.0204   1.0000   0.0212
  -5.750  -0.4636   0.04959   0.04425  -0.0241   1.0000   0.0215
  -5.500  -0.4410   0.04412   0.03841  -0.0268   1.0000   0.0221
  -5.250  -0.4176   0.03929   0.03313  -0.0284   1.0000   0.0225
  -5.000  -0.3941   0.03497   0.02835  -0.0295   1.0000   0.0224
  -4.750  -0.3694   0.03110   0.02398  -0.0300   1.0000   0.0223
  -4.500  -0.3436   0.02778   0.02011  -0.0301   1.0000   0.0224
  -4.250  -0.3170   0.02503   0.01683  -0.0299   1.0000   0.0227
  -4.000  -0.2900   0.02281   0.01415  -0.0296   1.0000   0.0231
  -3.750  -0.2637   0.02044   0.01147  -0.0292   1.0000   0.0241
  -3.500  -0.2378   0.01912   0.01007  -0.0289   1.0000   0.0275
  -3.250  -0.2112   0.01795   0.00872  -0.0284   1.0000   0.0314
  -3.000  -0.1848   0.01669   0.00733  -0.0278   1.0000   0.0341
  -2.750  -0.1589   0.01561   0.00626  -0.0274   1.0000   0.0417
  -2.500  -0.1326   0.01469   0.00531  -0.0270   1.0000   0.0535
  -2.250  -0.1063   0.01381   0.00452  -0.0266   1.0000   0.0857
  -2.000  -0.0808   0.01279   0.00403  -0.0264   1.0000   0.1869
  -1.500  -0.0390   0.01032   0.00369  -0.0235   1.0000   0.7242
  -1.250  -0.0021   0.00977   0.00338  -0.0241   1.0000   1.0000
  -1.000   0.0234   0.00978   0.00320  -0.0237   1.0000   1.0000
  -0.750   0.0487   0.00980   0.00308  -0.0233   1.0000   1.0000
  -0.500   0.0817   0.00985   0.00299  -0.0246   0.9878   1.0000
  -0.250   0.1245   0.00991   0.00289  -0.0277   0.9504   1.0000
   0.000   0.1657   0.00998   0.00281  -0.0302   0.9038   1.0000
   0.250   0.2027   0.01008   0.00272  -0.0316   0.8532   1.0000
   0.500   0.2330   0.01023   0.00265  -0.0315   0.8074   1.0000
   0.750   0.2599   0.01044   0.00263  -0.0307   0.7704   1.0000
   1.000   0.2861   0.01066   0.00266  -0.0299   0.7397   1.0000
   1.250   0.3123   0.01091   0.00273  -0.0292   0.7135   1.0000
   1.500   0.3386   0.01114   0.00286  -0.0286   0.6880   1.0000
   1.750   0.3650   0.01135   0.00298  -0.0281   0.6612   1.0000
   2.000   0.3913   0.01154   0.00309  -0.0275   0.6338   1.0000
   2.250   0.4175   0.01174   0.00320  -0.0269   0.6056   1.0000
   2.500   0.4436   0.01195   0.00332  -0.0263   0.5761   1.0000
   2.750   0.4697   0.01219   0.00349  -0.0258   0.5454   1.0000
   3.000   0.4957   0.01244   0.00364  -0.0252   0.5134   1.0000
   3.250   0.5217   0.01272   0.00383  -0.0247   0.4796   1.0000
   3.500   0.5476   0.01303   0.00404  -0.0242   0.4441   1.0000
   3.750   0.5734   0.01337   0.00433  -0.0237   0.4071   1.0000
   4.000   0.5991   0.01377   0.00462  -0.0233   0.3698   1.0000
   4.250   0.6247   0.01421   0.00496  -0.0229   0.3317   1.0000
   4.500   0.6503   0.01469   0.00535  -0.0225   0.2948   1.0000
   4.750   0.6756   0.01523   0.00583  -0.0222   0.2593   1.0000
   5.000   0.7010   0.01580   0.00634  -0.0219   0.2268   1.0000
   5.250   0.7262   0.01642   0.00690  -0.0216   0.1980   1.0000
   5.500   0.7512   0.01708   0.00753  -0.0213   0.1731   1.0000
   5.750   0.7760   0.01780   0.00822  -0.0209   0.1517   1.0000
   6.000   0.8009   0.01852   0.00902  -0.0206   0.1325   1.0000
   6.250   0.8254   0.01930   0.00984  -0.0202   0.1153   1.0000
   6.500   0.8496   0.02016   0.01075  -0.0198   0.1008   1.0000
   6.750   0.8735   0.02106   0.01170  -0.0195   0.0867   1.0000
   7.000   0.8971   0.02208   0.01283  -0.0190   0.0754   1.0000
   7.250   0.9203   0.02312   0.01403  -0.0185   0.0642   1.0000
   7.500   0.9425   0.02445   0.01548  -0.0179   0.0563   1.0000
   7.750   0.9643   0.02573   0.01684  -0.0175   0.0485   1.0000
   8.000   0.9861   0.02723   0.01861  -0.0167   0.0422   1.0000
   8.250   1.0062   0.02888   0.02037  -0.0161   0.0378   1.0000
   8.500   1.0262   0.03081   0.02262  -0.0154   0.0333   1.0000
   8.750   1.0452   0.03274   0.02482  -0.0147   0.0299   1.0000
   9.000   1.0620   0.03501   0.02732  -0.0140   0.0278   1.0000
   9.250   1.0756   0.03820   0.03087  -0.0131   0.0260   1.0000
   9.500   1.0882   0.04119   0.03439  -0.0124   0.0238   1.0000
   9.750   1.0971   0.04436   0.03801  -0.0118   0.0220   1.0000
  10.000   1.1010   0.04812   0.04217  -0.0113   0.0211   1.0000
  10.250   1.0995   0.05214   0.04657  -0.0110   0.0206   1.0000
  10.500   1.0913   0.05647   0.05123  -0.0113   0.0202   1.0000
  10.750   1.0748   0.06138   0.05641  -0.0127   0.0202   1.0000
  11.000   1.0523   0.06866   0.06396  -0.0184   0.0205   1.0000
  11.250   1.0215   0.07978   0.07531  -0.0282   0.0213   1.0000
  11.500   0.9780   0.09646   0.09208  -0.0409   0.0231   1.0000
<< Back to AG46ct -02f rot. (ag46ct02r-il)

Polar data table (+)

Polar graphs


<< Back to AG46ct -02f rot. (ag46ct02r-il)