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

FX 66-S-161 AIRFOIL (fx66s161-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: FX 66-S-161 AIRFOIL (fx66s161-il)
Reynolds number: 50,000
Max Cl/Cd: 16.16 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx66s161-il-50000-n5.txt
Download as CSV file: xf-fx66s161-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-S-161 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.3317   0.10179   0.09542  -0.0578   1.0000   0.0471
 -10.750  -0.3368   0.09658   0.09031  -0.0600   1.0000   0.0470
 -10.500  -0.3449   0.09121   0.08503  -0.0625   1.0000   0.0469
 -10.250  -0.3545   0.08625   0.08018  -0.0647   1.0000   0.0467
 -10.000  -0.3696   0.08134   0.07540  -0.0669   1.0000   0.0465
  -9.750  -0.3896   0.07740   0.07163  -0.0681   1.0000   0.0462
  -9.500  -0.3969   0.06980   0.06391  -0.0779   0.9579   0.0460
  -9.250  -0.4019   0.06326   0.05709  -0.0871   0.9265   0.0458
  -9.000  -0.4036   0.05778   0.05117  -0.0931   0.9017   0.0461
  -8.750  -0.4009   0.05313   0.04591  -0.0969   0.8826   0.0469
  -8.500  -0.3883   0.04939   0.04180  -0.0988   0.8666   0.0478
  -8.250  -0.3725   0.04635   0.03844  -0.0998   0.8518   0.0489
  -8.000  -0.3561   0.04373   0.03546  -0.1001   0.8378   0.0501
  -7.750  -0.3385   0.04161   0.03302  -0.1000   0.8250   0.0527
  -7.500  -0.3200   0.03953   0.03049  -0.0998   0.8134   0.0560
  -7.250  -0.2984   0.03752   0.02795  -0.0993   0.8034   0.0588
  -7.000  -0.2753   0.03577   0.02612  -0.0988   0.7936   0.0619
  -6.750  -0.2535   0.03454   0.02469  -0.0981   0.7841   0.0673
  -6.500  -0.2291   0.03322   0.02312  -0.0972   0.7764   0.0731
  -6.250  -0.2094   0.03218   0.02205  -0.0961   0.7673   0.0790
  -6.000  -0.1883   0.03113   0.02084  -0.0952   0.7604   0.0882
  -5.750  -0.1706   0.03024   0.01981  -0.0940   0.7519   0.0986
  -5.500  -0.1519   0.02909   0.01869  -0.0933   0.7459   0.1150
  -5.250  -0.1365   0.02797   0.01770  -0.0925   0.7383   0.1412
  -5.000  -0.1208   0.02626   0.01653  -0.0922   0.7325   0.2090
  -4.750  -0.1062   0.02568   0.01691  -0.0904   0.7262   0.3687
  -4.500  -0.0849   0.02628   0.01743  -0.0887   0.7198   0.4476
  -4.250  -0.0609   0.02692   0.01781  -0.0872   0.7152   0.4972
  -4.000  -0.0421   0.02766   0.01848  -0.0850   0.7084   0.5316
  -3.750  -0.0201   0.02820   0.01887  -0.0830   0.7034   0.5610
  -3.500   0.0036   0.02853   0.01904  -0.0813   0.6994   0.5814
  -3.250   0.0235   0.02879   0.01917  -0.0800   0.6933   0.5971
  -3.000   0.0476   0.02881   0.01899  -0.0793   0.6886   0.6108
  -2.750   0.0746   0.02870   0.01862  -0.0793   0.6849   0.6229
  -2.500   0.0948   0.02888   0.01868  -0.0786   0.6790   0.6322
  -2.250   0.1192   0.02889   0.01852  -0.0784   0.6744   0.6410
  -2.000   0.1465   0.02882   0.01821  -0.0787   0.6708   0.6507
  -1.750   0.1675   0.02903   0.01833  -0.0780   0.6658   0.6582
  -1.500   0.1905   0.02920   0.01835  -0.0779   0.6609   0.6675
  -1.250   0.2162   0.02925   0.01827  -0.0777   0.6571   0.6753
  -1.000   0.2421   0.02935   0.01820  -0.0778   0.6536   0.6845
  -0.750   0.2599   0.02982   0.01865  -0.0770   0.6479   0.6934
  -0.500   0.2836   0.03003   0.01877  -0.0767   0.6439   0.7026
   0.000   0.3284   0.03066   0.01927  -0.0759   0.6360   0.7224
   0.250   0.3472   0.03116   0.01976  -0.0753   0.6311   0.7324
   0.500   0.3723   0.03140   0.01992  -0.0752   0.6276   0.7435
   0.750   0.3995   0.03149   0.01995  -0.0751   0.6249   0.7546
   1.000   0.4078   0.03257   0.02112  -0.0735   0.6187   0.7659
   1.250   0.4274   0.03308   0.02164  -0.0729   0.6144   0.7790
   1.500   0.4528   0.03326   0.02182  -0.0726   0.6114   0.7941
   1.750   0.4679   0.03398   0.02260  -0.0714   0.6068   0.8110
   2.000   0.4774   0.03500   0.02372  -0.0699   0.6008   0.8312
   2.250   0.5026   0.03525   0.02404  -0.0697   0.5974   0.8571
   2.500   0.5399   0.03522   0.02406  -0.0712   0.5949   0.8941
   2.750   0.5523   0.03716   0.02612  -0.0725   0.5863   1.0000
   3.000   0.5798   0.03775   0.02660  -0.0736   0.5826   1.0000
   3.250   0.6137   0.03798   0.02671  -0.0749   0.5801   1.0000
   3.750   0.6288   0.04134   0.03001  -0.0733   0.5672   1.0000
   4.000   0.6608   0.04164   0.03023  -0.0740   0.5650   1.0000
   4.500   0.6645   0.04569   0.03428  -0.0712   0.5515   1.0000
   5.000   0.6694   0.05006   0.03864  -0.0692   0.5381   1.0000
   5.250   0.6979   0.05054   0.03910  -0.0694   0.5356   1.0000
   5.500   0.6801   0.05422   0.04280  -0.0678   0.5255   1.0000
   5.750   0.7048   0.05496   0.04355  -0.0678   0.5222   1.0000
   6.250   0.7144   0.05930   0.04795  -0.0665   0.5092   1.0000
   6.500   0.7410   0.05992   0.04861  -0.0665   0.5064   1.0000
   6.750   0.7290   0.06339   0.05212  -0.0655   0.4973   1.0000
   7.000   0.7488   0.06452   0.05331  -0.0653   0.4932   1.0000
   7.250   0.7766   0.06509   0.05395  -0.0653   0.4906   1.0000
   7.500   0.7595   0.06908   0.05798  -0.0645   0.4810   1.0000
   7.750   0.7824   0.07000   0.05898  -0.0644   0.4774   1.0000
   8.250   0.7923   0.07471   0.06382  -0.0637   0.4647   1.0000
   8.500   0.8177   0.07548   0.06471  -0.0636   0.4617   1.0000
   8.750   0.8070   0.07908   0.06838  -0.0632   0.4524   1.0000
   9.000   0.8285   0.08013   0.06954  -0.0631   0.4484   1.0000
   9.250   0.8267   0.08311   0.07260  -0.0629   0.4407   1.0000
   9.500   0.8414   0.08469   0.07430  -0.0627   0.4352   1.0000
   9.750   0.8683   0.08524   0.07501  -0.0625   0.4322   1.0000
  10.000   0.8568   0.08910   0.07895  -0.0624   0.4220   1.0000
  10.250   0.8817   0.08978   0.07979  -0.0622   0.4184   1.0000
  10.500   0.8741   0.09337   0.08347  -0.0622   0.4087   1.0000
  10.750   0.8989   0.09393   0.08419  -0.0619   0.4045   1.0000
  11.000   0.8937   0.09733   0.08770  -0.0620   0.3947   1.0000
  11.250   0.9193   0.09769   0.08827  -0.0615   0.3903   1.0000
  11.500   0.9138   0.10119   0.09188  -0.0618   0.3800   1.0000
  11.750   0.9412   0.10123   0.09213  -0.0612   0.3757   1.0000
<< Back to FX 66-S-161 AIRFOIL (fx66s161-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-S-161 AIRFOIL (fx66s161-il)