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=200,000 Ncrit=9


Details Polar file
Airfoil: FX 66-S-161 AIRFOIL (fx66s161-il)
Reynolds number: 200,000
Max Cl/Cd: 73.77 at α=10°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66s161-il-200000.txt
Download as CSV file: xf-fx66s161-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.2963   0.11162   0.10847  -0.0482   1.0000   0.0564
 -11.000  -0.3195   0.10517   0.10214  -0.0560   1.0000   0.0585
 -10.750  -0.3298   0.09944   0.09650  -0.0603   1.0000   0.0587
 -10.500  -0.2979   0.09602   0.09309  -0.0585   0.9608   0.0608
 -10.250  -0.2506   0.09069   0.08759  -0.0686   0.9128   0.0639
 -10.000  -0.1790   0.06974   0.06617  -0.0841   0.7865   0.0734
  -9.750  -0.1656   0.06804   0.06436  -0.0827   0.7724   0.0751
  -9.500  -0.1670   0.06400   0.06027  -0.0838   0.7626   0.0768
  -9.250  -0.1782   0.05823   0.05448  -0.0867   0.7549   0.0784
  -9.000  -0.2123   0.04934   0.04558  -0.0932   0.7501   0.0792
  -8.750  -0.3900   0.04165   0.03620  -0.1043   0.7702   0.0374
  -8.500  -0.3771   0.03766   0.03205  -0.1039   0.7602   0.0352
  -8.250  -0.3696   0.03424   0.02825  -0.1028   0.7498   0.0341
  -8.000  -0.3575   0.03122   0.02478  -0.1017   0.7411   0.0334
  -7.750  -0.3411   0.02876   0.02190  -0.1007   0.7326   0.0333
  -7.500  -0.3215   0.02688   0.01968  -0.0999   0.7242   0.0341
  -7.250  -0.2996   0.02562   0.01808  -0.0991   0.7168   0.0356
  -7.000  -0.2769   0.02450   0.01664  -0.0984   0.7092   0.0372
  -6.750  -0.2546   0.02230   0.01433  -0.0978   0.7028   0.0389
  -6.500  -0.2310   0.02124   0.01321  -0.0973   0.6967   0.0411
  -6.250  -0.2071   0.02039   0.01226  -0.0967   0.6903   0.0444
  -6.000  -0.1837   0.01945   0.01119  -0.0961   0.6854   0.0486
  -5.750  -0.1610   0.01864   0.01041  -0.0955   0.6797   0.0539
  -5.500  -0.1375   0.01792   0.00961  -0.0950   0.6745   0.0603
  -5.250  -0.1139   0.01723   0.00887  -0.0946   0.6702   0.0694
  -5.000  -0.0907   0.01650   0.00816  -0.0941   0.6650   0.0844
  -4.750  -0.0690   0.01535   0.00728  -0.0938   0.6602   0.1347
  -4.500  -0.0509   0.01382   0.00679  -0.0934   0.6564   0.3717
  -4.250  -0.0244   0.01396   0.00699  -0.0931   0.6524   0.4382
  -4.000   0.0026   0.01420   0.00715  -0.0928   0.6479   0.4703
  -3.750   0.0298   0.01446   0.00734  -0.0924   0.6439   0.4945
  -3.500   0.0575   0.01471   0.00743  -0.0922   0.6405   0.5138
  -3.250   0.0844   0.01491   0.00759  -0.0919   0.6365   0.5292
  -3.000   0.1112   0.01511   0.00775  -0.0916   0.6324   0.5450
  -2.750   0.1383   0.01528   0.00787  -0.0913   0.6288   0.5597
  -2.500   0.1657   0.01544   0.00798  -0.0910   0.6258   0.5729
  -2.250   0.1928   0.01557   0.00810  -0.0907   0.6227   0.5835
  -2.000   0.2199   0.01563   0.00812  -0.0907   0.6191   0.5930
  -1.750   0.2473   0.01567   0.00812  -0.0907   0.6158   0.5998
  -1.500   0.2758   0.01569   0.00802  -0.0909   0.6128   0.6070
  -1.250   0.3041   0.01573   0.00798  -0.0911   0.6102   0.6124
  -1.000   0.3315   0.01581   0.00803  -0.0912   0.6070   0.6191
  -0.750   0.3586   0.01586   0.00809  -0.0912   0.6038   0.6247
  -0.500   0.3862   0.01593   0.00814  -0.0913   0.6008   0.6312
  -0.250   0.4145   0.01600   0.00814  -0.0916   0.5982   0.6378
   0.000   0.4427   0.01606   0.00817  -0.0917   0.5958   0.6436
   0.250   0.4704   0.01621   0.00827  -0.0919   0.5931   0.6503
   0.500   0.4966   0.01631   0.00845  -0.0919   0.5896   0.6559
   0.750   0.5238   0.01642   0.00857  -0.0919   0.5865   0.6626
   1.000   0.5518   0.01648   0.00860  -0.0921   0.5835   0.6695
   1.250   0.5802   0.01653   0.00862  -0.0922   0.5808   0.6771
   1.500   0.6067   0.01669   0.00882  -0.0922   0.5774   0.6853
   1.750   0.6323   0.01684   0.00906  -0.0920   0.5738   0.6941
   2.000   0.6590   0.01695   0.00922  -0.0920   0.5706   0.7027
   2.250   0.6869   0.01706   0.00933  -0.0922   0.5680   0.7126
   2.500   0.7147   0.01716   0.00946  -0.0923   0.5658   0.7217
   2.750   0.7410   0.01738   0.00975  -0.0922   0.5633   0.7325
   3.000   0.7648   0.01760   0.01013  -0.0918   0.5595   0.7453
   3.250   0.7901   0.01774   0.01038  -0.0915   0.5562   0.7605
   3.500   0.8165   0.01779   0.01052  -0.0913   0.5533   0.7784
   3.750   0.8434   0.01783   0.01063  -0.0911   0.5510   0.8013
   4.000   0.8666   0.01799   0.01097  -0.0903   0.5484   0.8328
   4.250   0.8875   0.01812   0.01142  -0.0891   0.5445   0.8986
   4.500   0.9273   0.01831   0.01167  -0.0921   0.5409   1.0000
   4.750   0.9570   0.01850   0.01183  -0.0928   0.5380   1.0000
   5.000   0.9874   0.01868   0.01194  -0.0935   0.5356   1.0000
   5.250   1.0099   0.01920   0.01256  -0.0931   0.5316   1.0000
   5.500   1.0338   0.01958   0.01302  -0.0928   0.5275   1.0000
   5.750   1.0612   0.01979   0.01325  -0.0929   0.5242   1.0000
   6.000   1.0909   0.01994   0.01337  -0.0934   0.5216   1.0000
   6.250   1.1152   0.02037   0.01388  -0.0931   0.5181   1.0000
   6.500   1.1358   0.02089   0.01455  -0.0924   0.5134   1.0000
   6.750   1.1618   0.02117   0.01489  -0.0923   0.5100   1.0000
   7.000   1.1910   0.02134   0.01507  -0.0927   0.5073   1.0000
   7.250   1.2184   0.02165   0.01542  -0.0928   0.5043   1.0000
   7.500   1.2342   0.02236   0.01635  -0.0914   0.4988   1.0000
   7.750   1.2594   0.02264   0.01672  -0.0912   0.4950   1.0000
   8.000   1.2900   0.02268   0.01678  -0.0917   0.4921   1.0000
   8.250   1.3115   0.02311   0.01735  -0.0909   0.4874   1.0000
   8.500   1.3319   0.02335   0.01774  -0.0899   0.4814   1.0000
   8.750   1.3689   0.02218   0.01648  -0.0906   0.4740   1.0000
   9.000   1.3936   0.02067   0.01493  -0.0892   0.4591   1.0000
   9.250   1.4126   0.01984   0.01412  -0.0872   0.4448   1.0000
   9.500   1.4296   0.01961   0.01399  -0.0853   0.4331   1.0000
   9.750   1.4428   0.01967   0.01421  -0.0830   0.4208   1.0000
  10.000   1.4555   0.01973   0.01436  -0.0805   0.4064   1.0000
  10.250   1.4639   0.01988   0.01454  -0.0773   0.3886   1.0000
  10.500   1.4633   0.02037   0.01505  -0.0729   0.3674   1.0000
  10.750   1.4564   0.02128   0.01588  -0.0681   0.3397   1.0000
  11.000   1.4410   0.02292   0.01736  -0.0631   0.3083   1.0000
  11.250   1.4197   0.02545   0.01973  -0.0587   0.2719   1.0000
  11.500   1.3931   0.02900   0.02308  -0.0553   0.2353   1.0000
  11.750   1.3665   0.03320   0.02710  -0.0529   0.2031   1.0000
  12.000   1.3386   0.03792   0.03165  -0.0511   0.1715   1.0000
  12.250   1.3127   0.04274   0.03631  -0.0498   0.1428   1.0000
  12.500   1.2861   0.04786   0.04125  -0.0489   0.1132   1.0000
  12.750   1.2606   0.05315   0.04634  -0.0483   0.0849   1.0000
  13.000   1.2385   0.05839   0.05141  -0.0480   0.0638   1.0000
  13.250   1.2214   0.06334   0.05626  -0.0480   0.0515   1.0000
  13.500   1.2110   0.06766   0.06059  -0.0481   0.0439   1.0000
  13.750   1.2025   0.07191   0.06489  -0.0483   0.0393   1.0000
  14.000   1.1978   0.07575   0.06878  -0.0487   0.0354   1.0000
  14.250   1.1902   0.07999   0.07306  -0.0491   0.0333   1.0000
  14.500   1.1885   0.08354   0.07669  -0.0494   0.0308   1.0000
  14.750   1.1891   0.08680   0.08004  -0.0498   0.0285   1.0000
  15.000   1.1891   0.09012   0.08339  -0.0502   0.0269   1.0000
  15.250   1.1894   0.09298   0.08620  -0.0499   0.0252   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)