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FX 61-140 AIRFOIL (fx61140-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 61-140 AIRFOIL (fx61140-il)
Reynolds number: 50,000
Max Cl/Cd: 14.4 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx61140-il-50000.txt
Download as CSV file: xf-fx61140-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 61-140 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4120   0.08697   0.08094  -0.0341   1.0000   0.2177
  -7.750  -0.4683   0.08017   0.07443  -0.0390   1.0000   0.1948
  -7.500  -0.5202   0.07550   0.06988  -0.0429   1.0000   0.1930
  -7.000  -0.5570   0.06168   0.05534  -0.0488   1.0000   0.1355
  -6.750  -0.5533   0.05663   0.04987  -0.0492   1.0000   0.1210
  -6.500  -0.5426   0.05244   0.04560  -0.0482   1.0000   0.1146
  -6.250  -0.5315   0.04837   0.04078  -0.0488   1.0000   0.1057
  -6.000  -0.5165   0.04487   0.03702  -0.0485   1.0000   0.1020
  -5.750  -0.4982   0.04166   0.03334  -0.0484   1.0000   0.0992
  -5.500  -0.4788   0.03914   0.03045  -0.0481   1.0000   0.0999
  -5.250  -0.4585   0.03696   0.02792  -0.0476   1.0000   0.1024
  -5.000  -0.4373   0.03504   0.02566  -0.0468   1.0000   0.1038
  -4.750  -0.4170   0.03344   0.02378  -0.0454   1.0000   0.1054
  -4.500  -0.3992   0.03218   0.02235  -0.0435   1.0000   0.1085
  -4.250  -0.3857   0.03116   0.02149  -0.0411   1.0000   0.1171
  -4.000  -0.3732   0.03040   0.02079  -0.0382   1.0000   0.1260
  -3.750  -0.3602   0.02980   0.02013  -0.0356   1.0000   0.1363
  -3.500  -0.3452   0.02877   0.01939  -0.0345   1.0000   0.1594
  -3.250  -0.3241   0.02518   0.01886  -0.0350   1.0000   0.4761
  -3.000  -0.3465   0.02891   0.02286  -0.0204   1.0000   0.6706
  -2.750  -0.3492   0.03293   0.02670  -0.0071   0.9848   0.7435
  -2.500  -0.3521   0.03504   0.02871   0.0053   0.9703   0.8032
  -2.250  -0.2071   0.03683   0.02946  -0.0017   0.9693   0.9259
  -2.000  -0.1237   0.03625   0.02835  -0.0139   0.9574   0.9442
  -1.750  -0.0469   0.03565   0.02730  -0.0251   0.9451   0.9583
  -1.500   0.0325   0.03494   0.02622  -0.0368   0.9330   0.9703
  -1.250   0.1028   0.03423   0.02523  -0.0469   0.9198   0.9802
  -1.000   0.1690   0.03357   0.02435  -0.0561   0.9065   0.9889
  -0.750   0.2324   0.03296   0.02355  -0.0647   0.8936   0.9968
  -0.500   0.2799   0.03259   0.02305  -0.0701   0.8804   1.0000
  -0.250   0.3128   0.03247   0.02284  -0.0727   0.8671   1.0000
   0.000   0.3451   0.03240   0.02269  -0.0750   0.8543   1.0000
   0.250   0.3793   0.03233   0.02256  -0.0775   0.8422   1.0000
   0.500   0.4315   0.03183   0.02197  -0.0827   0.8331   1.0000
   0.750   0.4476   0.03221   0.02233  -0.0820   0.8198   1.0000
   1.250   0.4637   0.03352   0.02362  -0.0782   0.7944   1.0000
   1.500   0.4692   0.03428   0.02438  -0.0758   0.7828   1.0000
   1.750   0.4965   0.03448   0.02458  -0.0766   0.7734   1.0000
   2.000   0.5018   0.03527   0.02536  -0.0742   0.7625   1.0000
   2.250   0.4757   0.03686   0.02697  -0.0675   0.7507   1.0000
   2.500   0.4517   0.03828   0.02840  -0.0611   0.7406   1.0000
   2.750   0.4907   0.03836   0.02848  -0.0633   0.7326   1.0000
   3.000   0.4177   0.04077   0.03089  -0.0509   0.7219   1.0000
   3.250   0.4069   0.04191   0.03202  -0.0469   0.7138   1.0000
   3.500   0.4030   0.04295   0.03307  -0.0439   0.7056   1.0000
   3.750   0.3962   0.04448   0.03460  -0.0416   0.6972   1.0000
   4.000   0.4298   0.04520   0.03535  -0.0434   0.6893   1.0000
   4.250   0.4254   0.04730   0.03747  -0.0426   0.6805   1.0000
   4.500   0.4608   0.04828   0.03852  -0.0448   0.6727   1.0000
   4.750   0.4646   0.05051   0.04079  -0.0453   0.6643   1.0000
   5.000   0.4895   0.05210   0.04245  -0.0472   0.6565   1.0000
   5.250   0.5053   0.05413   0.04454  -0.0486   0.6486   1.0000
   5.500   0.5233   0.05621   0.04669  -0.0503   0.6411   1.0000
   5.750   0.5407   0.05839   0.04899  -0.0520   0.6340   1.0000
   6.000   0.5576   0.06069   0.05137  -0.0537   0.6267   1.0000
   6.250   0.5720   0.06316   0.05393  -0.0554   0.6200   1.0000
   6.500   0.5911   0.06554   0.05640  -0.0571   0.6129   1.0000
   6.750   0.5977   0.06850   0.05946  -0.0587   0.6074   1.0000
   7.000   0.6238   0.07073   0.06184  -0.0606   0.5993   1.0000
   7.250   0.6236   0.07424   0.06544  -0.0621   0.5963   1.0000
   7.500   0.6288   0.07776   0.06907  -0.0641   0.5954   1.0000
   7.750   0.6352   0.08153   0.07296  -0.0663   0.5975   1.0000
   8.000   0.6555   0.08550   0.07708  -0.0695   0.6017   1.0000
   9.750   0.7947   0.10156   0.09432  -0.0747   0.4856   1.0000
  10.000   0.7809   0.10688   0.09968  -0.0768   0.4839   1.0000
  10.250   0.7754   0.11193   0.10487  -0.0790   0.4831   1.0000
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