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FX 61-147 AIRFOIL (fx61147-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 61-147 AIRFOIL (fx61147-il)
Reynolds number: 50,000
Max Cl/Cd: 8.36 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx61147-il-50000.txt
Download as CSV file: xf-fx61147-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 61-147 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3024   0.11030   0.10386  -0.0255   1.0000   0.3324
  -8.750  -0.3063   0.10820   0.10186  -0.0243   1.0000   0.3454
  -8.500  -0.3120   0.10635   0.10013  -0.0228   1.0000   0.3592
  -8.250  -0.2963   0.10259   0.09638  -0.0217   1.0000   0.3697
  -8.000  -0.2947   0.09975   0.09363  -0.0204   1.0000   0.3785
  -7.750  -0.3024   0.09785   0.09186  -0.0185   1.0000   0.3906
  -7.500  -0.2969   0.09478   0.08885  -0.0173   1.0000   0.3961
  -7.250  -0.3107   0.09311   0.08734  -0.0149   1.0000   0.4055
  -7.000  -0.3980   0.08675   0.08134  -0.0211   1.0000   0.2875
  -6.750  -0.4481   0.08544   0.08033  -0.0181   1.0000   0.2846
  -6.500  -0.5458   0.06940   0.06411  -0.0376   1.0000   0.1810
  -6.250  -0.5516   0.06196   0.05613  -0.0420   1.0000   0.1462
  -6.000  -0.5457   0.05676   0.05041  -0.0440   1.0000   0.1313
  -5.750  -0.5309   0.05221   0.04487  -0.0467   1.0000   0.1197
  -5.500  -0.5149   0.04885   0.04135  -0.0468   1.0000   0.1180
  -5.250  -0.4961   0.04569   0.03787  -0.0473   1.0000   0.1154
  -5.000  -0.4742   0.04277   0.03448  -0.0479   1.0000   0.1121
  -4.750  -0.4516   0.04038   0.03161  -0.0481   1.0000   0.1103
  -4.500  -0.4313   0.03868   0.02963  -0.0479   1.0000   0.1129
  -4.250  -0.4106   0.03730   0.02789  -0.0475   1.0000   0.1169
  -4.000  -0.3902   0.03618   0.02640  -0.0468   1.0000   0.1197
  -3.750  -0.3712   0.03491   0.02523  -0.0455   0.9991   0.1240
  -3.500  -0.3280   0.03417   0.02452  -0.0475   0.9886   0.1387
  -3.250  -0.2892   0.03344   0.02390  -0.0489   0.9777   0.1582
  -3.000  -0.2467   0.03169   0.02271  -0.0526   0.9669   0.2211
  -2.750  -0.2516   0.03330   0.02700  -0.0400   0.9546   0.6815
  -2.500  -0.2571   0.03585   0.02940  -0.0282   0.9414   0.7465
  -2.250  -0.2614   0.03707   0.03050  -0.0179   0.9290   0.7932
  -2.000  -0.1044   0.03932   0.03181  -0.0246   0.9255   0.9415
  -1.750  -0.0313   0.03894   0.03097  -0.0348   0.9141   0.9578
  -1.500   0.0436   0.03851   0.03013  -0.0455   0.9030   0.9710
  -1.250   0.1207   0.03799   0.02926  -0.0565   0.8925   0.9825
  -1.000   0.1778   0.03763   0.02865  -0.0641   0.8798   0.9919
  -0.750   0.2308   0.03741   0.02824  -0.0709   0.8679   1.0000
  -0.500   0.2508   0.03756   0.02827  -0.0715   0.8555   1.0000
  -0.250   0.2765   0.03764   0.02824  -0.0728   0.8446   1.0000
   0.000   0.3008   0.03775   0.02825  -0.0738   0.8337   1.0000
   0.250   0.3019   0.03836   0.02882  -0.0712   0.8212   1.0000
   0.500   0.3038   0.03900   0.02940  -0.0688   0.8098   1.0000
   0.750   0.3224   0.03933   0.02966  -0.0687   0.8002   1.0000
   1.000   0.3328   0.03981   0.03008  -0.0674   0.7899   1.0000
   1.250   0.3160   0.04085   0.03112  -0.0624   0.7791   1.0000
   1.500   0.3269   0.04139   0.03160  -0.0611   0.7703   1.0000
   1.750   0.3240   0.04213   0.03232  -0.0579   0.7608   1.0000
   2.000   0.3055   0.04316   0.03334  -0.0529   0.7518   1.0000
   2.250   0.3256   0.04357   0.03370  -0.0527   0.7442   1.0000
   2.500   0.2941   0.04478   0.03491  -0.0464   0.7355   1.0000
   2.750   0.3275   0.04522   0.03529  -0.0479   0.7284   1.0000
   3.000   0.3101   0.04680   0.03687  -0.0447   0.7198   1.0000
   3.250   0.3437   0.04764   0.03766  -0.0468   0.7121   1.0000
   3.500   0.3464   0.04935   0.03937  -0.0465   0.7042   1.0000
   3.750   0.3775   0.05050   0.04050  -0.0486   0.6963   1.0000
   4.000   0.3881   0.05229   0.04230  -0.0493   0.6885   1.0000
   4.250   0.4161   0.05373   0.04374  -0.0513   0.6809   1.0000
   4.500   0.4293   0.05565   0.04568  -0.0524   0.6732   1.0000
   4.750   0.4553   0.05731   0.04736  -0.0543   0.6657   1.0000
   5.000   0.4657   0.05952   0.04961  -0.0554   0.6590   1.0000
   5.250   0.4951   0.06120   0.05131  -0.0575   0.6514   1.0000
   5.500   0.4981   0.06389   0.05405  -0.0583   0.6462   1.0000
   5.750   0.5364   0.06536   0.05558  -0.0607   0.6376   1.0000
   6.000   0.5335   0.06853   0.05880  -0.0615   0.6348   1.0000
   6.250   0.5388   0.07149   0.06182  -0.0628   0.6325   1.0000
   6.500   0.5458   0.07459   0.06498  -0.0643   0.6325   1.0000
   6.750   0.5580   0.07797   0.06844  -0.0665   0.6362   1.0000
   7.000   0.4755   0.08583   0.07642  -0.0684   0.7404   1.0000
   7.250   0.5078   0.08903   0.07969  -0.0715   0.7287   1.0000
   7.500   0.5160   0.09078   0.08149  -0.0717   0.7162   1.0000
   7.750   0.5292   0.09311   0.08388  -0.0727   0.7043   1.0000
   8.000   0.5482   0.09597   0.08684  -0.0743   0.6943   1.0000
   8.250   0.5818   0.09966   0.09062  -0.0773   0.6831   1.0000
   8.500   0.5965   0.10198   0.09303  -0.0782   0.6694   1.0000
   8.750   0.6048   0.10415   0.09528  -0.0786   0.6564   1.0000
   9.000   0.6161   0.10672   0.09796  -0.0794   0.6437   1.0000
   9.250   0.6299   0.10960   0.10092  -0.0805   0.6319   1.0000
   9.500   0.6489   0.11282   0.10425  -0.0821   0.6202   1.0000
   9.750   0.6767   0.11661   0.10816  -0.0842   0.6078   1.0000
  10.000   0.6969   0.11984   0.11155  -0.0855   0.5937   1.0000
  10.250   0.7030   0.12222   0.11403  -0.0859   0.5797   1.0000
  10.500   0.7076   0.12482   0.11671  -0.0864   0.5666   1.0000
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