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FX 66-H-60 (fx66h60-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 66-H-60 (fx66h60-il)
Reynolds number: 50,000
Max Cl/Cd: 28.56 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66h60-il-50000.txt
Download as CSV file: xf-fx66h60-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-H-60                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6390   0.11083   0.10476   0.0335   1.0000   0.1570
  -8.000  -0.6262   0.10579   0.09965   0.0346   1.0000   0.1661
  -7.750  -0.6406   0.10361   0.09760   0.0304   1.0000   0.1708
  -7.500  -0.6281   0.09880   0.09282   0.0317   1.0000   0.1818
  -7.000  -0.6405   0.09286   0.08698   0.0244   1.0000   0.1988
  -6.750  -0.6210   0.08726   0.08147   0.0289   1.0000   0.2174
  -6.500  -0.6188   0.08372   0.07797   0.0281   1.0000   0.2366
  -6.250  -0.6116   0.07976   0.07408   0.0293   1.0000   0.2590
  -6.000  -0.6068   0.07627   0.07065   0.0306   1.0000   0.2899
  -4.750  -0.3191   0.05237   0.04637   0.0545   1.0000   0.9037
  -4.500  -0.3772   0.05294   0.04719   0.0619   1.0000   0.8131
  -4.250  -0.4255   0.05219   0.04668   0.0658   1.0000   0.7528
  -4.000  -0.4683   0.04990   0.04465   0.0662   1.0000   0.6896
  -3.750  -0.5070   0.04632   0.04131   0.0610   1.0000   0.6248
  -3.500  -0.4012   0.03726   0.02927   0.0091   1.0000   0.2055
  -3.250  -0.3620   0.03468   0.02575   0.0090   1.0000   0.1560
  -3.000  -0.3314   0.03189   0.02252   0.0097   1.0000   0.1391
  -2.750  -0.3000   0.02947   0.01959   0.0106   1.0000   0.1244
  -2.500  -0.2685   0.02742   0.01702   0.0116   1.0000   0.1149
  -2.250  -0.2379   0.02522   0.01456   0.0123   1.0000   0.1114
  -2.000  -0.2062   0.02330   0.01241   0.0130   1.0000   0.1115
  -1.750  -0.1719   0.02159   0.01055   0.0133   1.0000   0.1199
  -1.500  -0.1411   0.02000   0.00900   0.0136   1.0000   0.1405
  -1.250  -0.0391   0.01461   0.00596   0.0040   1.0000   1.0000
  -1.000  -0.0146   0.01451   0.00548   0.0046   1.0000   1.0000
  -0.750   0.0100   0.01443   0.00518   0.0050   1.0000   1.0000
  -0.500   0.0351   0.01437   0.00502   0.0051   1.0000   1.0000
  -0.250   0.0613   0.01436   0.00497   0.0045   1.0000   1.0000
   0.000   0.0860   0.01460   0.00539   0.0022   1.0000   1.0000
   0.250   0.1702   0.01543   0.00561  -0.0088   0.8284   1.0000
   0.500   0.1903   0.01621   0.00600  -0.0065   0.7715   1.0000
   0.750   0.2106   0.01685   0.00637  -0.0046   0.7306   1.0000
   1.000   0.2320   0.01744   0.00674  -0.0029   0.6976   1.0000
   1.250   0.2546   0.01801   0.00715  -0.0017   0.6690   1.0000
   1.500   0.2774   0.01858   0.00757  -0.0005   0.6449   1.0000
   1.750   0.3006   0.01916   0.00809   0.0004   0.6222   1.0000
   2.000   0.3233   0.01979   0.00864   0.0014   0.6032   1.0000
   2.250   0.3466   0.02045   0.00930   0.0022   0.5847   1.0000
   2.500   0.3702   0.02115   0.01004   0.0029   0.5686   1.0000
   2.750   0.3940   0.02189   0.01079   0.0035   0.5537   1.0000
   3.000   0.4174   0.02269   0.01163   0.0041   0.5399   1.0000
   3.250   0.4410   0.02358   0.01260   0.0046   0.5271   1.0000
   3.500   0.4646   0.02453   0.01368   0.0049   0.5149   1.0000
   3.750   0.4881   0.02557   0.01486   0.0052   0.5036   1.0000
   4.000   0.5115   0.02668   0.01622   0.0055   0.4939   1.0000
   4.250   0.5346   0.02776   0.01744   0.0061   0.4842   1.0000
   4.500   0.5581   0.02925   0.01923   0.0056   0.4740   1.0000
   4.750   0.5810   0.03082   0.02107   0.0053   0.4655   1.0000
   5.000   0.6035   0.03211   0.02254   0.0060   0.4570   1.0000
   5.250   0.6257   0.03436   0.02520   0.0046   0.4481   1.0000
   5.500   0.6471   0.03627   0.02753   0.0044   0.4399   1.0000
   5.750   0.6681   0.03824   0.02982   0.0042   0.4311   1.0000
   6.000   0.6866   0.04118   0.03318   0.0025   0.4217   1.0000
   6.250   0.7048   0.04360   0.03597   0.0022   0.4110   1.0000
   6.500   0.7297   0.03062   0.02223   0.0205   0.3182   1.0000
   6.750   0.7438   0.02604   0.01654   0.0262   0.1350   1.0000
   7.000   0.7611   0.02875   0.01894   0.0274   0.1030   1.0000
   7.250   0.7776   0.03130   0.02148   0.0284   0.0865   1.0000
   7.500   0.7973   0.03393   0.02434   0.0300   0.0795   1.0000
   7.750   0.8177   0.03686   0.02746   0.0314   0.0754   1.0000
   8.000   0.8377   0.04038   0.03121   0.0326   0.0736   1.0000
   8.250   0.8549   0.04453   0.03577   0.0333   0.0731   1.0000
   8.500   0.8676   0.04902   0.04076   0.0336   0.0730   1.0000
   8.750   0.8741   0.05370   0.04604   0.0332   0.0736   1.0000
   9.000   0.8608   0.06009   0.05322   0.0303   0.0761   1.0000
   9.250   0.8393   0.06726   0.06075   0.0255   0.0791   1.0000
   9.500   0.8175   0.07406   0.06761   0.0199   0.0807   1.0000
   9.750   0.7992   0.08155   0.07508   0.0138   0.0823   1.0000
  10.000   0.7917   0.08793   0.08140   0.0104   0.0840   1.0000
  10.250   0.6579   0.09327   0.08682   0.0043   0.0912   1.0000
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