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FX 66-H-60 (fx66h60-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: FX 66-H-60 (fx66h60-il)
Reynolds number: 100,000
Max Cl/Cd: 37.93 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66h60-il-100000.txt
Download as CSV file: xf-fx66h60-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6258   0.11357   0.10903   0.0307   1.0000   0.0493
  -8.500  -0.6287   0.11063   0.10616   0.0264   1.0000   0.0496
  -8.250  -0.6325   0.10745   0.10304   0.0222   1.0000   0.0497
  -8.000  -0.6335   0.10429   0.09986   0.0186   1.0000   0.0499
  -7.750  -0.6320   0.10112   0.09662   0.0152   1.0000   0.0500
  -7.500  -0.6173   0.09257   0.08826   0.0216   1.0000   0.0524
  -7.250  -0.6110   0.08859   0.08422   0.0213   1.0000   0.0545
  -7.000  -0.6058   0.08470   0.08033   0.0194   1.0000   0.0563
  -6.750  -0.5996   0.08078   0.07639   0.0169   1.0000   0.0586
  -6.500  -0.5907   0.07717   0.07267   0.0129   1.0000   0.0618
  -6.250  -0.5791   0.07617   0.07119   0.0085   1.0000   0.0634
  -6.000  -0.5704   0.06846   0.06377   0.0095   1.0000   0.0653
  -5.750  -0.5585   0.06450   0.05983   0.0099   1.0000   0.0689
  -5.500  -0.5381   0.06347   0.05822   0.0074   1.0000   0.0766
  -5.250  -0.5268   0.05704   0.05198   0.0075   1.0000   0.0791
  -5.000  -0.5108   0.05356   0.04847   0.0078   1.0000   0.0860
  -4.750  -0.4941   0.04996   0.04473   0.0076   1.0000   0.0957
  -4.500  -0.4766   0.04677   0.04140   0.0077   1.0000   0.1108
  -4.250  -0.4573   0.04433   0.03860   0.0076   1.0000   0.1323
  -4.000  -0.4399   0.04055   0.03489   0.0082   1.0000   0.1491
  -3.750  -0.4219   0.03763   0.03190   0.0089   1.0000   0.1770
  -3.250  -0.3714   0.01592   0.01081   0.0099   1.0000   0.2721
  -3.000  -0.3540   0.01346   0.00839   0.0114   1.0000   0.3148
  -2.750  -0.3291   0.01138   0.00611   0.0117   1.0000   0.3293
  -2.500  -0.2668   0.02388   0.01566   0.0137   1.0000   0.0713
  -2.250  -0.2365   0.02182   0.01315   0.0151   1.0000   0.0606
  -2.000  -0.2082   0.01990   0.01102   0.0159   1.0000   0.0593
  -1.750  -0.1799   0.01883   0.00970   0.0168   1.0000   0.0637
  -1.500  -0.1524   0.01681   0.00775   0.0174   1.0000   0.0655
  -1.250  -0.1266   0.01541   0.00644   0.0183   1.0000   0.0700
  -1.000  -0.0138   0.01063   0.00445   0.0048   1.0000   1.0000
  -0.750   0.0116   0.01054   0.00420   0.0050   1.0000   1.0000
  -0.500   0.0649   0.01086   0.00405  -0.0003   0.8197   1.0000
  -0.250   0.0886   0.01137   0.00394   0.0012   0.7286   1.0000
   0.000   0.1116   0.01170   0.00387   0.0025   0.6749   1.0000
   0.250   0.1348   0.01199   0.00383   0.0037   0.6349   1.0000
   0.500   0.1572   0.01227   0.00383   0.0049   0.6030   1.0000
   0.750   0.1804   0.01255   0.00387   0.0059   0.5757   1.0000
   1.000   0.2041   0.01283   0.00394   0.0067   0.5523   1.0000
   1.250   0.2279   0.01310   0.00406   0.0075   0.5308   1.0000
   1.500   0.2515   0.01339   0.00419   0.0084   0.5127   1.0000
   1.750   0.2751   0.01370   0.00436   0.0092   0.4966   1.0000
   2.000   0.2989   0.01401   0.00460   0.0100   0.4807   1.0000
   2.250   0.3227   0.01433   0.00486   0.0108   0.4667   1.0000
   2.500   0.3466   0.01469   0.00519   0.0115   0.4542   1.0000
   2.750   0.3704   0.01506   0.00550   0.0123   0.4425   1.0000
   3.000   0.3945   0.01542   0.00588   0.0130   0.4308   1.0000
   3.250   0.4186   0.01583   0.00634   0.0136   0.4201   1.0000
   3.500   0.4427   0.01629   0.00684   0.0143   0.4106   1.0000
   3.750   0.4667   0.01674   0.00734   0.0150   0.4022   1.0000
   4.000   0.4909   0.01722   0.00798   0.0156   0.3924   1.0000
   4.250   0.5151   0.01779   0.00867   0.0162   0.3846   1.0000
   4.500   0.5392   0.01831   0.00927   0.0169   0.3768   1.0000
   4.750   0.5633   0.01895   0.01014   0.0174   0.3685   1.0000
   5.000   0.5876   0.01958   0.01084   0.0181   0.3623   1.0000
   5.250   0.6116   0.02034   0.01202   0.0185   0.3539   1.0000
   5.500   0.6359   0.02104   0.01288   0.0192   0.3473   1.0000
   5.750   0.6592   0.02112   0.01316   0.0202   0.3314   1.0000
   6.000   0.6801   0.01888   0.01072   0.0224   0.2865   1.0000
   6.250   0.6979   0.01840   0.00940   0.0238   0.0789   1.0000
   6.500   0.7168   0.02086   0.01186   0.0249   0.0536   1.0000
   6.750   0.7349   0.02298   0.01412   0.0261   0.0473   1.0000
   7.000   0.7550   0.02456   0.01585   0.0273   0.0422   1.0000
   7.250   0.7724   0.02687   0.01816   0.0285   0.0374   1.0000
   7.500   0.7927   0.02982   0.02113   0.0299   0.0363   1.0000
   7.750   0.8145   0.03269   0.02423   0.0310   0.0362   1.0000
   8.000   0.8350   0.03621   0.02805   0.0319   0.0366   1.0000
   8.250   0.8559   0.03826   0.03054   0.0330   0.0376   1.0000
   8.500   0.8715   0.04146   0.03454   0.0340   0.0402   1.0000
   8.750   0.8783   0.04637   0.04016   0.0343   0.0433   1.0000
   9.000   0.8819   0.05151   0.04570   0.0341   0.0458   1.0000
   9.250   0.8291   0.04660   0.04145   0.0336   0.0467   1.0000
   9.500   0.8695   0.06143   0.05649   0.0310   0.0515   1.0000
   9.750   0.8485   0.06665   0.06190   0.0271   0.0517   1.0000
  10.000   0.8301   0.07266   0.06797   0.0210   0.0513   1.0000
  10.250   0.8110   0.08016   0.07549   0.0141   0.0510   1.0000
  10.500   0.7900   0.08906   0.08437   0.0068   0.0510   1.0000
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