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FX 66-H-60 (fx66h60-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: FX 66-H-60 (fx66h60-il)
Reynolds number: 200,000
Max Cl/Cd: 51.31 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx66h60-il-200000-n5.txt
Download as CSV file: xf-fx66h60-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5263   0.08690   0.08391   0.0194   1.0000   0.0183
  -7.500  -0.5302   0.08268   0.07972   0.0168   1.0000   0.0184
  -7.250  -0.5343   0.07864   0.07567   0.0150   1.0000   0.0184
  -7.000  -0.5335   0.07419   0.07120   0.0131   1.0000   0.0184
  -6.750  -0.5304   0.06961   0.06657   0.0116   1.0000   0.0184
  -6.500  -0.5252   0.06499   0.06189   0.0103   1.0000   0.0185
  -6.250  -0.5180   0.06030   0.05712   0.0093   1.0000   0.0185
  -6.000  -0.5089   0.05561   0.05233   0.0085   1.0000   0.0185
  -5.750  -0.4982   0.05099   0.04760   0.0081   1.0000   0.0186
  -5.500  -0.4858   0.04645   0.04291   0.0078   1.0000   0.0186
  -5.250  -0.4719   0.04196   0.03827   0.0077   1.0000   0.0186
  -5.000  -0.4569   0.03758   0.03372   0.0079   1.0000   0.0186
  -4.750  -0.4407   0.03332   0.02928   0.0082   1.0000   0.0186
  -4.500  -0.4552   0.04011   0.03561   0.0092   1.0000   0.0160
  -4.250  -0.4347   0.03623   0.03147   0.0099   1.0000   0.0141
  -4.000  -0.4119   0.03249   0.02741   0.0110   1.0000   0.0132
  -3.750  -0.3880   0.02911   0.02368   0.0123   1.0000   0.0129
  -3.500  -0.3631   0.02682   0.02111   0.0133   1.0000   0.0153
  -3.250  -0.3372   0.02456   0.01845   0.0144   1.0000   0.0167
  -3.000  -0.3121   0.02199   0.01553   0.0155   1.0000   0.0160
  -2.750  -0.2860   0.01978   0.01301   0.0165   1.0000   0.0154
  -2.500  -0.2349   0.01742   0.01000   0.0126   0.8222   0.0147
  -2.250  -0.2077   0.01619   0.00828   0.0136   0.7428   0.0145
  -2.000  -0.1827   0.01517   0.00687   0.0148   0.6873   0.0144
  -1.750  -0.1582   0.01428   0.00570   0.0160   0.6420   0.0145
  -1.500  -0.1340   0.01353   0.00472   0.0171   0.6041   0.0149
  -1.250  -0.1097   0.01294   0.00385   0.0181   0.5712   0.0158
  -1.000  -0.0849   0.01253   0.00320   0.0189   0.5425   0.0174
  -0.750  -0.0597   0.01213   0.00260   0.0197   0.5170   0.0236
  -0.500  -0.0365   0.01140   0.00225   0.0204   0.4949   0.1502
  -0.250   0.0672   0.00955   0.00263   0.0067   0.4567   0.9950
   0.000   0.1045   0.00958   0.00238   0.0046   0.4362   1.0000
   0.250   0.1283   0.00962   0.00226   0.0053   0.4209   1.0000
   0.500   0.1521   0.00967   0.00216   0.0061   0.4065   1.0000
   0.750   0.1760   0.00973   0.00209   0.0068   0.3931   1.0000
   1.000   0.2000   0.00979   0.00204   0.0075   0.3806   1.0000
   1.250   0.2240   0.00985   0.00202   0.0082   0.3696   1.0000
   1.500   0.2480   0.00993   0.00201   0.0089   0.3592   1.0000
   1.750   0.2719   0.01003   0.00203   0.0097   0.3492   1.0000
   2.000   0.2958   0.01013   0.00207   0.0104   0.3399   1.0000
   2.250   0.3197   0.01023   0.00214   0.0111   0.3311   1.0000
   2.500   0.3436   0.01035   0.00226   0.0119   0.3235   1.0000
   2.750   0.3674   0.01047   0.00236   0.0126   0.3154   1.0000
   3.000   0.3913   0.01061   0.00249   0.0134   0.3079   1.0000
   3.250   0.4151   0.01076   0.00263   0.0141   0.3009   1.0000
   3.500   0.4390   0.01091   0.00281   0.0149   0.2944   1.0000
   3.750   0.4628   0.01108   0.00300   0.0156   0.2883   1.0000
   4.000   0.4867   0.01126   0.00329   0.0164   0.2821   1.0000
   4.250   0.5106   0.01145   0.00352   0.0171   0.2758   1.0000
   4.500   0.5345   0.01167   0.00379   0.0178   0.2706   1.0000
   4.750   0.5587   0.01187   0.00411   0.0185   0.2650   1.0000
   5.000   0.5827   0.01212   0.00441   0.0193   0.2597   1.0000
   5.250   0.6070   0.01227   0.00467   0.0199   0.2478   1.0000
   5.500   0.6311   0.01235   0.00478   0.0205   0.2206   1.0000
   5.750   0.6542   0.01275   0.00509   0.0211   0.1510   1.0000
   6.000   0.6711   0.01495   0.00657   0.0220   0.0230   1.0000
   6.250   0.6936   0.01592   0.00776   0.0228   0.0157   1.0000
   6.500   0.7159   0.01690   0.00893   0.0235   0.0116   1.0000
   6.750   0.7343   0.01879   0.01114   0.0245   0.0098   1.0000
   7.000   0.7541   0.02024   0.01274   0.0255   0.0093   1.0000
   7.250   0.7737   0.02181   0.01445   0.0265   0.0089   1.0000
   7.500   0.7930   0.02358   0.01639   0.0275   0.0086   1.0000
   7.750   0.8122   0.02554   0.01853   0.0285   0.0082   1.0000
   8.000   0.8322   0.02708   0.02037   0.0292   0.0073   1.0000
   8.250   0.8523   0.02799   0.02144   0.0294   0.0056   1.0000
   8.500   0.8686   0.03010   0.02376   0.0300   0.0052   1.0000
   8.750   0.8817   0.03294   0.02689   0.0306   0.0049   1.0000
   9.000   0.8898   0.03666   0.03097   0.0311   0.0047   1.0000
   9.250   0.8930   0.04077   0.03547   0.0313   0.0046   1.0000
   9.500   0.8923   0.04472   0.03976   0.0312   0.0045   1.0000
   9.750   0.8855   0.04875   0.04409   0.0304   0.0045   1.0000
  10.000   0.8753   0.05226   0.04778   0.0288   0.0045   1.0000
  10.250   0.8686   0.05633   0.05200   0.0252   0.0046   1.0000
  10.500   0.8533   0.06270   0.05854   0.0198   0.0045   1.0000
  10.750   0.8459   0.06800   0.06395   0.0154   0.0046   1.0000
  11.000   0.8355   0.07412   0.07018   0.0107   0.0046   1.0000
  11.250   0.8290   0.07966   0.07582   0.0066   0.0046   1.0000
  11.500   0.8186   0.08628   0.08253   0.0021   0.0047   1.0000
  11.750   0.8071   0.09342   0.08975  -0.0024   0.0047   1.0000
  12.000   0.7970   0.10084   0.09725  -0.0069   0.0049   1.0000
  12.250   0.7811   0.11081   0.10729  -0.0121   0.0052   1.0000
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