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FX 66-H-80 (fx66h80-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: FX 66-H-80 (fx66h80-il)
Reynolds number: 200,000
Max Cl/Cd: 62.97 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66h80-il-200000.txt
Download as CSV file: xf-fx66h80-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-H-80                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5413   0.09146   0.08851   0.0198   1.0000   0.0302
  -7.500  -0.5420   0.08779   0.08489   0.0171   1.0000   0.0307
  -7.250  -0.5418   0.08401   0.08114   0.0141   1.0000   0.0314
  -7.000  -0.5366   0.07985   0.07697   0.0101   1.0000   0.0323
  -6.750  -0.5266   0.07575   0.07282   0.0053   1.0000   0.0335
  -6.500  -0.5117   0.07219   0.06909   0.0010   1.0000   0.0343
  -6.250  -0.4971   0.06868   0.06542  -0.0007   1.0000   0.0346
  -6.000  -0.4825   0.06503   0.06161  -0.0018   1.0000   0.0347
  -5.750  -0.4791   0.05718   0.05381  -0.0029   1.0000   0.0359
  -5.500  -0.4657   0.05358   0.05024  -0.0030   1.0000   0.0370
  -5.250  -0.4493   0.05036   0.04699  -0.0034   1.0000   0.0387
  -5.000  -0.4299   0.04706   0.04360  -0.0040   1.0000   0.0410
  -4.500  -0.3814   0.04049   0.03575  -0.0041   0.7965   0.0487
  -4.250  -0.3660   0.03749   0.03262  -0.0032   0.7584   0.0508
  -4.000  -0.3463   0.03531   0.03016  -0.0023   0.7288   0.0545
  -3.750  -0.3238   0.03301   0.02720  -0.0009   0.7063   0.0627
  -3.500  -0.3030   0.03062   0.02471  -0.0004   0.6851   0.0672
  -3.250  -0.2804   0.02861   0.02229   0.0006   0.6682   0.0782
  -3.000  -0.2581   0.02710   0.02059   0.0011   0.6522   0.0949
  -2.500  -0.1956   0.02031   0.01233   0.0054   0.6289   0.0414
  -2.250  -0.1660   0.01892   0.01037   0.0067   0.6173   0.0353
  -2.000  -0.1384   0.01769   0.00891   0.0073   0.6066   0.0344
  -1.750  -0.1113   0.01636   0.00740   0.0080   0.5972   0.0344
  -1.500  -0.0857   0.01490   0.00586   0.0089   0.5880   0.0361
  -1.250  -0.0608   0.01400   0.00494   0.0097   0.5794   0.0421
  -1.000   0.0569   0.01010   0.00395  -0.0078   0.5658   1.0000
  -0.750   0.0817   0.01016   0.00374  -0.0071   0.5587   1.0000
  -0.500   0.1069   0.01019   0.00361  -0.0066   0.5508   1.0000
  -0.250   0.1318   0.01028   0.00348  -0.0060   0.5445   1.0000
   0.000   0.1572   0.01032   0.00341  -0.0056   0.5372   1.0000
   0.250   0.1822   0.01043   0.00334  -0.0050   0.5313   1.0000
   0.500   0.2075   0.01050   0.00335  -0.0046   0.5244   1.0000
   0.750   0.2326   0.01061   0.00333  -0.0040   0.5186   1.0000
   1.000   0.2580   0.01073   0.00338  -0.0036   0.5127   1.0000
   1.250   0.2833   0.01084   0.00342  -0.0031   0.5069   1.0000
   1.500   0.3086   0.01100   0.00348  -0.0026   0.5017   1.0000
   1.750   0.3341   0.01111   0.00359  -0.0022   0.4953   1.0000
   2.000   0.3593   0.01128   0.00367  -0.0017   0.4905   1.0000
   2.250   0.3849   0.01144   0.00385  -0.0013   0.4849   1.0000
   2.500   0.4102   0.01160   0.00398  -0.0009   0.4794   1.0000
   2.750   0.4355   0.01180   0.00413  -0.0004   0.4746   1.0000
   3.000   0.4611   0.01197   0.00435   0.0000   0.4686   1.0000
   3.250   0.4865   0.01217   0.00455   0.0005   0.4640   1.0000
   3.500   0.5120   0.01239   0.00482   0.0008   0.4588   1.0000
   3.750   0.5374   0.01259   0.00506   0.0012   0.4530   1.0000
   4.000   0.5628   0.01283   0.00527   0.0017   0.4488   1.0000
   4.250   0.5883   0.01308   0.00565   0.0020   0.4431   1.0000
   4.500   0.6138   0.01330   0.00596   0.0025   0.4379   1.0000
   4.750   0.6392   0.01359   0.00627   0.0029   0.4336   1.0000
   5.000   0.6647   0.01386   0.00671   0.0032   0.4276   1.0000
   5.250   0.6901   0.01410   0.00700   0.0037   0.4228   1.0000
   5.500   0.7154   0.01440   0.00745   0.0041   0.4172   1.0000
   5.750   0.7404   0.01426   0.00734   0.0049   0.4064   1.0000
   6.000   0.7648   0.01379   0.00699   0.0057   0.3877   1.0000
   6.250   0.7894   0.01331   0.00651   0.0066   0.3667   1.0000
   6.500   0.8142   0.01293   0.00629   0.0073   0.3328   1.0000
   6.750   0.8300   0.01459   0.00693   0.0079   0.1182   1.0000
   7.000   0.8446   0.01721   0.00913   0.0087   0.0437   1.0000
   7.250   0.8635   0.01864   0.01077   0.0096   0.0357   1.0000
   7.500   0.8805   0.02018   0.01244   0.0105   0.0315   1.0000
   7.750   0.8907   0.02264   0.01494   0.0123   0.0284   1.0000
   8.000   0.9099   0.02366   0.01608   0.0133   0.0253   1.0000
   8.250   0.9259   0.02534   0.01784   0.0148   0.0238   1.0000
   8.500   0.9427   0.02722   0.01980   0.0164   0.0226   1.0000
   8.750   0.9606   0.02938   0.02209   0.0179   0.0221   1.0000
   9.000   0.9797   0.03201   0.02492   0.0193   0.0223   1.0000
   9.250   0.9972   0.03533   0.02860   0.0209   0.0235   1.0000
   9.500   1.0108   0.03856   0.03218   0.0222   0.0239   1.0000
   9.750   1.0232   0.04042   0.03413   0.0228   0.0211   1.0000
  10.000   1.0274   0.04446   0.03872   0.0245   0.0225   1.0000
  10.250   1.0250   0.04978   0.04448   0.0259   0.0256   1.0000
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