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NACA 66 (p51htip-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 66 (p51htip-il)
Reynolds number: 200,000
Max Cl/Cd: 52.55 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-p51htip-il-200000.txt
Download as CSV file: xf-p51htip-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66                                         
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4689   0.09313   0.08972  -0.0499   1.0000   0.0522
 -10.250  -0.4747   0.08826   0.08490  -0.0533   1.0000   0.0534
 -10.000  -0.4919   0.08353   0.08026  -0.0561   1.0000   0.0542
  -9.750  -0.5232   0.08143   0.07824  -0.0536   0.9991   0.0534
  -9.500  -0.5306   0.07416   0.07086  -0.0625   0.9919   0.0540
  -9.250  -0.5465   0.06836   0.06478  -0.0691   0.9830   0.0559
  -9.000  -0.4645   0.04832   0.04482  -0.0792   0.9731   0.0610
  -8.750  -0.4665   0.04503   0.04145  -0.0788   0.9660   0.0625
  -8.500  -0.4657   0.04125   0.03749  -0.0791   0.9610   0.0651
  -8.250  -0.4858   0.04073   0.03637  -0.0747   0.9502   0.0704
  -8.000  -0.4856   0.03401   0.02954  -0.0746   0.9469   0.0724
  -7.750  -0.4717   0.03103   0.02662  -0.0742   0.9431   0.0745
  -7.500  -0.4634   0.02883   0.02431  -0.0726   0.9376   0.0782
  -7.250  -0.4640   0.02593   0.02097  -0.0698   0.9327   0.0868
  -7.000  -0.4496   0.02372   0.01880  -0.0690   0.9293   0.0903
  -6.750  -0.4490   0.02212   0.01685  -0.0656   0.9234   0.1009
  -6.500  -0.4303   0.02023   0.01501  -0.0652   0.9209   0.1082
  -6.250  -0.4335   0.02823   0.02072  -0.0594   0.9226   0.0465
  -6.000  -0.4076   0.02733   0.01957  -0.0588   0.9202   0.0456
  -5.750  -0.3801   0.02569   0.01771  -0.0587   0.9184   0.0454
  -5.500  -0.3497   0.02282   0.01473  -0.0596   0.9173   0.0481
  -5.250  -0.3323   0.02202   0.01388  -0.0580   0.9130   0.0495
  -5.000  -0.3084   0.02113   0.01292  -0.0574   0.9100   0.0508
  -4.750  -0.2834   0.02033   0.01208  -0.0569   0.9078   0.0530
  -4.500  -0.2583   0.01984   0.01152  -0.0566   0.9058   0.0565
  -4.250  -0.2359   0.01895   0.01064  -0.0558   0.9040   0.0596
  -4.000  -0.2203   0.01852   0.01024  -0.0539   0.9013   0.0630
  -3.750  -0.2122   0.01845   0.01015  -0.0506   0.8965   0.0670
  -3.500  -0.1966   0.01815   0.00983  -0.0486   0.8936   0.0738
  -3.250  -0.1772   0.01783   0.00950  -0.0473   0.8914   0.0862
  -3.000  -0.1709   0.01625   0.00901  -0.0442   0.8891   0.3072
  -2.750  -0.1949   0.01525   0.00941  -0.0350   0.8833   0.6008
  -2.500  -0.1905   0.01566   0.01033  -0.0292   0.8798   0.7698
  -2.250  -0.1720   0.01618   0.01083  -0.0264   0.8774   0.8111
  -2.000  -0.1516   0.01683   0.01147  -0.0236   0.8757   0.8407
  -1.750  -0.1287   0.01785   0.01252  -0.0201   0.8744   0.8724
  -1.500  -0.0777   0.01918   0.01381  -0.0216   0.8745   0.8978
  -1.250  -0.0062   0.01980   0.01431  -0.0293   0.8752   0.9069
  -1.000  -0.0060   0.01999   0.01447  -0.0245   0.8712   0.9167
  -0.750   0.0302   0.02024   0.01467  -0.0268   0.8686   0.9197
  -0.500   0.0633   0.02040   0.01478  -0.0285   0.8661   0.9234
  -0.250   0.0708   0.02043   0.01478  -0.0249   0.8623   0.9312
   0.000   0.1173   0.02045   0.01477  -0.0291   0.8612   0.9323
   0.250   0.1619   0.02044   0.01473  -0.0329   0.8602   0.9335
   0.500   0.2050   0.02040   0.01468  -0.0362   0.8591   0.9349
   0.750   0.1805   0.02092   0.01522  -0.0273   0.8495   0.9457
   1.000   0.2226   0.02089   0.01520  -0.0306   0.8476   0.9467
   1.250   0.2653   0.02082   0.01515  -0.0338   0.8459   0.9476
   1.500   0.3086   0.02071   0.01506  -0.0370   0.8445   0.9486
   1.750   0.3107   0.02117   0.01556  -0.0333   0.8358   0.9553
   2.000   0.3380   0.02105   0.01549  -0.0333   0.8322   0.9589
   2.250   0.3837   0.02079   0.01528  -0.0367   0.8303   0.9594
   2.500   0.4121   0.02097   0.01554  -0.0377   0.8235   0.9621
   2.750   0.4503   0.02071   0.01538  -0.0397   0.8187   0.9635
   3.000   0.5121   0.01886   0.01359  -0.0434   0.8141   0.9612
   3.250   0.5522   0.01692   0.01171  -0.0431   0.7950   0.9618
   3.500   0.5936   0.01451   0.00923  -0.0422   0.7707   0.9624
   3.750   0.6234   0.01353   0.00831  -0.0414   0.7458   0.9655
   4.000   0.6485   0.01280   0.00762  -0.0400   0.7139   0.9696
   4.250   0.6605   0.01257   0.00650  -0.0356   0.5047   0.9742
   4.500   0.6427   0.01570   0.00766  -0.0293   0.1507   0.9829
   4.750   0.6595   0.01723   0.00865  -0.0285   0.0723   0.9876
   5.000   0.6845   0.01807   0.00948  -0.0289   0.0612   0.9921
   5.250   0.7110   0.01881   0.01030  -0.0295   0.0557   0.9960
   5.500   0.7379   0.01971   0.01120  -0.0302   0.0517   0.9997
   5.750   0.7422   0.02047   0.01195  -0.0264   0.0499   1.0000
   6.000   0.7483   0.02149   0.01296  -0.0229   0.0480   1.0000
   6.250   0.7547   0.02188   0.01341  -0.0193   0.0464   1.0000
   6.500   0.7640   0.02248   0.01406  -0.0161   0.0447   1.0000
   6.750   0.7805   0.02337   0.01502  -0.0142   0.0435   1.0000
   7.000   0.8036   0.02455   0.01627  -0.0133   0.0425   1.0000
   7.250   0.8292   0.02600   0.01781  -0.0129   0.0417   1.0000
   7.500   0.8497   0.02737   0.01929  -0.0116   0.0408   1.0000
   7.750   0.8684   0.02892   0.02088  -0.0104   0.0392   1.0000
   8.000   0.8899   0.03107   0.02323  -0.0094   0.0389   1.0000
   8.250   0.9074   0.03306   0.02554  -0.0072   0.0398   1.0000
   8.500   0.9205   0.03807   0.03128  -0.0040   0.0457   1.0000
  13.500   0.5963   0.12709   0.12384   0.0079   0.0603   1.0000
  13.750   0.5776   0.13218   0.12892   0.0033   0.0600   1.0000
  14.000   0.5527   0.13736   0.13407  -0.0030   0.0587   1.0000
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