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NACA 66-209 (naca66209-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA 66-209 (naca66209-il)
Reynolds number: 100,000
Max Cl/Cd: 40.56 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca66209-il-100000-n5.txt
Download as CSV file: xf-naca66209-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-209                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5180   0.09237   0.08741  -0.0376   1.0000   0.0425
  -9.250  -0.5113   0.08857   0.08358  -0.0368   1.0000   0.0350
  -8.750  -0.5346   0.07603   0.07107  -0.0466   1.0000   0.0263
  -8.500  -0.5462   0.07272   0.06778  -0.0460   1.0000   0.0259
  -8.250  -0.5571   0.06925   0.06433  -0.0451   1.0000   0.0257
  -8.000  -0.5664   0.06616   0.06117  -0.0434   1.0000   0.0252
  -7.750  -0.5765   0.06311   0.05805  -0.0411   1.0000   0.0249
  -7.500  -0.5858   0.06027   0.05513  -0.0382   1.0000   0.0247
  -7.250  -0.5936   0.05748   0.05220  -0.0351   1.0000   0.0250
  -7.000  -0.5991   0.05457   0.04905  -0.0319   1.0000   0.0258
  -6.750  -0.6006   0.05179   0.04608  -0.0290   1.0000   0.0258
  -6.500  -0.5846   0.04836   0.04216  -0.0289   0.9963   0.0270
  -6.250  -0.5653   0.04395   0.03735  -0.0299   0.9913   0.0268
  -6.000  -0.5430   0.03985   0.03276  -0.0307   0.9869   0.0267
  -5.750  -0.5201   0.03651   0.02890  -0.0309   0.9824   0.0269
  -5.500  -0.4942   0.03386   0.02569  -0.0311   0.9785   0.0275
  -5.250  -0.4699   0.03020   0.02167  -0.0319   0.9756   0.0297
  -5.000  -0.4452   0.02845   0.01969  -0.0321   0.9713   0.0322
  -4.750  -0.4160   0.02637   0.01717  -0.0324   0.9682   0.0329
  -4.500  -0.3850   0.02458   0.01507  -0.0329   0.9659   0.0341
  -4.250  -0.3564   0.02315   0.01341  -0.0330   0.9630   0.0355
  -4.000  -0.3298   0.02226   0.01231  -0.0329   0.9593   0.0393
  -3.750  -0.3014   0.02107   0.01103  -0.0331   0.9564   0.0414
  -3.500  -0.2730   0.01987   0.00981  -0.0336   0.9538   0.0439
  -3.250  -0.2450   0.01908   0.00896  -0.0340   0.9508   0.0475
  -3.000  -0.2220   0.01859   0.00834  -0.0333   0.9462   0.0528
  -2.750  -0.1953   0.01801   0.00774  -0.0335   0.9429   0.0631
  -2.500  -0.1658   0.01747   0.00717  -0.0341   0.9403   0.0783
  -2.250  -0.1576   0.01465   0.00711  -0.0312   0.9371   0.6341
  -2.000  -0.1502   0.01502   0.00785  -0.0248   0.9313   0.8097
  -1.750  -0.1321   0.01550   0.00828  -0.0210   0.9279   0.8643
  -1.500  -0.1044   0.01601   0.00869  -0.0189   0.9262   0.9059
  -1.250  -0.0570   0.01639   0.00890  -0.0216   0.9265   0.9348
  -1.000  -0.0204   0.01639   0.00874  -0.0238   0.9247   0.9404
  -0.750   0.0101   0.01640   0.00865  -0.0249   0.9209   0.9453
  -0.500   0.0389   0.01641   0.00858  -0.0257   0.9172   0.9507
  -0.250   0.0734   0.01642   0.00851  -0.0276   0.9146   0.9543
   0.000   0.1112   0.01643   0.00847  -0.0302   0.9125   0.9568
   0.250   0.1490   0.01644   0.00846  -0.0327   0.9108   0.9598
   0.500   0.1735   0.01652   0.00855  -0.0327   0.9052   0.9658
   0.750   0.2101   0.01655   0.00859  -0.0351   0.9019   0.9683
   1.000   0.2484   0.01656   0.00863  -0.0377   0.8990   0.9710
   1.250   0.2849   0.01657   0.00873  -0.0399   0.8959   0.9743
   1.500   0.3152   0.01666   0.00890  -0.0410   0.8900   0.9789
   1.750   0.3518   0.01666   0.00901  -0.0432   0.8861   0.9818
   2.000   0.3907   0.01662   0.00910  -0.0457   0.8829   0.9842
   2.250   0.4205   0.01670   0.00937  -0.0467   0.8751   0.9896
   2.500   0.4598   0.01642   0.00927  -0.0486   0.8664   0.9916
   2.750   0.5007   0.01479   0.00774  -0.0476   0.8266   0.9902
   3.000   0.5263   0.01380   0.00681  -0.0449   0.7650   0.9945
   3.250   0.5475   0.01350   0.00556  -0.0410   0.5175   0.9983
   3.500   0.5384   0.01589   0.00612  -0.0350   0.1768   1.0000
   3.750   0.5436   0.01719   0.00673  -0.0315   0.0748   1.0000
   4.000   0.5553   0.01785   0.00734  -0.0287   0.0608   1.0000
   4.250   0.5683   0.01840   0.00799  -0.0260   0.0545   1.0000
   4.500   0.5799   0.01916   0.00874  -0.0233   0.0479   1.0000
   4.750   0.5955   0.01979   0.00953  -0.0212   0.0434   1.0000
   5.000   0.6118   0.02066   0.01047  -0.0193   0.0407   1.0000
   5.250   0.6301   0.02171   0.01156  -0.0178   0.0385   1.0000
   5.500   0.6512   0.02302   0.01287  -0.0168   0.0367   1.0000
   5.750   0.6768   0.02510   0.01496  -0.0166   0.0351   1.0000
   6.000   0.7024   0.02641   0.01648  -0.0161   0.0331   1.0000
   6.250   0.7278   0.02789   0.01827  -0.0156   0.0309   1.0000
   6.500   0.7532   0.02998   0.02066  -0.0151   0.0298   1.0000
   6.750   0.7765   0.03231   0.02337  -0.0142   0.0291   1.0000
   7.000   0.7971   0.03494   0.02641  -0.0129   0.0286   1.0000
   7.250   0.8149   0.03775   0.02967  -0.0114   0.0281   1.0000
   7.500   0.8305   0.04003   0.03224  -0.0100   0.0262   1.0000
   7.750   0.8444   0.04223   0.03462  -0.0090   0.0245   1.0000
   8.000   0.8535   0.04568   0.03837  -0.0073   0.0237   1.0000
   8.250   0.8596   0.04949   0.04261  -0.0053   0.0235   1.0000
   8.500   0.8629   0.05331   0.04680  -0.0032   0.0234   1.0000
   8.750   0.8641   0.05700   0.05082  -0.0010   0.0234   1.0000
   9.000   0.8623   0.06065   0.05478   0.0012   0.0235   1.0000
   9.250   0.8572   0.06425   0.05867   0.0034   0.0237   1.0000
   9.500   0.8472   0.06781   0.06248   0.0058   0.0241   1.0000
   9.750   0.8255   0.07199   0.06693   0.0082   0.0247   1.0000
  10.000   0.8006   0.07730   0.07244   0.0083   0.0255   1.0000
  10.250   0.7785   0.08335   0.07861   0.0060   0.0262   1.0000
  10.500   0.7590   0.09069   0.08597   0.0013   0.0269   1.0000
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