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NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-210 (naca66210-il)
Reynolds number: 200,000
Max Cl/Cd: 51.39 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66210-il-200000.txt
Download as CSV file: xf-naca66210-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-210                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4029   0.09135   0.08798  -0.0394   1.0000   0.0439
 -10.000  -0.4064   0.08692   0.08357  -0.0409   1.0000   0.0450
  -9.750  -0.4125   0.08212   0.07880  -0.0429   1.0000   0.0464
  -9.500  -0.4205   0.07710   0.07382  -0.0453   1.0000   0.0472
  -9.250  -0.4325   0.07160   0.06837  -0.0483   1.0000   0.0477
  -9.000  -0.4493   0.06602   0.06282  -0.0515   1.0000   0.0475
  -8.750  -0.4705   0.06212   0.05895  -0.0516   1.0000   0.0471
  -8.500  -0.4992   0.06023   0.05711  -0.0477   1.0000   0.0465
  -8.250  -0.5306   0.05947   0.05640  -0.0420   1.0000   0.0455
  -8.000  -0.5613   0.05849   0.05544  -0.0364   1.0000   0.0447
  -7.750  -0.5750   0.05483   0.05168  -0.0357   0.9977   0.0455
  -7.500  -0.5738   0.04938   0.04592  -0.0384   0.9924   0.0484
  -7.250  -0.5677   0.04729   0.04312  -0.0394   0.9853   0.0503
  -7.000  -0.6123   0.05711   0.05242  -0.0341   0.9878   0.0503
  -6.250  -0.5155   0.02748   0.02288  -0.0427   0.9716   0.0680
  -6.000  -0.5413   0.03971   0.03456  -0.0387   0.9734   0.0700
  -5.750  -0.5161   0.03682   0.03128  -0.0400   0.9706   0.0806
  -5.500  -0.4996   0.03525   0.02935  -0.0390   0.9656   0.0929
  -5.250  -0.4764   0.03291   0.02703  -0.0392   0.9623   0.1003
  -5.000  -0.4426   0.02789   0.02091  -0.0374   0.9598   0.0661
  -4.750  -0.4066   0.02453   0.01703  -0.0370   0.9584   0.0476
  -4.500  -0.3729   0.02230   0.01447  -0.0378   0.9570   0.0459
  -4.250  -0.3455   0.02091   0.01290  -0.0375   0.9542   0.0461
  -4.000  -0.3225   0.02018   0.01202  -0.0366   0.9501   0.0486
  -3.750  -0.2938   0.01944   0.01117  -0.0368   0.9475   0.0500
  -3.500  -0.2659   0.01796   0.00973  -0.0370   0.9456   0.0524
  -3.250  -0.2358   0.01714   0.00892  -0.0378   0.9437   0.0561
  -3.000  -0.2022   0.01669   0.00843  -0.0393   0.9422   0.0633
  -2.750  -0.1927   0.01622   0.00799  -0.0361   0.9360   0.0692
  -2.500  -0.1669   0.01578   0.00755  -0.0361   0.9331   0.0854
  -2.250  -0.1624   0.01296   0.00748  -0.0326   0.9302   0.7063
  -2.000  -0.1398   0.01352   0.00822  -0.0297   0.9282   0.8161
  -1.750  -0.1390   0.01399   0.00872  -0.0234   0.9218   0.8438
  -1.500  -0.1300   0.01468   0.00946  -0.0175   0.9180   0.8788
  -1.250  -0.1117   0.01526   0.01002  -0.0136   0.9156   0.9065
  -1.000  -0.0775   0.01567   0.01035  -0.0137   0.9145   0.9236
  -0.750  -0.0247   0.01613   0.01072  -0.0179   0.9147   0.9385
  -0.500  -0.0082   0.01622   0.01077  -0.0162   0.9093   0.9464
  -0.250   0.0317   0.01628   0.01077  -0.0192   0.9068   0.9493
   0.000   0.0715   0.01630   0.01075  -0.0221   0.9046   0.9521
   0.250   0.1107   0.01629   0.01072  -0.0249   0.9026   0.9546
   0.500   0.1477   0.01627   0.01068  -0.0272   0.9008   0.9572
   0.750   0.1832   0.01633   0.01075  -0.0294   0.8975   0.9591
   1.000   0.2156   0.01641   0.01085  -0.0310   0.8929   0.9619
   1.250   0.2565   0.01639   0.01087  -0.0342   0.8903   0.9635
   1.500   0.2992   0.01629   0.01083  -0.0375   0.8881   0.9646
   1.750   0.3438   0.01613   0.01074  -0.0410   0.8862   0.9653
   2.000   0.3604   0.01628   0.01095  -0.0394   0.8777   0.9713
   2.250   0.4199   0.01512   0.00989  -0.0439   0.8675   0.9688
   2.500   0.4638   0.01382   0.00866  -0.0449   0.8504   0.9689
   2.750   0.4978   0.01247   0.00730  -0.0437   0.8276   0.9704
   3.000   0.5280   0.01146   0.00632  -0.0425   0.7911   0.9725
   3.250   0.5565   0.01083   0.00565  -0.0414   0.7343   0.9755
   3.500   0.5529   0.01223   0.00523  -0.0346   0.3590   0.9828
   3.750   0.5627   0.01478   0.00626  -0.0330   0.0751   0.9882
   4.000   0.5900   0.01551   0.00700  -0.0334   0.0619   0.9923
   4.250   0.6164   0.01645   0.00796  -0.0338   0.0545   0.9963
   4.500   0.6462   0.01723   0.00878  -0.0348   0.0493   1.0000
   4.750   0.6562   0.01790   0.00945  -0.0318   0.0471   1.0000
   5.000   0.6667   0.01891   0.01044  -0.0289   0.0452   1.0000
   5.250   0.6869   0.02091   0.01242  -0.0276   0.0437   1.0000
   5.500   0.7024   0.02163   0.01328  -0.0253   0.0428   1.0000
   5.750   0.7159   0.02224   0.01404  -0.0227   0.0409   1.0000
   6.000   0.7346   0.02362   0.01557  -0.0209   0.0404   1.0000
   6.250   0.7541   0.02541   0.01756  -0.0191   0.0405   1.0000
   6.500   0.7737   0.02777   0.02018  -0.0174   0.0414   1.0000
   6.750   0.7982   0.03102   0.02356  -0.0167   0.0446   1.0000
   7.000   0.8238   0.03878   0.03296  -0.0104   0.0919   1.0000
   7.250   0.8406   0.04090   0.03515  -0.0093   0.0855   1.0000
   7.500   0.8472   0.04331   0.03794  -0.0069   0.0759   1.0000
   7.750   0.8634   0.04606   0.04070  -0.0062   0.0717   1.0000
   8.000   0.8641   0.04951   0.04453  -0.0036   0.0660   1.0000
   8.250   0.8715   0.05179   0.04708  -0.0015   0.0611   1.0000
   8.500   0.8834   0.05479   0.05010  -0.0008   0.0586   1.0000
   8.750   0.8890   0.06485   0.05996  -0.0020   0.0561   1.0000
   9.000   0.8751   0.06358   0.05929   0.0032   0.0548   1.0000
   9.250   0.8607   0.06551   0.06162   0.0069   0.0512   1.0000
   9.500   0.8491   0.06861   0.06487   0.0092   0.0497   1.0000
   9.750   0.8344   0.07168   0.06805   0.0114   0.0488   1.0000
  10.000   0.8172   0.07528   0.07174   0.0125   0.0483   1.0000
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