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NACA 66-209 (naca66209-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-209 (naca66209-il)
Reynolds number: 200,000
Max Cl/Cd: 53.3 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66209-il-200000.txt
Download as CSV file: xf-naca66209-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.4151   0.11308   0.10960  -0.0288   1.0000   0.0358
 -11.000  -0.4186   0.10928   0.10583  -0.0315   1.0000   0.0363
 -10.750  -0.4225   0.10532   0.10189  -0.0343   1.0000   0.0366
 -10.500  -0.4241   0.10111   0.09770  -0.0364   1.0000   0.0366
 -10.250  -0.4275   0.09665   0.09326  -0.0390   1.0000   0.0367
 -10.000  -0.4304   0.08961   0.08626  -0.0383   1.0000   0.0380
  -9.750  -0.4239   0.08621   0.08287  -0.0373   1.0000   0.0391
  -9.500  -0.4231   0.08234   0.07901  -0.0378   1.0000   0.0399
  -9.250  -0.4252   0.07816   0.07486  -0.0389   1.0000   0.0408
  -9.000  -0.4299   0.07373   0.07046  -0.0405   1.0000   0.0415
  -8.750  -0.4381   0.06889   0.06566  -0.0427   1.0000   0.0421
  -8.500  -0.4531   0.06367   0.06048  -0.0461   1.0000   0.0418
  -8.250  -0.4755   0.05964   0.05648  -0.0468   1.0000   0.0416
  -8.000  -0.5059   0.05785   0.05473  -0.0426   1.0000   0.0409
  -7.750  -0.5969   0.06517   0.06160  -0.0385   1.0000   0.0384
  -7.500  -0.6043   0.06306   0.05949  -0.0351   1.0000   0.0390
  -7.250  -0.6119   0.06092   0.05731  -0.0318   1.0000   0.0396
  -7.000  -0.6172   0.05864   0.05499  -0.0288   1.0000   0.0405
  -6.750  -0.6202   0.05622   0.05247  -0.0262   1.0000   0.0416
  -6.500  -0.6183   0.05346   0.04957  -0.0243   0.9996   0.0432
  -6.250  -0.5873   0.05177   0.04711  -0.0266   0.9949   0.0493
  -6.000  -0.5741   0.04457   0.03972  -0.0282   0.9912   0.0514
  -5.750  -0.5518   0.04126   0.03641  -0.0296   0.9879   0.0543
  -5.500  -0.5242   0.03903   0.03344  -0.0307   0.9841   0.0641
  -5.250  -0.5003   0.03534   0.02990  -0.0321   0.9820   0.0688
  -5.000  -0.4797   0.03315   0.02737  -0.0321   0.9777   0.0803
  -4.750  -0.4549   0.03115   0.02527  -0.0330   0.9744   0.0967
  -4.250  -0.3810   0.02335   0.01594  -0.0316   0.9708   0.0452
  -4.000  -0.3468   0.02129   0.01355  -0.0324   0.9694   0.0441
  -3.750  -0.3243   0.02030   0.01238  -0.0312   0.9654   0.0463
  -3.500  -0.2960   0.01901   0.01092  -0.0311   0.9627   0.0462
  -3.250  -0.2660   0.01795   0.00977  -0.0315   0.9604   0.0476
  -3.000  -0.2338   0.01725   0.00900  -0.0324   0.9582   0.0497
  -2.750  -0.2027   0.01611   0.00796  -0.0336   0.9564   0.0573
  -2.500  -0.1811   0.01568   0.00748  -0.0326   0.9523   0.0621
  -2.250  -0.1568   0.01515   0.00691  -0.0323   0.9487   0.0751
  -2.000  -0.1417   0.01244   0.00642  -0.0315   0.9458   0.5724
  -1.750  -0.1283   0.01259   0.00738  -0.0262   0.9428   0.8332
  -1.500  -0.1260   0.01312   0.00799  -0.0194   0.9373   0.8779
  -1.250  -0.1175   0.01358   0.00845  -0.0136   0.9326   0.9132
  -1.000  -0.0893   0.01395   0.00874  -0.0123   0.9305   0.9382
  -0.750  -0.0071   0.01460   0.00927  -0.0221   0.9346   0.9605
  -0.500   0.0399   0.01466   0.00925  -0.0265   0.9336   0.9642
  -0.250   0.0805   0.01468   0.00920  -0.0297   0.9316   0.9679
   0.000   0.1215   0.01468   0.00917  -0.0329   0.9298   0.9705
   0.250   0.1679   0.01465   0.00912  -0.0372   0.9286   0.9715
   0.500   0.1995   0.01468   0.00916  -0.0387   0.9238   0.9750
   0.750   0.2359   0.01468   0.00918  -0.0411   0.9203   0.9777
   1.000   0.2778   0.01464   0.00918  -0.0444   0.9179   0.9796
   1.250   0.3224   0.01456   0.00917  -0.0482   0.9159   0.9808
   1.500   0.3561   0.01456   0.00923  -0.0499   0.9112   0.9837
   1.750   0.4004   0.01418   0.00895  -0.0530   0.9041   0.9843
   2.000   0.4482   0.01313   0.00801  -0.0553   0.8894   0.9834
   2.250   0.4844   0.01185   0.00677  -0.0545   0.8640   0.9844
   2.500   0.5146   0.01079   0.00570  -0.0528   0.8316   0.9869
   2.750   0.5405   0.01014   0.00503  -0.0511   0.7825   0.9908
   3.000   0.5484   0.01083   0.00424  -0.0456   0.4496   0.9959
   3.250   0.5501   0.01378   0.00528  -0.0424   0.0758   1.0000
   3.500   0.5635   0.01433   0.00582  -0.0400   0.0610   1.0000
   3.750   0.5734   0.01510   0.00659  -0.0369   0.0546   1.0000
   4.000   0.5850   0.01565   0.00722  -0.0340   0.0512   1.0000
   4.250   0.5952   0.01631   0.00790  -0.0309   0.0482   1.0000
   4.500   0.6051   0.01704   0.00860  -0.0278   0.0448   1.0000
   4.750   0.6175   0.01854   0.01007  -0.0252   0.0421   1.0000
   5.000   0.6349   0.01944   0.01108  -0.0231   0.0410   1.0000
   5.250   0.6563   0.02074   0.01248  -0.0218   0.0404   1.0000
   5.500   0.6811   0.02248   0.01438  -0.0209   0.0403   1.0000
   5.750   0.7068   0.02481   0.01692  -0.0202   0.0409   1.0000
   6.000   0.7299   0.02640   0.01879  -0.0191   0.0396   1.0000
   6.250   0.7553   0.03068   0.02324  -0.0187   0.0428   1.0000
   7.500   0.8383   0.04685   0.04177  -0.0075   0.0588   1.0000
   7.750   0.8555   0.05248   0.04714  -0.0084   0.0557   1.0000
   8.000   0.8479   0.05377   0.04924  -0.0036   0.0508   1.0000
   8.250   0.8521   0.05687   0.05258  -0.0018   0.0476   1.0000
   8.500   0.8569   0.05997   0.05580  -0.0005   0.0456   1.0000
   8.750   0.8642   0.06347   0.05931   0.0002   0.0441   1.0000
   9.000   0.8644   0.07277   0.06851  -0.0007   0.0425   1.0000
   9.250   0.7674   0.06244   0.05903   0.0096   0.0450   1.0000
   9.500   0.7502   0.06643   0.06310   0.0108   0.0447   1.0000
   9.750   0.7135   0.07211   0.06894   0.0100   0.0456   1.0000
  10.000   0.6747   0.08069   0.07764   0.0050   0.0472   1.0000
  10.250   0.6508   0.08996   0.08690  -0.0016   0.0481   1.0000
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