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NACA 65-206 (naca65206-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA 65-206 (naca65206-il)
Reynolds number: 500,000
Max Cl/Cd: 56.73 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca65206-il-500000-n5.txt
Download as CSV file: xf-naca65206-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65-206                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.4956   0.15073   0.14856   0.0143   1.0000   0.0046
 -13.000  -0.4943   0.14834   0.14616   0.0129   1.0000   0.0046
  -9.500  -0.5485   0.10669   0.10436  -0.0023   1.0000   0.0046
  -8.000  -0.5433   0.08204   0.07985  -0.0132   1.0000   0.0042
  -7.750  -0.5451   0.07828   0.07607  -0.0157   1.0000   0.0042
  -7.500  -0.5457   0.07323   0.07104  -0.0210   1.0000   0.0040
  -7.250  -0.5408   0.06848   0.06625  -0.0253   1.0000   0.0038
  -7.000  -0.5341   0.06360   0.06132  -0.0289   1.0000   0.0037
  -6.750  -0.5251   0.05879   0.05642  -0.0318   1.0000   0.0036
  -6.500  -0.5141   0.05402   0.05154  -0.0340   1.0000   0.0034
  -6.250  -0.5014   0.04914   0.04651  -0.0355   1.0000   0.0033
  -6.000  -0.4868   0.04444   0.04162  -0.0363   1.0000   0.0031
  -5.750  -0.4710   0.03973   0.03669  -0.0365   1.0000   0.0031
  -5.500  -0.4539   0.03525   0.03194  -0.0361   1.0000   0.0030
  -5.250  -0.4359   0.03078   0.02717  -0.0352   1.0000   0.0030
  -5.000  -0.4042   0.02531   0.02121  -0.0365   0.9961   0.0032
  -4.750  -0.3721   0.01998   0.01530  -0.0371   0.9920   0.0038
  -4.500  -0.3396   0.01716   0.01211  -0.0378   0.9886   0.0045
  -4.250  -0.3083   0.01480   0.00937  -0.0389   0.9852   0.0056
  -4.000  -0.2772   0.01300   0.00730  -0.0396   0.9809   0.0059
  -3.750  -0.2450   0.01170   0.00583  -0.0405   0.9772   0.0068
  -3.500  -0.2146   0.01071   0.00471  -0.0410   0.9714   0.0086
  -3.250  -0.1818   0.01054   0.00445  -0.0422   0.9669   0.0103
  -3.000  -0.1529   0.00938   0.00318  -0.0427   0.9597   0.0142
  -2.750  -0.1218   0.00896   0.00269  -0.0435   0.9535   0.0164
  -2.500  -0.0924   0.00863   0.00229  -0.0439   0.9453   0.0189
  -2.250  -0.0632   0.00831   0.00189  -0.0443   0.9374   0.0208
  -2.000  -0.0346   0.00803   0.00153  -0.0445   0.9290   0.0261
  -1.750  -0.0069   0.00785   0.00129  -0.0444   0.9195   0.0354
  -1.500   0.0200   0.00729   0.00111  -0.0446   0.9101   0.1549
  -1.250   0.0434   0.00569   0.00094  -0.0448   0.9005   0.5951
  -1.000   0.0673   0.00527   0.00099  -0.0439   0.8906   0.7387
  -0.750   0.0916   0.00512   0.00102  -0.0429   0.8807   0.8020
  -0.500   0.1163   0.00506   0.00103  -0.0420   0.8711   0.8388
  -0.250   0.1427   0.00506   0.00101  -0.0416   0.8614   0.8521
   0.000   0.1691   0.00506   0.00100  -0.0412   0.8510   0.8642
   0.250   0.1955   0.00507   0.00100  -0.0408   0.8406   0.8760
   0.500   0.2216   0.00508   0.00103  -0.0403   0.8301   0.8879
   0.750   0.2474   0.00510   0.00105  -0.0398   0.8192   0.8997
   1.000   0.2726   0.00514   0.00106  -0.0390   0.7994   0.9119
   1.250   0.2967   0.00523   0.00105  -0.0380   0.7654   0.9246
   1.750   0.3394   0.00613   0.00108  -0.0350   0.5329   0.9537
   2.000   0.3608   0.00775   0.00144  -0.0348   0.2064   0.9718
   2.500   0.4171   0.00895   0.00199  -0.0359   0.0306   1.0000
   2.750   0.4440   0.00924   0.00231  -0.0358   0.0243   1.0000
   3.000   0.4709   0.00953   0.00269  -0.0357   0.0212   1.0000
   3.250   0.4975   0.00986   0.00307  -0.0356   0.0187   1.0000
   3.500   0.5237   0.01030   0.00356  -0.0354   0.0167   1.0000
   3.750   0.5469   0.01144   0.00484  -0.0346   0.0138   1.0000
   4.000   0.5742   0.01156   0.00499  -0.0347   0.0126   1.0000
   4.250   0.6000   0.01206   0.00556  -0.0344   0.0107   1.0000
   4.500   0.6255   0.01270   0.00627  -0.0340   0.0087   1.0000
   4.750   0.6499   0.01363   0.00731  -0.0336   0.0061   1.0000
   5.000   0.6757   0.01419   0.00797  -0.0333   0.0053   1.0000
   5.250   0.7002   0.01551   0.00947  -0.0326   0.0043   1.0000
   5.500   0.7246   0.01725   0.01144  -0.0318   0.0037   1.0000
   5.750   0.7483   0.01974   0.01428  -0.0308   0.0034   1.0000
   6.000   0.7691   0.02459   0.01972  -0.0289   0.0034   1.0000
   6.250   0.7863   0.03090   0.02666  -0.0264   0.0035   1.0000
   6.500   0.8010   0.03712   0.03339  -0.0242   0.0037   1.0000
   6.750   0.8138   0.04277   0.03943  -0.0225   0.0039   1.0000
   7.000   0.8246   0.04817   0.04514  -0.0213   0.0041   1.0000
   7.250   0.8325   0.05358   0.05082  -0.0205   0.0043   1.0000
   7.500   0.8378   0.05881   0.05627  -0.0201   0.0045   1.0000
   7.750   0.8411   0.06387   0.06151  -0.0201   0.0044   1.0000
   8.000   0.8408   0.06903   0.06681  -0.0207   0.0043   1.0000
   8.250   0.8346   0.07419   0.07208  -0.0218   0.0047   1.0000
   8.500   0.8255   0.07891   0.07688  -0.0230   0.0045   1.0000
   8.750   0.8106   0.08389   0.08190  -0.0259   0.0048   1.0000
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