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NACA 0012-64 a=0.8 c(li)=0.2 (naca001264a08cli02-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0012-64 a=0.8 c(li)=0.2 (naca001264a08cli02-il)
Reynolds number: 50,000
Max Cl/Cd: 33.04 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001264a08cli02-il-50000.txt
Download as CSV file: xf-naca001264a08cli02-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0012-64 a=0.8 c(li)=0.2                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4561   0.11087   0.10380  -0.0158   1.0000   0.3338
  -9.250  -0.4615   0.10848   0.10149  -0.0146   1.0000   0.3508
  -9.000  -0.4645   0.10570   0.09880  -0.0132   1.0000   0.3681
  -8.750  -0.4409   0.10141   0.09449  -0.0115   1.0000   0.3907
  -8.500  -0.4500   0.09939   0.09257  -0.0092   1.0000   0.4128
  -8.250  -0.4410   0.09655   0.08976  -0.0069   1.0000   0.4395
  -8.000  -0.4279   0.09334   0.08659  -0.0046   1.0000   0.4682
  -7.750  -0.4297   0.09143   0.08475  -0.0011   1.0000   0.5000
  -7.500  -0.3963   0.08696   0.08023   0.0000   1.0000   0.5361
  -7.250  -0.3844   0.08399   0.07730   0.0021   1.0000   0.5665
  -6.000  -0.4625   0.07110   0.06502   0.0132   1.0000   0.5410
  -5.750  -0.5162   0.06898   0.06313   0.0186   1.0000   0.5192
  -5.500  -0.5696   0.06559   0.05994   0.0208   1.0000   0.4884
  -5.250  -0.6237   0.05982   0.05405   0.0157   1.0000   0.4255
  -5.000  -0.6078   0.04914   0.04137  -0.0001   1.0000   0.2178
  -4.750  -0.5859   0.04500   0.03635   0.0011   1.0000   0.1743
  -4.500  -0.5636   0.04220   0.03271   0.0032   1.0000   0.1485
  -4.250  -0.5415   0.03967   0.02959   0.0052   1.0000   0.1346
  -4.000  -0.5186   0.03756   0.02680   0.0072   1.0000   0.1240
  -3.750  -0.4941   0.03529   0.02427   0.0084   1.0000   0.1196
  -3.500  -0.4688   0.03347   0.02210   0.0095   1.0000   0.1182
  -3.250  -0.4423   0.03189   0.02023   0.0104   1.0000   0.1178
  -3.000  -0.4133   0.03041   0.01851   0.0111   1.0000   0.1163
  -2.750  -0.3820   0.02913   0.01708   0.0114   1.0000   0.1157
  -2.500  -0.3503   0.02812   0.01594   0.0115   1.0000   0.1175
  -2.250  -0.3227   0.02738   0.01504   0.0120   1.0000   0.1241
  -2.000  -0.2992   0.02661   0.01422   0.0129   1.0000   0.1325
  -1.750  -0.1121   0.02340   0.01403  -0.0137   1.0000   1.0000
  -1.500  -0.1014   0.02346   0.01379  -0.0109   1.0000   1.0000
  -1.250  -0.0902   0.02356   0.01362  -0.0083   1.0000   1.0000
  -1.000  -0.0784   0.02369   0.01350  -0.0059   1.0000   1.0000
  -0.750  -0.0663   0.02385   0.01342  -0.0035   1.0000   1.0000
  -0.500  -0.0537   0.02404   0.01340  -0.0014   1.0000   1.0000
  -0.250  -0.0408   0.02425   0.01343   0.0007   1.0000   1.0000
   0.000  -0.0276   0.02449   0.01350   0.0027   1.0000   1.0000
   0.250  -0.0141   0.02476   0.01363   0.0046   1.0000   1.0000
   0.500  -0.0003   0.02507   0.01379   0.0064   1.0000   1.0000
   0.750   0.0137   0.02540   0.01401   0.0081   1.0000   1.0000
   1.000   0.0278   0.02576   0.01428   0.0097   1.0000   1.0000
   1.250   0.0419   0.02616   0.01460   0.0112   1.0000   1.0000
   1.500   0.0562   0.02659   0.01496   0.0127   1.0000   1.0000
   1.750   0.0704   0.02706   0.01538   0.0141   1.0000   1.0000
   2.000   0.0847   0.02758   0.01586   0.0154   1.0000   1.0000
   2.250   0.0990   0.02813   0.01640   0.0166   1.0000   1.0000
   2.500   0.1131   0.02874   0.01700   0.0178   1.0000   1.0000
   2.750   0.1272   0.02940   0.01767   0.0188   1.0000   1.0000
   3.000   0.1411   0.03012   0.01841   0.0198   1.0000   1.0000
   3.250   0.1548   0.03091   0.01922   0.0206   1.0000   1.0000
   3.500   0.1683   0.03176   0.02012   0.0214   1.0000   1.0000
   3.750   0.1984   0.03307   0.02153   0.0187   0.9941   1.0000
   4.000   0.2732   0.03503   0.02372   0.0074   0.9629   1.0000
   4.250   0.3434   0.03639   0.02534  -0.0018   0.9305   1.0000
   4.500   0.4106   0.03716   0.02644  -0.0096   0.8992   1.0000
   4.750   0.4713   0.03730   0.02691  -0.0153   0.8666   1.0000
   5.000   0.5431   0.03636   0.02646  -0.0212   0.8286   1.0000
   5.250   0.6505   0.03237   0.02325  -0.0289   0.7787   1.0000
   5.500   0.7495   0.02627   0.01794  -0.0317   0.7084   1.0000
   5.750   0.7877   0.02384   0.01440  -0.0250   0.4666   1.0000
   6.000   0.7893   0.02619   0.01538  -0.0196   0.3384   1.0000
   6.250   0.8090   0.02851   0.01699  -0.0180   0.2632   1.0000
   6.500   0.8394   0.03066   0.01883  -0.0183   0.2169   1.0000
   6.750   0.8780   0.03308   0.02106  -0.0198   0.1920   1.0000
   7.000   0.9064   0.03523   0.02331  -0.0196   0.1771   1.0000
   7.250   0.9289   0.03715   0.02542  -0.0186   0.1654   1.0000
   7.500   0.9498   0.03960   0.02828  -0.0172   0.1592   1.0000
   7.750   0.9688   0.04218   0.03124  -0.0155   0.1549   1.0000
   8.000   0.9875   0.04498   0.03435  -0.0140   0.1518   1.0000
   8.250   1.0038   0.04827   0.03784  -0.0124   0.1491   1.0000
   8.500   1.0061   0.05111   0.04127  -0.0085   0.1471   1.0000
   8.750   1.0056   0.05418   0.04484  -0.0045   0.1451   1.0000
   9.000   1.0033   0.05764   0.04873  -0.0007   0.1447   1.0000
   9.250   0.9984   0.06145   0.05290   0.0030   0.1456   1.0000
   9.500   0.9920   0.06545   0.05719   0.0066   0.1470   1.0000
   9.750   0.9886   0.06977   0.06172   0.0094   0.1485   1.0000
  10.000   0.9488   0.07369   0.06604   0.0153   0.1526   1.0000
  10.250   0.9032   0.07813   0.07064   0.0203   0.1566   1.0000
  10.500   0.8733   0.08293   0.07549   0.0228   0.1600   1.0000
  10.750   0.8596   0.08842   0.08104   0.0233   0.1649   1.0000
  11.000   0.7717   0.10186   0.09431   0.0151   0.1810   1.0000
  11.250   0.6759   0.13167   0.12374  -0.0129   0.3545   1.0000
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