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NACA 25112 (naca25112-jf) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 25112 (naca25112-jf)
Reynolds number: 50,000
Max Cl/Cd: 17.68 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca25112-jf-50000.txt
Download as CSV file: xf-naca25112-jf-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 25112                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4517   0.09133   0.08486   0.0130   1.0000   0.3743
  -7.500  -0.4336   0.08811   0.08166   0.0156   1.0000   0.4064
  -7.250  -0.4327   0.08599   0.07962   0.0184   1.0000   0.4427
  -7.000  -0.3976   0.08218   0.07579   0.0212   1.0000   0.4835
  -6.750  -0.3720   0.07941   0.07302   0.0244   1.0000   0.5356
  -6.500  -0.3417   0.07660   0.07020   0.0277   1.0000   0.5993
  -6.250  -0.3125   0.07376   0.06737   0.0297   1.0000   0.6583
  -5.500  -0.4830   0.04731   0.03948  -0.0107   1.0000   0.1825
  -5.250  -0.4739   0.04424   0.03611  -0.0087   1.0000   0.1769
  -5.000  -0.4661   0.04169   0.03305  -0.0062   1.0000   0.1730
  -4.750  -0.4567   0.03998   0.03079  -0.0036   1.0000   0.1707
  -4.500  -0.4460   0.03797   0.02853  -0.0014   1.0000   0.1705
  -4.250  -0.4349   0.03599   0.02654   0.0002   1.0000   0.1735
  -4.000  -0.4233   0.03464   0.02511   0.0018   1.0000   0.1775
  -3.750  -0.4098   0.03340   0.02363   0.0032   1.0000   0.1805
  -3.500  -0.3947   0.03240   0.02230   0.0042   1.0000   0.1849
  -3.250  -0.3783   0.03136   0.02111   0.0049   1.0000   0.1911
  -3.000  -0.3207   0.02987   0.01930  -0.0013   0.9854   0.2062
  -2.750  -0.2555   0.02829   0.01775  -0.0084   0.9691   0.2298
  -2.500  -0.1894   0.02688   0.01642  -0.0153   0.9524   0.2669
  -2.250   0.0152   0.02311   0.01531  -0.0393   0.9647   1.0000
  -2.000   0.0972   0.02291   0.01457  -0.0501   0.9426   1.0000
  -1.750   0.1591   0.02274   0.01406  -0.0569   0.9180   1.0000
  -1.500   0.2046   0.02269   0.01376  -0.0603   0.8936   1.0000
  -1.250   0.2378   0.02278   0.01364  -0.0612   0.8706   1.0000
  -1.000   0.2621   0.02305   0.01374  -0.0606   0.8484   1.0000
  -0.750   0.2851   0.02333   0.01387  -0.0597   0.8284   1.0000
  -0.500   0.0507   0.00674   0.00137   0.0051   0.4333   0.6298
  -0.250   0.3280   0.02405   0.01432  -0.0574   0.7933   1.0000
   0.000   0.3487   0.02448   0.01464  -0.0561   0.7773   1.0000
   0.250   0.3692   0.02497   0.01502  -0.0549   0.7624   1.0000
   0.500   0.3895   0.02545   0.01541  -0.0536   0.7483   1.0000
   0.750   0.4105   0.02586   0.01570  -0.0521   0.7357   1.0000
   1.000   0.4299   0.02651   0.01631  -0.0510   0.7220   1.0000
   1.250   0.4486   0.02725   0.01701  -0.0499   0.7085   1.0000
   1.500   0.4680   0.02795   0.01765  -0.0487   0.6965   1.0000
   1.750   0.4892   0.02841   0.01803  -0.0471   0.6860   1.0000
   2.000   0.5060   0.02942   0.01906  -0.0461   0.6730   1.0000
   2.250   0.5238   0.03035   0.01997  -0.0449   0.6617   1.0000
   2.500   0.5453   0.03085   0.02040  -0.0433   0.6521   1.0000
   2.750   0.5593   0.03217   0.02176  -0.0423   0.6397   1.0000
   3.000   0.5765   0.03320   0.02280  -0.0410   0.6296   1.0000
   3.250   0.5958   0.03400   0.02358  -0.0396   0.6195   1.0000
   3.500   0.6067   0.03564   0.02527  -0.0384   0.6082   1.0000
   3.750   0.6311   0.03603   0.02561  -0.0367   0.6001   1.0000
   4.000   0.6374   0.03809   0.02777  -0.0355   0.5882   1.0000
   4.250   0.6487   0.03973   0.02944  -0.0341   0.5783   1.0000
   4.500   0.6679   0.04066   0.03038  -0.0327   0.5690   1.0000
   4.750   0.6670   0.04335   0.03313  -0.0312   0.5581   1.0000
   5.000   0.6992   0.04326   0.03305  -0.0297   0.5504   1.0000
   5.250   0.6829   0.04731   0.03718  -0.0283   0.5390   1.0000
   5.500   0.6974   0.04883   0.03873  -0.0268   0.5299   1.0000
   5.750   0.6967   0.05166   0.04160  -0.0256   0.5202   1.0000
   6.000   0.6888   0.05520   0.04516  -0.0245   0.5111   1.0000
   6.250   0.7108   0.05622   0.04624  -0.0233   0.5017   1.0000
   6.500   0.6759   0.06204   0.05201  -0.0227   0.4943   1.0000
   6.750   0.7328   0.06034   0.05048  -0.0211   0.4833   1.0000
   7.000   0.6711   0.06831   0.05831  -0.0211   0.4788   1.0000
   7.250   0.6597   0.07226   0.06226  -0.0208   0.4726   1.0000
   7.500   0.6741   0.07441   0.06449  -0.0202   0.4632   1.0000
   7.750   0.6567   0.07910   0.06918  -0.0206   0.4599   1.0000
   8.000   0.6824   0.08049   0.07065  -0.0197   0.4479   1.0000
   8.250   0.6646   0.08541   0.07558  -0.0205   0.4460   1.0000
   8.500   0.6568   0.08991   0.08012  -0.0214   0.4458   1.0000
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