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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 1,000,000
Max Cl/Cd: 78.03 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa49-il-1000000.txt
Download as CSV file: xf-usa49-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 49 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6084   0.08854   0.08702   0.0057   1.0000   0.0079
  -8.250  -0.6089   0.08385   0.08234   0.0022   1.0000   0.0081
  -8.000  -0.6109   0.07851   0.07703  -0.0033   1.0000   0.0083
  -7.750  -0.6065   0.07187   0.07035  -0.0112   1.0000   0.0086
  -5.000  -0.4529   0.01740   0.01309  -0.0202   1.0000   0.0095
  -4.750  -0.4280   0.01422   0.00950  -0.0195   0.9992   0.0088
  -4.500  -0.3957   0.01273   0.00779  -0.0203   0.9975   0.0092
  -4.250  -0.3622   0.01209   0.00704  -0.0214   0.9959   0.0098
  -4.000  -0.3291   0.01107   0.00590  -0.0225   0.9945   0.0103
  -3.750  -0.2984   0.00972   0.00446  -0.0233   0.9921   0.0109
  -3.500  -0.2657   0.00911   0.00383  -0.0244   0.9896   0.0120
  -3.250  -0.2327   0.00863   0.00331  -0.0256   0.9866   0.0130
  -3.000  -0.1990   0.00819   0.00283  -0.0268   0.9835   0.0143
  -2.750  -0.1691   0.00791   0.00252  -0.0271   0.9767   0.0152
  -2.500  -0.1397   0.00748   0.00199  -0.0273   0.9682   0.0173
  -2.250  -0.1141   0.00724   0.00171  -0.0265   0.9565   0.0212
  -2.000  -0.0886   0.00697   0.00146  -0.0258   0.9456   0.0399
  -1.750  -0.0637   0.00651   0.00133  -0.0252   0.9349   0.1303
  -1.500  -0.0382   0.00622   0.00121  -0.0247   0.9233   0.1866
  -1.250  -0.0136   0.00560   0.00109  -0.0243   0.9101   0.3488
  -1.000   0.0110   0.00497   0.00098  -0.0238   0.8949   0.5203
  -0.750   0.0348   0.00438   0.00092  -0.0230   0.8765   0.6998
  -0.500   0.0555   0.00398   0.00092  -0.0211   0.8419   0.8473
  -0.250   0.0785   0.00402   0.00087  -0.0196   0.7887   0.8946
   0.000   0.1023   0.00414   0.00085  -0.0184   0.7355   0.9243
   0.250   0.1278   0.00428   0.00084  -0.0176   0.6901   0.9480
   0.500   0.1575   0.00446   0.00085  -0.0179   0.6414   0.9706
   0.750   0.1921   0.00470   0.00087  -0.0194   0.5873   0.9853
   1.000   0.2302   0.00493   0.00090  -0.0218   0.5419   0.9952
   1.250   0.2689   0.00511   0.00094  -0.0243   0.5091   1.0000
   1.500   0.2947   0.00525   0.00097  -0.0240   0.4799   1.0000
   1.750   0.3204   0.00541   0.00099  -0.0236   0.4438   1.0000
   2.000   0.3459   0.00568   0.00104  -0.0232   0.3894   1.0000
   2.250   0.3709   0.00606   0.00114  -0.0229   0.3213   1.0000
   2.500   0.3958   0.00656   0.00129  -0.0226   0.2374   1.0000
   2.750   0.4190   0.00755   0.00162  -0.0222   0.0791   1.0000
   3.000   0.4453   0.00783   0.00181  -0.0219   0.0616   1.0000
   3.250   0.4720   0.00805   0.00203  -0.0217   0.0549   1.0000
   3.500   0.4989   0.00821   0.00219  -0.0216   0.0535   1.0000
   3.750   0.5259   0.00840   0.00241  -0.0214   0.0515   1.0000
   4.000   0.5528   0.00862   0.00265  -0.0213   0.0496   1.0000
   4.250   0.5797   0.00887   0.00291  -0.0211   0.0480   1.0000
   4.500   0.6065   0.00914   0.00321  -0.0210   0.0467   1.0000
   4.750   0.6330   0.00947   0.00358  -0.0208   0.0449   1.0000
   5.000   0.6588   0.00998   0.00415  -0.0206   0.0419   1.0000
   5.250   0.6862   0.01006   0.00423  -0.0206   0.0411   1.0000
   5.500   0.7135   0.01016   0.00434  -0.0205   0.0395   1.0000
   5.750   0.7406   0.01031   0.00452  -0.0205   0.0374   1.0000
   6.000   0.7677   0.01047   0.00467  -0.0205   0.0350   1.0000
   6.250   0.7933   0.01097   0.00520  -0.0202   0.0308   1.0000
   6.500   0.8217   0.01080   0.00500  -0.0204   0.0292   1.0000
   6.750   0.8490   0.01088   0.00499  -0.0205   0.0223   1.0000
   7.000   0.8748   0.01131   0.00535  -0.0202   0.0151   1.0000
   7.250   0.9006   0.01172   0.00582  -0.0200   0.0132   1.0000
   7.500   0.9253   0.01236   0.00650  -0.0196   0.0110   1.0000
   7.750   0.9497   0.01305   0.00729  -0.0191   0.0098   1.0000
   8.000   0.9746   0.01356   0.00788  -0.0188   0.0091   1.0000
   8.250   0.9989   0.01418   0.00859  -0.0184   0.0084   1.0000
   8.500   1.0228   0.01485   0.00934  -0.0179   0.0078   1.0000
   8.750   1.0449   0.01581   0.01041  -0.0172   0.0072   1.0000
   9.000   1.0607   0.01799   0.01286  -0.0156   0.0066   1.0000
   9.250   1.0830   0.01887   0.01386  -0.0150   0.0064   1.0000
   9.500   1.1052   0.01971   0.01482  -0.0144   0.0062   1.0000
   9.750   1.1259   0.02079   0.01605  -0.0136   0.0059   1.0000
  10.000   1.1457   0.02197   0.01738  -0.0127   0.0056   1.0000
  10.250   1.1653   0.02308   0.01865  -0.0119   0.0053   1.0000
  10.500   1.1847   0.02411   0.01980  -0.0111   0.0050   1.0000
  10.750   1.2028   0.02524   0.02105  -0.0102   0.0048   1.0000
  11.000   1.2180   0.02672   0.02268  -0.0091   0.0046   1.0000
  11.250   1.2290   0.02869   0.02485  -0.0076   0.0045   1.0000
  11.500   1.2340   0.03128   0.02769  -0.0057   0.0044   1.0000
  11.750   1.2313   0.03445   0.03115  -0.0033   0.0043   1.0000
  12.000   1.2195   0.03762   0.03458  -0.0001   0.0042   1.0000
  12.250   1.2041   0.04114   0.03834   0.0020   0.0042   1.0000
  12.500   1.1875   0.04549   0.04291   0.0020   0.0042   1.0000
  12.750   1.1720   0.05080   0.04842  -0.0005   0.0042   1.0000
  13.000   1.1561   0.05748   0.05529  -0.0053   0.0042   1.0000
  13.250   1.1415   0.06497   0.06295  -0.0112   0.0042   1.0000
  13.500   1.1242   0.07355   0.07167  -0.0175   0.0043   1.0000
  13.750   1.1058   0.08281   0.08106  -0.0237   0.0043   1.0000
  14.000   1.0772   0.09463   0.09303  -0.0306   0.0044   1.0000
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