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USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 40 AIRFOIL (usa40-il)
Reynolds number: 50,000
Max Cl/Cd: 23.34 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa40-il-50000.txt
Download as CSV file: xf-usa40-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 40 AIRFOIL                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.2618   0.10912   0.10207  -0.0260   1.0000   0.2561
  -8.250  -0.2560   0.10670   0.09972  -0.0250   1.0000   0.2640
  -8.000  -0.2988   0.10920   0.10247  -0.0240   1.0000   0.2698
  -7.750  -0.2579   0.10230   0.09553  -0.0229   1.0000   0.2767
  -7.500  -0.2722   0.10182   0.09520  -0.0210   1.0000   0.2860
  -7.250  -0.2778   0.09963   0.09314  -0.0191   1.0000   0.2913
  -7.000  -0.2720   0.09742   0.09099  -0.0168   1.0000   0.3014
  -6.750  -0.3213   0.09964   0.09347  -0.0120   1.0000   0.3060
  -6.500  -0.2876   0.09443   0.08825  -0.0111   1.0000   0.3163
  -6.250  -0.3270   0.09567   0.08969  -0.0060   1.0000   0.3232
  -6.000  -0.3257   0.09300   0.08712  -0.0035   1.0000   0.3296
  -5.750  -0.3387   0.09223   0.08646   0.0002   1.0000   0.3388
  -5.500  -0.3656   0.09168   0.08608   0.0029   1.0000   0.3457
  -5.250  -0.3599   0.08951   0.08396   0.0060   1.0000   0.3551
  -5.000  -0.3820   0.08862   0.08321   0.0079   1.0000   0.3649
  -4.750  -0.3830   0.08697   0.08163   0.0107   1.0000   0.3760
  -4.500  -0.3903   0.08514   0.07989   0.0127   1.0000   0.3863
  -4.250  -0.4024   0.08386   0.07871   0.0138   1.0000   0.4019
  -4.000  -0.3228   0.05494   0.04826  -0.0505   1.0000   0.1817
  -3.750  -0.2939   0.05086   0.04384  -0.0553   1.0000   0.1749
  -3.500  -0.2641   0.04736   0.03997  -0.0597   1.0000   0.1725
  -3.250  -0.1926   0.04276   0.03459  -0.0713   0.9889   0.1701
  -3.000  -0.1281   0.04001   0.03118  -0.0799   0.9767   0.1744
  -2.750  -0.0692   0.03810   0.02883  -0.0867   0.9634   0.1836
  -2.500  -0.0125   0.03674   0.02702  -0.0925   0.9494   0.1973
  -2.250   0.0390   0.03558   0.02586  -0.0972   0.9349   0.2183
  -2.000   0.0903   0.03456   0.02484  -0.1015   0.9203   0.2535
  -1.750   0.1424   0.03345   0.02411  -0.1055   0.9062   0.3167
  -1.500   0.1898   0.03217   0.02353  -0.1082   0.8914   0.4529
  -1.250   0.2199   0.03177   0.02392  -0.1068   0.8751   0.5908
  -1.000   0.2394   0.03188   0.02440  -0.1027   0.8586   0.7037
  -0.750   0.2495   0.03170   0.02441  -0.0967   0.8426   0.8065
  -0.500   0.2622   0.03066   0.02359  -0.0906   0.8274   0.9505
  -0.250   0.3246   0.03057   0.02301  -0.0984   0.8096   1.0000
   0.000   0.3741   0.03059   0.02268  -0.1024   0.7936   1.0000
   0.250   0.4206   0.03052   0.02231  -0.1052   0.7791   1.0000
   0.500   0.4681   0.03021   0.02174  -0.1076   0.7659   1.0000
   0.750   0.5012   0.03045   0.02179  -0.1083   0.7494   1.0000
   1.000   0.5322   0.03081   0.02197  -0.1087   0.7335   1.0000
   1.250   0.5625   0.03119   0.02220  -0.1090   0.7181   1.0000
   1.500   0.5937   0.03152   0.02238  -0.1092   0.7037   1.0000
   1.750   0.6382   0.03096   0.02163  -0.1101   0.6941   1.0000
   2.000   0.6618   0.03175   0.02232  -0.1097   0.6787   1.0000
   2.250   0.6847   0.03264   0.02313  -0.1092   0.6640   1.0000
   2.500   0.7108   0.03337   0.02377  -0.1090   0.6513   1.0000
   2.750   0.7487   0.03317   0.02344  -0.1092   0.6416   1.0000
   3.000   0.7653   0.03464   0.02489  -0.1084   0.6274   1.0000
   3.250   0.7876   0.03576   0.02596  -0.1080   0.6158   1.0000
   3.500   0.8214   0.03588   0.02598  -0.1079   0.6064   1.0000
   3.750   0.8319   0.03802   0.02816  -0.1071   0.5935   1.0000
   4.000   0.8743   0.03746   0.02748  -0.1073   0.5866   1.0000
   4.250   0.8746   0.04055   0.03067  -0.1061   0.5735   1.0000
   4.500   0.8974   0.04174   0.03185  -0.1056   0.5647   1.0000
   4.750   0.9110   0.04373   0.03387  -0.1049   0.5546   1.0000
   5.000   0.9153   0.04676   0.03696  -0.1041   0.5451   1.0000
   5.250   0.9333   0.04843   0.03865  -0.1036   0.5366   1.0000
   5.500   0.9124   0.05407   0.04437  -0.1026   0.5265   1.0000
   5.750   0.9236   0.05656   0.04690  -0.1021   0.5190   1.0000
   6.000   0.8779   0.06530   0.05567  -0.1024   0.5111   1.0000
   6.250   0.8500   0.07208   0.06247  -0.1030   0.5049   1.0000
   6.500   0.9084   0.06993   0.06036  -0.1019   0.4987   1.0000
   6.750   0.8640   0.07887   0.06932  -0.1032   0.4961   1.0000
   7.000   0.8417   0.08523   0.07571  -0.1042   0.4952   1.0000
   7.250   0.8311   0.09047   0.08098  -0.1051   0.4951   1.0000
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