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USA 40 B AIRFOIL (usa40b-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 40 B AIRFOIL (usa40b-il)
Reynolds number: 50,000
Max Cl/Cd: 25.06 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa40b-il-50000.txt
Download as CSV file: xf-usa40b-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 40 B 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2811   0.10872   0.10205  -0.0191   1.0000   0.2697
  -7.750  -0.2365   0.10155   0.09485  -0.0176   1.0000   0.2773
  -7.500  -0.2541   0.10148   0.09493  -0.0164   1.0000   0.2872
  -7.250  -0.2423   0.09776   0.09130  -0.0151   1.0000   0.2934
  -7.000  -0.2459   0.09641   0.09007  -0.0132   1.0000   0.3043
  -6.750  -0.2582   0.09485   0.08867  -0.0111   1.0000   0.3104
  -6.500  -0.2476   0.09229   0.08619  -0.0090   1.0000   0.3205
  -6.250  -0.3067   0.09511   0.08930  -0.0036   1.0000   0.3262
  -6.000  -0.2641   0.08933   0.08351  -0.0033   1.0000   0.3356
  -5.750  -0.3066   0.09076   0.08516   0.0022   1.0000   0.3439
  -5.500  -0.3107   0.08850   0.08303   0.0051   1.0000   0.3497
  -5.250  -0.3160   0.08720   0.08182   0.0084   1.0000   0.3588
  -5.000  -0.3492   0.08711   0.08191   0.0116   1.0000   0.3673
  -4.750  -0.3422   0.08493   0.07979   0.0147   1.0000   0.3767
  -4.500  -0.3633   0.08400   0.07900   0.0171   1.0000   0.3880
  -4.250  -0.3863   0.08377   0.07889   0.0183   1.0000   0.4044
  -4.000  -0.3721   0.08084   0.07602   0.0224   1.0000   0.4117
  -3.750  -0.3235   0.05288   0.04647  -0.0383   1.0000   0.1803
  -3.500  -0.2883   0.04853   0.04166  -0.0443   0.9966   0.1734
  -3.250  -0.2204   0.04388   0.03636  -0.0550   0.9809   0.1698
  -3.000  -0.1585   0.04036   0.03220  -0.0631   0.9632   0.1695
  -2.750  -0.0964   0.03797   0.02906  -0.0701   0.9445   0.1765
  -2.500  -0.0364   0.03603   0.02705  -0.0759   0.9253   0.1862
  -2.250   0.0193   0.03439   0.02520  -0.0804   0.9042   0.2027
  -2.000   0.0716   0.03287   0.02358  -0.0839   0.8816   0.2279
  -1.750   0.1262   0.03114   0.02203  -0.0873   0.8598   0.2738
  -1.500   0.1841   0.02909   0.02055  -0.0904   0.8394   0.3823
  -1.250   0.2292   0.02728   0.01982  -0.0899   0.8191   0.5622
  -1.000   0.2592   0.02653   0.01954  -0.0853   0.7986   0.6999
  -0.750   0.2728   0.02586   0.01902  -0.0784   0.7756   0.8064
  -0.500   0.4068   0.02315   0.01601  -0.0903   0.7437   1.0000
  -0.250   0.4186   0.02319   0.01573  -0.0885   0.7155   1.0000
   0.000   0.4466   0.02320   0.01534  -0.0883   0.6889   1.0000
   0.250   0.4759   0.02338   0.01513  -0.0881   0.6639   1.0000
   0.500   0.5002   0.02383   0.01530  -0.0875   0.6390   1.0000
   0.750   0.5275   0.02422   0.01539  -0.0872   0.6174   1.0000
   1.000   0.5549   0.02466   0.01554  -0.0869   0.5979   1.0000
   1.250   0.5819   0.02518   0.01581  -0.0865   0.5803   1.0000
   1.500   0.6086   0.02576   0.01616  -0.0862   0.5643   1.0000
   1.750   0.6363   0.02633   0.01649  -0.0859   0.5500   1.0000
   2.000   0.6652   0.02685   0.01675  -0.0858   0.5370   1.0000
   2.250   0.6872   0.02783   0.01767  -0.0852   0.5238   1.0000
   2.500   0.7128   0.02870   0.01839  -0.0848   0.5130   1.0000
   2.750   0.7390   0.02949   0.01902  -0.0845   0.5025   1.0000
   3.000   0.7601   0.03069   0.02020  -0.0839   0.4923   1.0000
   3.250   0.7874   0.03149   0.02084  -0.0837   0.4837   1.0000
   3.500   0.8056   0.03304   0.02244  -0.0830   0.4757   1.0000
   3.750   0.8281   0.03427   0.02363  -0.0825   0.4682   1.0000
   4.000   0.8537   0.03539   0.02463  -0.0822   0.4618   1.0000
   4.250   0.8627   0.03763   0.02707  -0.0810   0.4543   1.0000
   4.500   0.8874   0.03877   0.02812  -0.0806   0.4482   1.0000
   4.750   0.9073   0.04048   0.02982  -0.0801   0.4433   1.0000
   5.000   0.9016   0.04422   0.03384  -0.0785   0.4385   1.0000
   5.250   0.8996   0.04776   0.03752  -0.0773   0.4342   1.0000
   5.500   0.9132   0.05003   0.03979  -0.0766   0.4300   1.0000
   5.750   0.9426   0.05116   0.04084  -0.0765   0.4261   1.0000
   6.000   0.8712   0.06113   0.05112  -0.0745   0.4249   1.0000
   6.250   0.8208   0.07012   0.06019  -0.0748   0.4269   1.0000
   6.500   0.5459   0.10135   0.09187  -0.0925   0.5976   1.0000
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