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R140 (original) (r140-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: R140 (original) (r140-il)
Reynolds number: 50,000
Max Cl/Cd: 27.11 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-r140-il-50000.txt
Download as CSV file: xf-r140-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: R140 (original)                                 
  
 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  -0.4915   0.10439   0.09776  -0.0075   1.0000   0.3400
  -9.250  -0.5942   0.08743   0.08114  -0.0343   1.0000   0.1721
  -9.000  -0.6224   0.07935   0.07302  -0.0390   1.0000   0.1355
  -8.750  -0.6574   0.07589   0.06949  -0.0376   1.0000   0.1283
  -8.500  -0.6972   0.07437   0.06772  -0.0343   1.0000   0.1305
  -8.250  -0.6972   0.06737   0.06049  -0.0333   1.0000   0.1057
  -8.000  -0.6953   0.06273   0.05568  -0.0313   1.0000   0.0956
  -7.750  -0.7015   0.05919   0.05174  -0.0279   1.0000   0.0881
  -7.500  -0.7017   0.05536   0.04766  -0.0249   1.0000   0.0840
  -7.250  -0.7099   0.05275   0.04421  -0.0196   1.0000   0.0781
  -7.000  -0.6995   0.04890   0.04023  -0.0175   1.0000   0.0765
  -6.750  -0.6920   0.04555   0.03654  -0.0143   1.0000   0.0753
  -6.500  -0.6820   0.04240   0.03299  -0.0111   1.0000   0.0747
  -6.250  -0.6676   0.03952   0.02970  -0.0085   1.0000   0.0747
  -6.000  -0.6498   0.03677   0.02649  -0.0060   1.0000   0.0764
  -5.750  -0.6273   0.03391   0.02345  -0.0044   1.0000   0.0825
  -5.500  -0.6027   0.03148   0.02092  -0.0031   1.0000   0.0969
  -5.250  -0.5867   0.02921   0.01887  -0.0011   1.0000   0.1355
  -5.000  -0.5605   0.02650   0.01637   0.0004   1.0000   0.1855
  -4.750  -0.5499   0.02429   0.01459   0.0034   1.0000   0.2335
  -4.500  -0.5512   0.02197   0.01330   0.0080   1.0000   0.3167
  -4.250  -0.5701   0.02096   0.01431   0.0199   1.0000   0.6562
  -4.000  -0.2193   0.03183   0.02234  -0.0087   1.0000   0.9181
  -3.750  -0.1815   0.03086   0.02109  -0.0121   1.0000   0.9337
  -3.500  -0.1462   0.02991   0.01992  -0.0153   1.0000   0.9479
  -3.250  -0.1103   0.02895   0.01877  -0.0187   1.0000   0.9614
  -3.000  -0.0728   0.02798   0.01756  -0.0226   1.0000   0.9745
  -2.750  -0.0335   0.02699   0.01642  -0.0270   1.0000   0.9869
  -2.500   0.0108   0.02595   0.01524  -0.0325   1.0000   0.9996
  -2.250   0.0169   0.02564   0.01494  -0.0307   1.0000   1.0000
  -2.000   0.0180   0.02548   0.01481  -0.0281   1.0000   1.0000
  -1.750   0.0144   0.02546   0.01482  -0.0248   1.0000   1.0000
  -1.500   0.0060   0.02556   0.01491  -0.0209   1.0000   1.0000
  -1.250  -0.0054   0.02572   0.01509  -0.0166   1.0000   1.0000
  -1.000  -0.0178   0.02587   0.01524  -0.0124   1.0000   1.0000
  -0.750  -0.0301   0.02596   0.01531  -0.0082   1.0000   1.0000
  -0.500  -0.0420   0.02596   0.01530  -0.0042   1.0000   1.0000
  -0.250  -0.0535   0.02590   0.01521  -0.0001   1.0000   1.0000
   0.000  -0.0640   0.02580   0.01507   0.0040   1.0000   1.0000
   0.250  -0.0723   0.02570   0.01492   0.0079   1.0000   1.0000
   0.500  -0.0775   0.02565   0.01480   0.0115   1.0000   1.0000
   0.750  -0.0787   0.02569   0.01476   0.0146   1.0000   1.0000
   1.000  -0.0760   0.02583   0.01482   0.0171   1.0000   1.0000
   1.500  -0.0115   0.02722   0.01615   0.0112   0.9837   1.0000
   1.750   0.0181   0.02790   0.01682   0.0088   0.9747   1.0000
   2.000   0.0493   0.02872   0.01764   0.0063   0.9664   1.0000
   2.250   0.0748   0.02944   0.01836   0.0049   0.9581   1.0000
   2.500   0.0985   0.03019   0.01914   0.0039   0.9504   1.0000
   2.750   0.1304   0.03115   0.02021   0.0015   0.9421   1.0000
   3.000   0.1496   0.03184   0.02095   0.0014   0.9328   1.0000
   3.250   0.1746   0.03270   0.02189   0.0003   0.9233   1.0000
   3.500   0.2041   0.03364   0.02295  -0.0015   0.9126   1.0000
   3.750   0.2353   0.03458   0.02405  -0.0035   0.9004   1.0000
   4.000   0.2645   0.03545   0.02516  -0.0049   0.8860   1.0000
   4.250   0.2921   0.03633   0.02623  -0.0061   0.8718   1.0000
   4.500   0.3190   0.03728   0.02738  -0.0071   0.8589   1.0000
   4.750   0.3448   0.03822   0.02855  -0.0079   0.8454   1.0000
   5.000   0.3738   0.03912   0.02973  -0.0089   0.8285   1.0000
   5.250   0.4189   0.03965   0.03077  -0.0114   0.8034   1.0000
   5.500   0.6392   0.02358   0.01438  -0.0064   0.2407   1.0000
   5.750   0.6222   0.02667   0.01626   0.0002   0.1486   1.0000
   6.000   0.6158   0.02930   0.01834   0.0055   0.1121   1.0000
   6.250   0.6390   0.03180   0.02069   0.0075   0.0878   1.0000
   6.500   0.6936   0.03483   0.02381   0.0055   0.0770   1.0000
   6.750   0.7310   0.03833   0.02752   0.0049   0.0720   1.0000
   7.000   0.7515   0.04073   0.03041   0.0069   0.0698   1.0000
   7.250   0.7682   0.04339   0.03349   0.0092   0.0681   1.0000
   7.500   0.7809   0.04620   0.03669   0.0118   0.0667   1.0000
   7.750   0.7909   0.04931   0.04019   0.0145   0.0673   1.0000
   8.000   0.7967   0.05256   0.04382   0.0176   0.0685   1.0000
   8.250   0.8016   0.05621   0.04776   0.0203   0.0705   1.0000
   8.500   0.8148   0.06109   0.05267   0.0214   0.0727   1.0000
   8.750   0.7944   0.06274   0.05514   0.0271   0.0774   1.0000
   9.000   0.7837   0.06655   0.05924   0.0302   0.0814   1.0000
   9.250   0.7913   0.07115   0.06395   0.0315   0.0874   1.0000
   9.500   0.7570   0.07368   0.06681   0.0354   0.0899   1.0000
   9.750   0.7291   0.07700   0.07024   0.0380   0.0919   1.0000
  10.000   0.7058   0.08110   0.07440   0.0385   0.0942   1.0000
  10.250   0.6949   0.08615   0.07952   0.0376   0.1013   1.0000
  10.500   0.6561   0.09270   0.08608   0.0330   0.1036   1.0000
  10.750   0.6392   0.10192   0.09527   0.0272   0.1228   1.0000
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