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NASA/LANGLEY LS(1)-0417MOD AIRFOIL (ls417mod-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY LS(1)-0417MOD AIRFOIL (ls417mod-il)
Reynolds number: 50,000
Max Cl/Cd: 24.32 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ls417mod-il-50000.txt
Download as CSV file: xf-ls417mod-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY LS(1)-0417MOD 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4771   0.09203   0.08484  -0.0483   1.0000   0.1609
  -9.750  -0.4697   0.08812   0.08095  -0.0477   1.0000   0.1577
  -9.500  -0.5706   0.07799   0.07088  -0.0503   1.0000   0.1449
  -9.250  -0.5953   0.07443   0.06737  -0.0481   1.0000   0.1426
  -9.000  -0.6325   0.07044   0.06331  -0.0456   1.0000   0.1397
  -8.750  -0.7024   0.06493   0.05714  -0.0425   1.0000   0.1339
  -8.500  -0.7027   0.06154   0.05363  -0.0407   1.0000   0.1329
  -8.250  -0.7046   0.05811   0.04996  -0.0391   1.0000   0.1323
  -8.000  -0.7042   0.05472   0.04622  -0.0378   1.0000   0.1324
  -7.750  -0.6993   0.05145   0.04249  -0.0368   1.0000   0.1332
  -7.500  -0.6888   0.04840   0.03898  -0.0362   1.0000   0.1343
  -7.250  -0.6719   0.04602   0.03663  -0.0352   1.0000   0.1366
  -7.000  -0.6543   0.04369   0.03409  -0.0344   1.0000   0.1386
  -6.750  -0.6352   0.04158   0.03172  -0.0337   1.0000   0.1420
  -6.500  -0.6144   0.03960   0.02926  -0.0334   1.0000   0.1469
  -6.250  -0.5938   0.03798   0.02780  -0.0323   1.0000   0.1527
  -6.000  -0.5708   0.03645   0.02597  -0.0316   1.0000   0.1595
  -5.750  -0.5493   0.03512   0.02481  -0.0304   1.0000   0.1688
  -5.500  -0.5276   0.03391   0.02369  -0.0291   1.0000   0.1811
  -5.250  -0.5067   0.03282   0.02274  -0.0274   1.0000   0.1980
  -5.000  -0.4859   0.03166   0.02177  -0.0258   1.0000   0.2263
  -4.750  -0.4657   0.02999   0.02076  -0.0249   1.0000   0.2833
  -4.500  -0.4538   0.02788   0.02084  -0.0213   1.0000   0.4742
  -4.250  -0.4547   0.03055   0.02401  -0.0107   1.0000   0.5919
  -4.000  -0.4494   0.03306   0.02642  -0.0027   1.0000   0.6449
  -3.750  -0.4448   0.03517   0.02845   0.0053   1.0000   0.6799
  -3.500  -0.4380   0.03675   0.02992   0.0118   1.0000   0.7165
  -3.250  -0.4338   0.03814   0.03127   0.0199   1.0000   0.7435
  -3.000  -0.4268   0.03901   0.03205   0.0263   1.0000   0.7748
  -2.750  -0.4176   0.03943   0.03238   0.0318   1.0000   0.8053
  -2.500  -0.4078   0.03948   0.03232   0.0360   1.0000   0.8370
  -2.250  -0.3861   0.03952   0.03223   0.0387   1.0000   0.8678
  -2.000  -0.3532   0.03952   0.03209   0.0385   1.0000   0.9011
  -1.750  -0.1958   0.04065   0.03278   0.0159   0.9919   0.9417
  -1.500  -0.0269   0.04044   0.03219  -0.0114   0.9670   0.9655
  -1.250   0.1235   0.03909   0.03061  -0.0353   0.9388   0.9845
  -1.000   0.2746   0.03674   0.02809  -0.0587   0.9103   1.0000
  -0.750   0.3640   0.03457   0.02583  -0.0700   0.8795   1.0000
  -0.500   0.4262   0.03271   0.02389  -0.0755   0.8437   1.0000
  -0.250   0.4744   0.03106   0.02211  -0.0780   0.8100   1.0000
   0.000   0.4994   0.03032   0.02124  -0.0768   0.7721   1.0000
   0.250   0.5258   0.02952   0.02027  -0.0758   0.7428   1.0000
   0.500   0.5461   0.02920   0.01980  -0.0744   0.7147   1.0000
   0.750   0.5659   0.02901   0.01948  -0.0730   0.6899   1.0000
   1.000   0.5863   0.02885   0.01918  -0.0719   0.6687   1.0000
   1.250   0.6074   0.02874   0.01893  -0.0708   0.6504   1.0000
   1.500   0.6297   0.02862   0.01864  -0.0699   0.6343   1.0000
   1.750   0.6478   0.02890   0.01887  -0.0689   0.6177   1.0000
   2.000   0.6666   0.02921   0.01911  -0.0679   0.6033   1.0000
   2.250   0.6899   0.02929   0.01902  -0.0672   0.5914   1.0000
   2.500   0.7060   0.02988   0.01964  -0.0661   0.5783   1.0000
   2.750   0.7260   0.03029   0.01997  -0.0652   0.5674   1.0000
   3.000   0.7429   0.03090   0.02058  -0.0641   0.5564   1.0000
   3.250   0.7619   0.03149   0.02110  -0.0631   0.5472   1.0000
   3.500   0.7743   0.03239   0.02208  -0.0614   0.5371   1.0000
   3.750   0.7982   0.03282   0.02235  -0.0609   0.5288   1.0000
   4.000   0.8006   0.03426   0.02402  -0.0582   0.5201   1.0000
   4.250   0.8159   0.03506   0.02480  -0.0567   0.5119   1.0000
   4.500   0.8294   0.03608   0.02582  -0.0551   0.5047   1.0000
   4.750   0.8238   0.03786   0.02780  -0.0514   0.4975   1.0000
   5.000   0.8346   0.03888   0.02884  -0.0494   0.4904   1.0000
   5.250   0.8519   0.03981   0.02972  -0.0480   0.4840   1.0000
   5.500   0.8192   0.04258   0.03279  -0.0414   0.4787   1.0000
   5.750   0.7845   0.04542   0.03582  -0.0348   0.4737   1.0000
   6.000   0.7797   0.04725   0.03770  -0.0317   0.4680   1.0000
   6.250   0.8237   0.04738   0.03771  -0.0331   0.4612   1.0000
   6.500   0.6829   0.05979   0.05055  -0.0257   0.4584   1.0000
   6.750   0.5957   0.07395   0.06487  -0.0304   0.4583   1.0000
   7.000   0.5794   0.08041   0.07139  -0.0334   0.4581   1.0000
   7.250   0.5784   0.08564   0.07668  -0.0361   0.4594   1.0000
   7.500   0.5879   0.09027   0.08137  -0.0388   0.4609   1.0000
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