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NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il)
Reynolds number: 50,000
Max Cl/Cd: 34.87 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ls413mod-il-50000.txt
Download as CSV file: xf-ls413mod-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY LS(1)-0413MOD 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3632   0.10545   0.09819  -0.0158   1.0000   0.3440
  -9.000  -0.3621   0.10275   0.09556  -0.0154   1.0000   0.3568
  -8.750  -0.3664   0.10058   0.09348  -0.0148   1.0000   0.3707
  -8.500  -0.3743   0.09899   0.09200  -0.0138   1.0000   0.3858
  -8.250  -0.3453   0.09406   0.08703  -0.0136   1.0000   0.3918
  -8.000  -0.3507   0.09182   0.08490  -0.0125   1.0000   0.4047
  -7.250  -0.5517   0.05799   0.05062  -0.0455   1.0000   0.1546
  -7.000  -0.5429   0.05408   0.04656  -0.0454   1.0000   0.1500
  -6.750  -0.5306   0.04880   0.04031  -0.0481   1.0000   0.1401
  -6.500  -0.5135   0.04554   0.03686  -0.0481   1.0000   0.1379
  -6.250  -0.4934   0.04239   0.03335  -0.0486   1.0000   0.1354
  -6.000  -0.4707   0.03956   0.03010  -0.0490   1.0000   0.1337
  -5.750  -0.4465   0.03714   0.02728  -0.0493   1.0000   0.1334
  -5.500  -0.4220   0.03521   0.02501  -0.0494   1.0000   0.1357
  -5.250  -0.3964   0.03361   0.02300  -0.0494   1.0000   0.1396
  -5.000  -0.3723   0.03197   0.02125  -0.0491   1.0000   0.1441
  -4.750  -0.3484   0.03070   0.01995  -0.0484   1.0000   0.1496
  -4.500  -0.3244   0.02963   0.01873  -0.0474   1.0000   0.1580
  -4.250  -0.3020   0.02870   0.01785  -0.0462   1.0000   0.1724
  -4.000  -0.2803   0.02760   0.01700  -0.0449   1.0000   0.1928
  -3.750  -0.2530   0.02581   0.01584  -0.0454   1.0000   0.2578
  -3.500  -0.2556   0.02499   0.01779  -0.0351   1.0000   0.6902
  -3.250  -0.2655   0.02642   0.01922  -0.0242   1.0000   0.7564
  -3.000  -0.2745   0.02732   0.02007  -0.0141   1.0000   0.8056
  -2.750  -0.2808   0.02764   0.02031  -0.0052   1.0000   0.8503
  -2.500  -0.2830   0.02745   0.01999   0.0024   1.0000   0.8941
  -2.250  -0.0971   0.02805   0.01967  -0.0211   1.0000   1.0000
  -2.000  -0.1023   0.02784   0.01939  -0.0182   1.0000   1.0000
  -1.750  -0.1067   0.02764   0.01912  -0.0153   1.0000   1.0000
  -1.500  -0.1101   0.02745   0.01887  -0.0125   1.0000   1.0000
  -1.250  -0.1122   0.02728   0.01862  -0.0100   1.0000   1.0000
  -1.000  -0.1124   0.02714   0.01842  -0.0077   1.0000   1.0000
  -0.750  -0.1065   0.02715   0.01835  -0.0065   1.0000   1.0000
  -0.500  -0.0937   0.02735   0.01846  -0.0065   1.0000   1.0000
  -0.250  -0.0758   0.02774   0.01875  -0.0074   1.0000   1.0000
   0.000  -0.0344   0.02872   0.01957  -0.0125   0.9926   1.0000
   0.250   0.0159   0.02988   0.02058  -0.0189   0.9795   1.0000
   0.500   0.0635   0.03098   0.02157  -0.0247   0.9662   1.0000
   0.750   0.1105   0.03206   0.02256  -0.0303   0.9522   1.0000
   1.000   0.1570   0.03312   0.02354  -0.0355   0.9377   1.0000
   1.250   0.2028   0.03412   0.02450  -0.0405   0.9227   1.0000
   1.500   0.2471   0.03506   0.02540  -0.0449   0.9068   1.0000
   1.750   0.2893   0.03591   0.02625  -0.0488   0.8900   1.0000
   2.000   0.3294   0.03670   0.02705  -0.0521   0.8722   1.0000
   2.250   0.3688   0.03743   0.02780  -0.0551   0.8539   1.0000
   2.500   0.4087   0.03809   0.02850  -0.0579   0.8353   1.0000
   2.750   0.4505   0.03861   0.02908  -0.0606   0.8163   1.0000
   3.000   0.4950   0.03890   0.02943  -0.0632   0.7972   1.0000
   3.250   0.5429   0.03888   0.02950  -0.0658   0.7785   1.0000
   3.500   0.5939   0.03846   0.02919  -0.0681   0.7605   1.0000
   3.750   0.6201   0.03883   0.02966  -0.0678   0.7368   1.0000
   4.000   0.6718   0.03762   0.02858  -0.0689   0.7164   1.0000
   4.250   0.7273   0.03559   0.02668  -0.0693   0.6969   1.0000
   4.500   0.7660   0.03445   0.02564  -0.0685   0.6727   1.0000
   4.750   0.8141   0.03249   0.02379  -0.0680   0.6494   1.0000
   5.000   0.8532   0.03113   0.02250  -0.0670   0.6214   1.0000
   5.250   0.8930   0.02972   0.02111  -0.0659   0.5898   1.0000
   5.500   0.9294   0.02865   0.01994  -0.0646   0.5529   1.0000
   5.750   0.9605   0.02825   0.01935  -0.0633   0.5118   1.0000
   6.000   0.9894   0.02837   0.01920  -0.0621   0.4696   1.0000
   6.250   1.0141   0.02916   0.01977  -0.0611   0.4300   1.0000
   6.500   1.0393   0.03022   0.02062  -0.0604   0.3960   1.0000
   6.750   1.0659   0.03143   0.02160  -0.0601   0.3680   1.0000
   7.000   1.0902   0.03285   0.02302  -0.0598   0.3444   1.0000
   7.250   1.1174   0.03425   0.02428  -0.0598   0.3247   1.0000
   7.500   1.1435   0.03590   0.02592  -0.0598   0.3085   1.0000
   7.750   1.1667   0.03773   0.02789  -0.0596   0.2944   1.0000
   8.000   1.1884   0.03970   0.03003  -0.0592   0.2819   1.0000
   8.250   1.2127   0.04179   0.03218  -0.0592   0.2716   1.0000
   8.500   1.2328   0.04409   0.03471  -0.0587   0.2627   1.0000
   8.750   1.2493   0.04672   0.03761  -0.0580   0.2548   1.0000
   9.000   1.2676   0.04905   0.04011  -0.0574   0.2463   1.0000
   9.250   1.2792   0.05219   0.04354  -0.0563   0.2405   1.0000
   9.500   1.2760   0.05600   0.04784  -0.0541   0.2359   1.0000
  10.000   1.2870   0.06268   0.05496  -0.0514   0.2245   1.0000
  10.250   1.2597   0.06788   0.06058  -0.0484   0.2226   1.0000
  10.500   1.2184   0.07397   0.06696  -0.0454   0.2230   1.0000
  10.750   1.1660   0.08206   0.07524  -0.0446   0.2251   1.0000
  11.000   1.1177   0.09213   0.08540  -0.0471   0.2270   1.0000
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