Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il)
Reynolds number: 100,000
Max Cl/Cd: 53.31 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ls413mod-il-100000.txt
Download as CSV file: xf-ls413mod-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4570   0.09290   0.08792  -0.0397   1.0000   0.1580
  -9.250  -0.4197   0.09132   0.08628  -0.0347   1.0000   0.1629
  -9.000  -0.4496   0.08767   0.08278  -0.0383   1.0000   0.1713
  -8.750  -0.4581   0.08379   0.07900  -0.0378   1.0000   0.1752
  -8.500  -0.4358   0.08200   0.07720  -0.0349   1.0000   0.1805
  -8.250  -0.4837   0.07876   0.07418  -0.0355   1.0000   0.1866
  -8.000  -0.5220   0.07447   0.06998  -0.0365   1.0000   0.1910
  -7.750  -0.5010   0.07335   0.06893  -0.0317   1.0000   0.1956
  -7.000  -0.5572   0.04386   0.03722  -0.0471   1.0000   0.0938
  -6.750  -0.5349   0.03911   0.03150  -0.0479   1.0000   0.0823
  -6.500  -0.5146   0.03601   0.02825  -0.0481   1.0000   0.0805
  -6.250  -0.4917   0.03353   0.02547  -0.0483   1.0000   0.0790
  -6.000  -0.4671   0.03142   0.02304  -0.0485   1.0000   0.0779
  -5.750  -0.4413   0.02968   0.02098  -0.0487   1.0000   0.0775
  -5.500  -0.4159   0.02837   0.01943  -0.0488   1.0000   0.0786
  -5.250  -0.3910   0.02738   0.01823  -0.0488   1.0000   0.0808
  -5.000  -0.3662   0.02658   0.01721  -0.0486   1.0000   0.0826
  -4.750  -0.3414   0.02526   0.01593  -0.0485   1.0000   0.0846
  -4.500  -0.3059   0.02429   0.01505  -0.0503   0.9971   0.0887
  -4.250  -0.2627   0.02359   0.01429  -0.0534   0.9928   0.0959
  -4.000  -0.2215   0.02266   0.01354  -0.0566   0.9876   0.1100
  -3.750  -0.1726   0.02111   0.01253  -0.0616   0.9844   0.1647
  -3.500  -0.1316   0.01955   0.01350  -0.0638   0.9821   0.6658
  -3.250  -0.1072   0.02054   0.01447  -0.0615   0.9727   0.7174
  -3.000  -0.0881   0.02154   0.01546  -0.0578   0.9634   0.7519
  -2.750  -0.0656   0.02247   0.01633  -0.0545   0.9557   0.7831
  -2.500  -0.0548   0.02307   0.01690  -0.0493   0.9449   0.8084
  -2.250  -0.0381   0.02347   0.01725  -0.0453   0.9363   0.8333
  -2.000  -0.0201   0.02362   0.01733  -0.0419   0.9273   0.8567
  -1.750  -0.0018   0.02361   0.01725  -0.0393   0.9180   0.8771
  -1.500   0.0246   0.02339   0.01694  -0.0380   0.9104   0.8948
  -1.250   0.0453   0.02321   0.01668  -0.0364   0.9007   0.9105
  -1.000   0.0805   0.02288   0.01626  -0.0373   0.8941   0.9253
  -0.750   0.1072   0.02272   0.01603  -0.0374   0.8845   0.9366
  -0.500   0.1512   0.02244   0.01567  -0.0405   0.8782   0.9451
  -0.250   0.1873   0.02227   0.01544  -0.0423   0.8696   0.9544
   0.000   0.2356   0.02197   0.01508  -0.0462   0.8629   0.9625
   0.250   0.2950   0.02155   0.01460  -0.0518   0.8588   0.9679
   0.500   0.3335   0.02142   0.01446  -0.0543   0.8473   0.9748
   0.750   0.3896   0.02095   0.01396  -0.0592   0.8403   0.9789
   1.000   0.4375   0.02051   0.01351  -0.0627   0.8296   0.9838
   1.250   0.4880   0.02008   0.01308  -0.0667   0.8182   0.9877
   1.500   0.5405   0.01931   0.01229  -0.0704   0.8099   0.9914
   1.750   0.5803   0.01899   0.01200  -0.0726   0.7957   0.9965
   2.000   0.6116   0.01876   0.01179  -0.0731   0.7818   1.0000
   2.250   0.6318   0.01854   0.01157  -0.0715   0.7676   1.0000
   2.500   0.6530   0.01823   0.01124  -0.0698   0.7525   1.0000
   2.750   0.6741   0.01791   0.01092  -0.0680   0.7366   1.0000
   3.000   0.6947   0.01764   0.01062  -0.0663   0.7201   1.0000
   3.250   0.7157   0.01743   0.01041  -0.0647   0.7035   1.0000
   3.500   0.7390   0.01725   0.01023  -0.0636   0.6860   1.0000
   3.750   0.7649   0.01709   0.01004  -0.0630   0.6671   1.0000
   4.000   0.7932   0.01693   0.00984  -0.0628   0.6471   1.0000
   4.250   0.8195   0.01691   0.00983  -0.0624   0.6227   1.0000
   4.500   0.8487   0.01684   0.00969  -0.0624   0.5976   1.0000
   4.750   0.8756   0.01692   0.00972  -0.0623   0.5665   1.0000
   5.000   0.9026   0.01708   0.00979  -0.0621   0.5314   1.0000
   5.250   0.9281   0.01741   0.00998  -0.0618   0.4898   1.0000
   5.500   0.9524   0.01794   0.01028  -0.0614   0.4440   1.0000
   5.750   0.9752   0.01870   0.01075  -0.0609   0.3968   1.0000
   6.000   0.9973   0.01965   0.01139  -0.0605   0.3532   1.0000
   6.250   1.0199   0.02067   0.01216  -0.0602   0.3160   1.0000
   6.500   1.0435   0.02179   0.01302  -0.0602   0.2879   1.0000
   6.750   1.0684   0.02287   0.01398  -0.0603   0.2651   1.0000
   7.000   1.0941   0.02397   0.01494  -0.0605   0.2473   1.0000
   7.250   1.1206   0.02508   0.01598  -0.0609   0.2327   1.0000
   7.500   1.1481   0.02625   0.01713  -0.0614   0.2209   1.0000
   7.750   1.1771   0.02753   0.01825  -0.0621   0.2109   1.0000
   8.000   1.2028   0.02860   0.01947  -0.0622   0.2013   1.0000
   8.250   1.2309   0.02995   0.02078  -0.0628   0.1933   1.0000
   8.500   1.2570   0.03117   0.02215  -0.0630   0.1860   1.0000
   8.750   1.2869   0.03278   0.02366  -0.0639   0.1798   1.0000
   9.000   1.3081   0.03401   0.02520  -0.0633   0.1731   1.0000
   9.250   1.3364   0.03545   0.02643  -0.0641   0.1657   1.0000
   9.500   1.3515   0.03654   0.02793  -0.0627   0.1589   1.0000
   9.750   1.3766   0.03774   0.02900  -0.0629   0.1519   1.0000
  10.000   1.3915   0.03931   0.03090  -0.0616   0.1460   1.0000
  10.250   1.4099   0.04050   0.03220  -0.0608   0.1397   1.0000
  10.500   1.4278   0.04231   0.03411  -0.0602   0.1338   1.0000
  10.750   1.4383   0.04381   0.03590  -0.0583   0.1278   1.0000
  11.000   1.4615   0.04559   0.03752  -0.0585   0.1215   1.0000
  11.250   1.4600   0.04749   0.03992  -0.0552   0.1165   1.0000
  11.500   1.4718   0.04892   0.04142  -0.0537   0.1110   1.0000
  11.750   1.4787   0.05139   0.04403  -0.0520   0.1062   1.0000
  12.000   1.4686   0.05371   0.04673  -0.0481   0.1026   1.0000
  12.250   1.4704   0.05548   0.04861  -0.0457   0.0987   1.0000
  12.500   1.4866   0.05809   0.05113  -0.0454   0.0946   1.0000
  12.750   1.4628   0.06133   0.05480  -0.0413   0.0934   1.0000
  13.000   1.4383   0.06524   0.05910  -0.0383   0.0922   1.0000
  13.250   1.4126   0.06978   0.06398  -0.0365   0.0912   1.0000
  13.500   1.3837   0.07511   0.06964  -0.0357   0.0905   1.0000
  13.750   1.3486   0.08171   0.07655  -0.0363   0.0903   1.0000
  14.000   1.3037   0.09040   0.08556  -0.0389   0.0909   1.0000
  14.250   1.2481   0.10206   0.09749  -0.0445   0.0923   1.0000
  14.500   1.1849   0.11751   0.11312  -0.0537   0.0940   1.0000
<< Back to NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il)

Polar data table (+)

Polar graphs


<< Back to NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il)