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=500,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY LS(1)-0413MOD AIRFOIL (ls413mod-il)
Reynolds number: 500,000
Max Cl/Cd: 86.73 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ls413mod-il-500000.txt
Download as CSV file: xf-ls413mod-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.4972   0.08271   0.08031  -0.0452   1.0000   0.0368
 -10.500  -0.5228   0.07033   0.06794  -0.0561   1.0000   0.0372
 -10.250  -0.5650   0.06134   0.05883  -0.0626   1.0000   0.0372
 -10.000  -0.6033   0.05786   0.05530  -0.0604   1.0000   0.0371
  -9.750  -0.6290   0.05538   0.05276  -0.0574   1.0000   0.0373
  -9.500  -0.6271   0.05171   0.04893  -0.0593   0.9976   0.0379
  -9.250  -0.6092   0.04792   0.04433  -0.0660   0.9921   0.0415
  -9.000  -0.5966   0.03987   0.03612  -0.0698   0.9882   0.0428
  -8.750  -0.5785   0.02712   0.02206  -0.0730   0.9848   0.0288
  -8.500  -0.5465   0.02388   0.01854  -0.0754   0.9832   0.0280
  -8.250  -0.5144   0.02168   0.01609  -0.0771   0.9803   0.0276
  -8.000  -0.4817   0.02001   0.01423  -0.0787   0.9769   0.0275
  -7.750  -0.4469   0.01865   0.01272  -0.0805   0.9744   0.0276
  -7.500  -0.4120   0.01749   0.01145  -0.0823   0.9719   0.0278
  -7.250  -0.3770   0.01649   0.01037  -0.0840   0.9693   0.0280
  -7.000  -0.3474   0.01568   0.00951  -0.0845   0.9623   0.0283
  -6.750  -0.3150   0.01493   0.00870  -0.0856   0.9574   0.0287
  -6.500  -0.2848   0.01432   0.00804  -0.0862   0.9507   0.0291
  -6.250  -0.2559   0.01343   0.00712  -0.0867   0.9433   0.0299
  -6.000  -0.2269   0.01272   0.00641  -0.0872   0.9355   0.0309
  -5.750  -0.1980   0.01223   0.00591  -0.0875   0.9273   0.0319
  -5.500  -0.1690   0.01180   0.00544  -0.0879   0.9188   0.0330
  -5.250  -0.1399   0.01142   0.00501  -0.0882   0.9105   0.0343
  -5.000  -0.1104   0.01098   0.00453  -0.0887   0.9017   0.0360
  -4.750  -0.0809   0.01065   0.00418  -0.0890   0.8935   0.0387
  -4.500  -0.0512   0.01038   0.00388  -0.0895   0.8845   0.0427
  -4.250  -0.0212   0.01007   0.00359  -0.0899   0.8762   0.0531
  -4.000   0.0107   0.00940   0.00324  -0.0913   0.8670   0.1230
  -3.750   0.0436   0.00864   0.00294  -0.0931   0.8589   0.2529
  -3.500   0.0786   0.00783   0.00271  -0.0955   0.8497   0.4099
  -3.250   0.1127   0.00727   0.00263  -0.0973   0.8422   0.5554
  -3.000   0.1433   0.00719   0.00269  -0.0978   0.8334   0.6130
  -2.750   0.1730   0.00724   0.00272  -0.0979   0.8260   0.6431
  -2.500   0.2030   0.00729   0.00275  -0.0982   0.8174   0.6637
  -2.000   0.2617   0.00745   0.00287  -0.0983   0.8010   0.6957
  -1.500   0.3185   0.00765   0.00306  -0.0979   0.7838   0.7218
  -1.250   0.3473   0.00775   0.00311  -0.0979   0.7738   0.7322
  -1.000   0.3753   0.00782   0.00313  -0.0976   0.7618   0.7383
  -0.750   0.4046   0.00790   0.00316  -0.0977   0.7497   0.7459
  -0.500   0.4327   0.00796   0.00321  -0.0975   0.7395   0.7514
  -0.250   0.4616   0.00802   0.00323  -0.0976   0.7295   0.7555
   0.000   0.4911   0.00806   0.00323  -0.0978   0.7180   0.7594
   0.250   0.5209   0.00812   0.00322  -0.0982   0.7045   0.7636
   0.500   0.5487   0.00817   0.00324  -0.0980   0.6898   0.7672
   0.750   0.5768   0.00826   0.00330  -0.0978   0.6757   0.7715
   1.000   0.6055   0.00837   0.00335  -0.0979   0.6620   0.7762
   1.250   0.6351   0.00846   0.00338  -0.0982   0.6469   0.7804
   1.500   0.6630   0.00853   0.00344  -0.0981   0.6316   0.7835
   1.750   0.6908   0.00864   0.00351  -0.0980   0.6146   0.7871
   2.000   0.7190   0.00877   0.00359  -0.0980   0.5953   0.7908
   2.250   0.7478   0.00891   0.00366  -0.0982   0.5738   0.7943
   2.500   0.7760   0.00910   0.00374  -0.0983   0.5489   0.7974
   2.750   0.8028   0.00929   0.00385  -0.0981   0.5199   0.8000
   3.000   0.8291   0.00956   0.00400  -0.0978   0.4878   0.8031
   3.250   0.8555   0.00989   0.00418  -0.0976   0.4536   0.8066
   3.500   0.8822   0.01023   0.00437  -0.0975   0.4199   0.8102
   3.750   0.9088   0.01060   0.00459  -0.0974   0.3864   0.8137
   4.000   0.9339   0.01096   0.00483  -0.0970   0.3509   0.8167
   4.250   0.9587   0.01140   0.00511  -0.0965   0.3134   0.8202
   4.500   0.9837   0.01187   0.00542  -0.0962   0.2799   0.8240
   4.750   1.0094   0.01233   0.00573  -0.0960   0.2515   0.8278
   5.000   1.0342   0.01277   0.00605  -0.0957   0.2263   0.8310
   5.250   1.0586   0.01317   0.00638  -0.0951   0.2046   0.8343
   5.500   1.0830   0.01360   0.00673  -0.0947   0.1855   0.8382
   5.750   1.1081   0.01404   0.00709  -0.0944   0.1694   0.8424
   6.000   1.1331   0.01444   0.00745  -0.0940   0.1572   0.8460
   6.250   1.1567   0.01484   0.00783  -0.0934   0.1478   0.8495
   6.500   1.1804   0.01526   0.00823  -0.0927   0.1397   0.8538
   6.750   1.2048   0.01567   0.00865  -0.0923   0.1330   0.8585
   7.000   1.2284   0.01608   0.00905  -0.0917   0.1259   0.8625
   7.250   1.2510   0.01648   0.00948  -0.0909   0.1196   0.8666
   7.500   1.2748   0.01684   0.00985  -0.0903   0.1142   0.8713
   7.750   1.2963   0.01743   0.01042  -0.0894   0.1087   0.8762
   8.000   1.3189   0.01772   0.01079  -0.0886   0.1055   0.8807
   8.250   1.3414   0.01807   0.01119  -0.0878   0.1015   0.8861
   8.500   1.3625   0.01859   0.01170  -0.0868   0.0974   0.8914
   8.750   1.3813   0.01910   0.01227  -0.0854   0.0942   0.8964
   9.000   1.4032   0.01943   0.01269  -0.0845   0.0916   0.9023
   9.250   1.4234   0.01982   0.01313  -0.0834   0.0884   0.9082
   9.500   1.4406   0.02035   0.01367  -0.0818   0.0851   0.9150
   9.750   1.4567   0.02093   0.01431  -0.0800   0.0822   0.9221
  10.000   1.4747   0.02125   0.01473  -0.0784   0.0798   0.9299
  10.250   1.4886   0.02163   0.01519  -0.0762   0.0771   0.9390
  10.500   1.5001   0.02216   0.01576  -0.0736   0.0741   0.9514
  10.750   1.5079   0.02278   0.01644  -0.0706   0.0711   1.0000
  11.000   1.5297   0.02328   0.01703  -0.0702   0.0678   1.0000
  11.250   1.5466   0.02409   0.01782  -0.0693   0.0634   1.0000
  11.500   1.5630   0.02492   0.01871  -0.0683   0.0588   1.0000
  11.750   1.5767   0.02595   0.01971  -0.0670   0.0532   1.0000
  12.000   1.5904   0.02698   0.02080  -0.0658   0.0480   1.0000
  12.250   1.5995   0.02838   0.02218  -0.0642   0.0436   1.0000
  12.500   1.6105   0.02964   0.02350  -0.0629   0.0403   1.0000
  12.750   1.6171   0.03130   0.02516  -0.0614   0.0378   1.0000
  13.000   1.6243   0.03295   0.02689  -0.0600   0.0360   1.0000
  13.250   1.6314   0.03465   0.02867  -0.0588   0.0345   1.0000
  13.500   1.6359   0.03664   0.03072  -0.0576   0.0332   1.0000
  13.750   1.6364   0.03906   0.03320  -0.0564   0.0320   1.0000
  14.000   1.6349   0.04178   0.03601  -0.0553   0.0311   1.0000
  14.250   1.6390   0.04400   0.03835  -0.0546   0.0304   1.0000
  14.500   1.6406   0.04654   0.04100  -0.0539   0.0298   1.0000
  14.750   1.6408   0.04930   0.04388  -0.0535   0.0290   1.0000
  15.000   1.6393   0.05234   0.04702  -0.0532   0.0283   1.0000
  15.250   1.6349   0.05581   0.05058  -0.0531   0.0278   1.0000
  15.500   1.6282   0.05970   0.05457  -0.0533   0.0272   1.0000
  15.750   1.6187   0.06413   0.05910  -0.0538   0.0268   1.0000
  16.000   1.6064   0.06905   0.06415  -0.0546   0.0264   1.0000
  16.250   1.5957   0.07393   0.06914  -0.0556   0.0260   1.0000
  16.500   1.5905   0.07823   0.07358  -0.0567   0.0257   1.0000
  16.750   1.5836   0.08289   0.07839  -0.0581   0.0254   1.0000
  17.000   1.5748   0.08789   0.08352  -0.0597   0.0251   1.0000
  17.250   1.5648   0.09319   0.08896  -0.0615   0.0248   1.0000
  17.500   1.5538   0.09874   0.09463  -0.0636   0.0245   1.0000
  17.750   1.5423   0.10448   0.10050  -0.0659   0.0242   1.0000
  18.000   1.5301   0.11037   0.10652  -0.0684   0.0239   1.0000
  18.250   1.5180   0.11635   0.11260  -0.0711   0.0236   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)