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)-0013 AIRFOIL (ls013-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il)
Reynolds number: 200,000
Max Cl/Cd: 39.84 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ls013-il-200000.txt
Download as CSV file: xf-ls013-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY LS(1)-0013 AIRFOIL                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.8467   0.10241   0.09874  -0.0047   1.0000   0.0515
 -13.000  -0.8492   0.09698   0.09334  -0.0056   1.0000   0.0509
 -12.750  -0.8763   0.08764   0.08392  -0.0112   1.0000   0.0500
 -12.500  -0.9122   0.07844   0.07458  -0.0167   1.0000   0.0491
 -12.250  -0.9497   0.07054   0.06646  -0.0201   1.0000   0.0482
 -12.000  -0.9864   0.06381   0.05946  -0.0213   1.0000   0.0475
 -11.750  -1.0195   0.05823   0.05357  -0.0204   1.0000   0.0469
 -11.500  -1.0445   0.05387   0.04890  -0.0178   1.0000   0.0467
 -11.250  -1.0566   0.04992   0.04461  -0.0155   1.0000   0.0468
 -11.000  -1.0595   0.04635   0.04069  -0.0137   1.0000   0.0472
 -10.750  -1.0557   0.04317   0.03718  -0.0121   1.0000   0.0478
 -10.500  -1.0470   0.04046   0.03414  -0.0108   1.0000   0.0489
 -10.250  -1.0354   0.03803   0.03132  -0.0096   1.0000   0.0502
 -10.000  -1.0211   0.03611   0.02897  -0.0084   1.0000   0.0513
  -9.750  -1.0026   0.03291   0.02554  -0.0081   1.0000   0.0526
  -9.500  -0.9795   0.03119   0.02379  -0.0082   1.0000   0.0541
  -9.250  -0.9561   0.02997   0.02249  -0.0081   1.0000   0.0561
  -9.000  -0.9322   0.02871   0.02106  -0.0081   1.0000   0.0584
  -8.750  -0.9078   0.02757   0.01967  -0.0080   1.0000   0.0604
  -8.500  -0.8818   0.02550   0.01754  -0.0083   1.0000   0.0627
  -8.250  -0.8558   0.02446   0.01653  -0.0087   1.0000   0.0655
  -8.000  -0.8297   0.02360   0.01558  -0.0089   1.0000   0.0689
  -7.750  -0.8038   0.02291   0.01473  -0.0091   1.0000   0.0716
  -7.500  -0.7799   0.02139   0.01331  -0.0091   1.0000   0.0753
  -7.250  -0.7604   0.02078   0.01271  -0.0083   1.0000   0.0790
  -7.000  -0.7461   0.02036   0.01220  -0.0064   0.9993   0.0825
  -6.750  -0.7119   0.01918   0.01108  -0.0085   0.9942   0.0878
  -6.500  -0.6770   0.01850   0.01039  -0.0105   0.9885   0.0945
  -6.250  -0.6430   0.01754   0.00949  -0.0125   0.9833   0.1021
  -6.000  -0.6096   0.01689   0.00882  -0.0141   0.9771   0.1118
  -5.750  -0.5771   0.01614   0.00813  -0.0157   0.9713   0.1254
  -5.500  -0.5487   0.01526   0.00744  -0.0167   0.9648   0.1468
  -5.250  -0.5222   0.01432   0.00680  -0.0173   0.9583   0.1924
  -5.000  -0.5007   0.01313   0.00625  -0.0173   0.9512   0.3047
  -4.750  -0.4803   0.01208   0.00584  -0.0167   0.9441   0.4315
  -4.500  -0.4594   0.01153   0.00596  -0.0155   0.9378   0.5651
  -4.250  -0.4342   0.01161   0.00620  -0.0146   0.9317   0.6279
  -4.000  -0.4084   0.01183   0.00642  -0.0138   0.9266   0.6632
  -3.750  -0.3818   0.01206   0.00665  -0.0132   0.9207   0.6883
  -3.500  -0.3559   0.01235   0.00692  -0.0123   0.9151   0.7085
  -3.250  -0.3308   0.01269   0.00722  -0.0111   0.9109   0.7255
  -3.000  -0.3037   0.01298   0.00750  -0.0107   0.9055   0.7399
  -2.750  -0.2779   0.01326   0.00776  -0.0098   0.9003   0.7528
  -2.500  -0.2532   0.01355   0.00800  -0.0086   0.8961   0.7660
  -2.250  -0.2277   0.01395   0.00843  -0.0074   0.8915   0.7777
  -2.000  -0.2034   0.01435   0.00885  -0.0059   0.8864   0.7896
  -1.750  -0.1791   0.01459   0.00906  -0.0046   0.8820   0.8008
  -1.500  -0.1543   0.01479   0.00925  -0.0035   0.8782   0.8075
  -1.250  -0.1258   0.01484   0.00927  -0.0040   0.8731   0.8156
  -1.000  -0.1009   0.01498   0.00941  -0.0030   0.8684   0.8209
  -0.750  -0.0751   0.01504   0.00944  -0.0024   0.8647   0.8277
  -0.500  -0.0484   0.01511   0.00952  -0.0023   0.8601   0.8338
  -0.250  -0.0235   0.01524   0.00967  -0.0013   0.8544   0.8407
   0.000   0.0000   0.01523   0.00964   0.0000   0.8489   0.8489
   0.250   0.0235   0.01524   0.00967   0.0013   0.8407   0.8544
   0.500   0.0483   0.01511   0.00952   0.0023   0.8338   0.8601
   0.750   0.0751   0.01504   0.00944   0.0024   0.8277   0.8647
   1.000   0.1009   0.01497   0.00941   0.0030   0.8209   0.8684
   1.250   0.1258   0.01484   0.00927   0.0040   0.8156   0.8731
   1.500   0.1543   0.01479   0.00925   0.0035   0.8076   0.8782
   1.750   0.1791   0.01459   0.00906   0.0046   0.8008   0.8820
   2.000   0.2034   0.01435   0.00885   0.0059   0.7896   0.8864
   2.250   0.2277   0.01395   0.00843   0.0074   0.7777   0.8915
   2.500   0.2531   0.01356   0.00800   0.0086   0.7661   0.8961
   2.750   0.2779   0.01326   0.00775   0.0098   0.7528   0.9003
   3.000   0.3037   0.01298   0.00750   0.0107   0.7399   0.9055
   3.250   0.3308   0.01269   0.00722   0.0112   0.7255   0.9109
   3.500   0.3558   0.01235   0.00692   0.0123   0.7086   0.9151
   3.750   0.3817   0.01206   0.00665   0.0132   0.6883   0.9207
   4.000   0.4084   0.01183   0.00642   0.0138   0.6632   0.9266
   4.250   0.4341   0.01161   0.00620   0.0146   0.6279   0.9317
   4.500   0.4594   0.01153   0.00596   0.0155   0.5651   0.9378
   4.750   0.4802   0.01208   0.00584   0.0167   0.4315   0.9441
   5.000   0.5007   0.01313   0.00625   0.0173   0.3048   0.9512
   5.250   0.5222   0.01432   0.00680   0.0173   0.1924   0.9583
   5.500   0.5487   0.01526   0.00744   0.0167   0.1468   0.9648
   5.750   0.5771   0.01614   0.00813   0.0157   0.1253   0.9713
   6.000   0.6096   0.01689   0.00882   0.0141   0.1118   0.9772
   6.250   0.6430   0.01754   0.00949   0.0125   0.1021   0.9834
   6.500   0.6771   0.01850   0.01039   0.0105   0.0945   0.9885
   6.750   0.7119   0.01918   0.01108   0.0084   0.0878   0.9943
   7.000   0.7462   0.02036   0.01220   0.0064   0.0824   0.9994
   7.250   0.7603   0.02078   0.01271   0.0083   0.0790   1.0000
   7.500   0.7799   0.02139   0.01331   0.0091   0.0753   1.0000
   7.750   0.8038   0.02291   0.01473   0.0091   0.0716   1.0000
   8.000   0.8298   0.02360   0.01558   0.0089   0.0688   1.0000
   8.250   0.8559   0.02446   0.01653   0.0086   0.0655   1.0000
   8.500   0.8819   0.02550   0.01754   0.0083   0.0627   1.0000
   8.750   0.9079   0.02757   0.01967   0.0080   0.0604   1.0000
   9.000   0.9323   0.02872   0.02107   0.0081   0.0584   1.0000
   9.250   0.9562   0.02997   0.02250   0.0081   0.0561   1.0000
   9.500   0.9796   0.03120   0.02380   0.0081   0.0541   1.0000
   9.750   1.0027   0.03291   0.02554   0.0081   0.0526   1.0000
  10.000   1.0212   0.03613   0.02899   0.0084   0.0513   1.0000
  10.250   1.0356   0.03805   0.03134   0.0095   0.0502   1.0000
  10.500   1.0471   0.04048   0.03416   0.0108   0.0489   1.0000
  10.750   1.0559   0.04319   0.03720   0.0121   0.0478   1.0000
  11.000   1.0597   0.04638   0.04073   0.0136   0.0472   1.0000
  11.250   1.0568   0.04995   0.04464   0.0155   0.0468   1.0000
  11.500   1.0446   0.05392   0.04896   0.0177   0.0467   1.0000
  11.750   1.0197   0.05828   0.05363   0.0203   0.0470   1.0000
  12.000   0.9869   0.06385   0.05951   0.0211   0.0475   1.0000
  12.250   0.9501   0.07063   0.06655   0.0199   0.0482   1.0000
  12.500   0.9127   0.07854   0.07468   0.0165   0.0491   1.0000
  12.750   0.8768   0.08778   0.08407   0.0110   0.0500   1.0000
  13.000   0.8503   0.09705   0.09341   0.0054   0.0509   1.0000
  13.250   0.8480   0.10245   0.09878   0.0045   0.0515   1.0000
<< Back to NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il)

Polar data table (+)

Polar graphs


<< Back to NASA/LANGLEY LS(1)-0013 AIRFOIL (ls013-il)