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 NLF(1)-0215F AIRFOIL (nlf0215f-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)
Reynolds number: 50,000
Max Cl/Cd: 20.82 at α=11.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-nlf0215f-il-50000-n5.txt
Download as CSV file: xf-nlf0215f-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY NLF(1)-0215F 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3662   0.10804   0.10180  -0.0505   0.9988   0.0619
  -9.000  -0.3413   0.10425   0.09794  -0.0525   0.9932   0.0587
  -8.750  -0.3292   0.09899   0.09268  -0.0578   0.9853   0.0569
  -8.500  -0.3223   0.09331   0.08700  -0.0638   0.9753   0.0551
  -8.250  -0.3555   0.08297   0.07653  -0.0762   0.9557   0.0500
  -8.000  -0.3527   0.07870   0.07219  -0.0796   0.9434   0.0496
  -7.750  -0.3520   0.07383   0.06717  -0.0839   0.9307   0.0495
  -7.500  -0.3446   0.06911   0.06224  -0.0880   0.9212   0.0492
  -7.250  -0.3421   0.06508   0.05798  -0.0901   0.9093   0.0492
  -7.000  -0.3355   0.06131   0.05391  -0.0919   0.8989   0.0492
  -6.750  -0.3197   0.05692   0.04904  -0.0951   0.8919   0.0495
  -6.500  -0.3084   0.05387   0.04565  -0.0956   0.8820   0.0494
  -6.250  -0.2837   0.05022   0.04154  -0.0980   0.8768   0.0494
  -6.000  -0.2666   0.04759   0.03854  -0.0983   0.8684   0.0493
  -5.750  -0.2392   0.04468   0.03517  -0.0999   0.8630   0.0494
  -5.250  -0.1869   0.04025   0.02987  -0.1011   0.8508   0.0500
  -5.000  -0.1553   0.03815   0.02752  -0.1023   0.8466   0.0511
  -4.750  -0.1200   0.03651   0.02573  -0.1039   0.8436   0.0541
  -4.500  -0.1048   0.03582   0.02489  -0.1020   0.8351   0.0567
  -4.250  -0.0739   0.03467   0.02349  -0.1019   0.8309   0.0596
  -4.000  -0.0404   0.03358   0.02223  -0.1016   0.8279   0.0621
  -3.750  -0.0288   0.03324   0.02188  -0.0987   0.8199   0.0645
  -3.500  -0.0030   0.03261   0.02115  -0.0980   0.8153   0.0700
  -3.250   0.0274   0.03177   0.02028  -0.0985   0.8121   0.0793
  -2.750   0.0626   0.03077   0.01937  -0.0962   0.7996   0.1111
  -2.500   0.0862   0.02805   0.01875  -0.0978   0.7964   0.4067
  -2.250   0.0571   0.02995   0.02155  -0.0823   0.7878   0.5702
  -2.000   0.0771   0.03157   0.02280  -0.0797   0.7828   0.7444
  -1.750   0.0728   0.03242   0.02362  -0.0689   0.7791   0.8050
  -1.500   0.0547   0.03304   0.02426  -0.0595   0.7703   0.8360
  -1.250   0.0597   0.03293   0.02401  -0.0525   0.7662   0.8789
  -1.000   0.0900   0.03269   0.02346  -0.0524   0.7635   0.8936
  -0.750   0.0900   0.03312   0.02377  -0.0489   0.7559   0.9011
  -0.500   0.1096   0.03320   0.02363  -0.0483   0.7511   0.9065
  -0.250   0.1416   0.03310   0.02328  -0.0495   0.7482   0.9101
   0.000   0.1507   0.03355   0.02360  -0.0477   0.7421   0.9149
   0.250   0.1636   0.03394   0.02385  -0.0464   0.7361   0.9192
   0.500   0.1939   0.03399   0.02371  -0.0475   0.7328   0.9216
   0.750   0.2287   0.03395   0.02349  -0.0491   0.7304   0.9236
   1.000   0.2250   0.03505   0.02457  -0.0461   0.7215   0.9280
   1.250   0.2518   0.03529   0.02468  -0.0468   0.7175   0.9303
   1.500   0.2849   0.03536   0.02461  -0.0483   0.7148   0.9320
   1.750   0.2893   0.03649   0.02572  -0.0467   0.7069   0.9352
   2.000   0.3149   0.03691   0.02606  -0.0473   0.7022   0.9374
   2.250   0.3480   0.03709   0.02615  -0.0488   0.6992   0.9390
   2.500   0.3580   0.03820   0.02724  -0.0480   0.6919   0.9416
   2.750   0.3814   0.03881   0.02781  -0.0486   0.6866   0.9439
   3.000   0.4139   0.03908   0.02802  -0.0499   0.6833   0.9457
   3.250   0.4268   0.04022   0.02918  -0.0496   0.6759   0.9486
   3.500   0.4511   0.04090   0.02985  -0.0503   0.6702   0.9505
   3.750   0.4856   0.04114   0.03007  -0.0519   0.6668   0.9519
   4.000   0.4953   0.04255   0.03152  -0.0514   0.6582   0.9555
   4.250   0.5239   0.04307   0.03206  -0.0525   0.6530   0.9583
   4.500   0.5613   0.04318   0.03219  -0.0542   0.6498   0.9606
   4.750   0.5674   0.04489   0.03397  -0.0537   0.6392   0.9658
   5.000   0.6025   0.04512   0.03423  -0.0553   0.6350   0.9691
   5.250   0.6134   0.04659   0.03577  -0.0551   0.6252   0.9777
   5.500   0.6442   0.04685   0.03611  -0.0561   0.6197   1.0000
   6.000   0.6881   0.04867   0.03806  -0.0571   0.6042   1.0000
   6.250   0.7291   0.04841   0.03790  -0.0587   0.6008   1.0000
   6.500   0.7341   0.05032   0.03990  -0.0581   0.5882   1.0000
   7.000   0.7829   0.05168   0.04145  -0.0591   0.5716   1.0000
   7.250   0.7955   0.05312   0.04302  -0.0589   0.5599   1.0000
   7.500   0.8345   0.05265   0.04267  -0.0600   0.5546   1.0000
   7.750   0.8446   0.05424   0.04438  -0.0597   0.5416   1.0000
   8.250   0.8990   0.05460   0.04504  -0.0602   0.5235   1.0000
   8.500   0.9118   0.05587   0.04644  -0.0600   0.5100   1.0000
   8.750   0.9274   0.05687   0.04759  -0.0598   0.4968   1.0000
   9.250   0.9867   0.05593   0.04701  -0.0597   0.4771   1.0000
   9.750   1.0204   0.05723   0.04863  -0.0589   0.4486   1.0000
  10.000   1.0399   0.05756   0.04914  -0.0585   0.4342   1.0000
  10.250   1.0613   0.05763   0.04937  -0.0581   0.4193   1.0000
  10.500   1.0847   0.05745   0.04935  -0.0576   0.4037   1.0000
  10.750   1.1107   0.05693   0.04899  -0.0571   0.3871   1.0000
  11.000   1.1370   0.05638   0.04852  -0.0565   0.3684   1.0000
  11.250   1.1533   0.05710   0.04931  -0.0559   0.3464   1.0000
  11.500   1.1792   0.05663   0.04878  -0.0551   0.3242   1.0000
  11.750   1.1905   0.05808   0.05023  -0.0545   0.3013   1.0000
  12.000   1.2031   0.05934   0.05141  -0.0539   0.2796   1.0000
  12.250   1.2115   0.06122   0.05321  -0.0534   0.2597   1.0000
  12.500   1.2174   0.06354   0.05552  -0.0531   0.2413   1.0000
  12.750   1.2230   0.06594   0.05790  -0.0529   0.2244   1.0000
  13.000   1.2287   0.06841   0.06038  -0.0527   0.2093   1.0000
  13.250   1.2346   0.07093   0.06290  -0.0527   0.1956   1.0000
  13.500   1.2405   0.07350   0.06547  -0.0527   0.1831   1.0000
  13.750   1.2482   0.07584   0.06776  -0.0527   0.1719   1.0000
  14.000   1.2525   0.07885   0.07093  -0.0530   0.1613   1.0000
  14.250   1.2578   0.08170   0.07387  -0.0532   0.1518   1.0000
  14.750   1.2697   0.08741   0.07981  -0.0539   0.1354   1.0000
  15.000   1.2791   0.08969   0.08206  -0.0541   0.1283   1.0000
  15.250   1.2792   0.09359   0.08625  -0.0550   0.1219   1.0000
  15.500   1.2926   0.09530   0.08787  -0.0550   0.1156   1.0000
  15.750   1.2861   0.10032   0.09328  -0.0565   0.1108   1.0000
  16.000   1.2878   0.10396   0.09707  -0.0575   0.1059   1.0000
  16.250   1.2965   0.10654   0.09969  -0.0581   0.1013   1.0000
  16.500   1.2828   0.11293   0.10645  -0.0606   0.0982   1.0000
  16.750   1.2729   0.11879   0.11255  -0.0632   0.0950   1.0000
  17.000   1.2885   0.11989   0.11362  -0.0633   0.0904   1.0000
  17.250   1.2715   0.12729   0.12129  -0.0670   0.0885   1.0000
  17.500   1.2405   0.13797   0.13229  -0.0731   0.0876   1.0000
  17.750   1.1922   0.15380   0.14835  -0.0830   0.0874   1.0000
<< Back to NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)

Polar data table (+)

Polar graphs


<< Back to NASA/LANGLEY NLF(1)-0215F AIRFOIL (nlf0215f-il)