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

BOEING 106 AIRFOIL (boe106-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: BOEING 106 AIRFOIL (boe106-il)
Reynolds number: 100,000
Max Cl/Cd: 52.76 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-boe106-il-100000-n5.txt
Download as CSV file: xf-boe106-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING 106 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4166   0.09011   0.08495  -0.0482   1.0000   0.0445
 -10.000  -0.4270   0.08432   0.07922  -0.0512   1.0000   0.0444
  -9.750  -0.4448   0.07749   0.07248  -0.0551   1.0000   0.0444
  -9.500  -0.4770   0.06959   0.06469  -0.0601   1.0000   0.0440
  -9.250  -0.5158   0.06588   0.06103  -0.0578   1.0000   0.0435
  -9.000  -0.5520   0.06248   0.05760  -0.0541   1.0000   0.0432
  -8.750  -0.5677   0.05439   0.04908  -0.0583   0.9906   0.0439
  -8.500  -0.5705   0.04649   0.04037  -0.0618   0.9794   0.0453
  -8.250  -0.5562   0.04157   0.03481  -0.0639   0.9723   0.0464
  -8.000  -0.5352   0.03923   0.03227  -0.0646   0.9646   0.0474
  -7.750  -0.5058   0.03742   0.03027  -0.0666   0.9599   0.0491
  -7.500  -0.4833   0.03538   0.02791  -0.0670   0.9521   0.0510
  -7.250  -0.4555   0.03279   0.02481  -0.0683   0.9467   0.0528
  -7.000  -0.4302   0.03067   0.02220  -0.0686   0.9397   0.0541
  -6.750  -0.4022   0.02896   0.02019  -0.0693   0.9331   0.0560
  -6.500  -0.3687   0.02770   0.01884  -0.0709   0.9288   0.0581
  -6.250  -0.3444   0.02659   0.01757  -0.0704   0.9200   0.0597
  -6.000  -0.3125   0.02534   0.01612  -0.0714   0.9146   0.0615
  -5.750  -0.2845   0.02436   0.01492  -0.0714   0.9076   0.0639
  -5.500  -0.2559   0.02341   0.01385  -0.0716   0.9008   0.0664
  -5.250  -0.2237   0.02253   0.01298  -0.0726   0.8961   0.0692
  -5.000  -0.2003   0.02188   0.01228  -0.0717   0.8869   0.0716
  -4.750  -0.1699   0.02113   0.01144  -0.0721   0.8813   0.0749
  -4.500  -0.1458   0.02056   0.01085  -0.0714   0.8727   0.0789
  -4.250  -0.1174   0.02003   0.01030  -0.0714   0.8663   0.0848
  -4.000  -0.0926   0.01953   0.00978  -0.0708   0.8582   0.0912
  -3.750  -0.0656   0.01903   0.00927  -0.0705   0.8510   0.1004
  -3.500  -0.0403   0.01858   0.00885  -0.0699   0.8431   0.1152
  -3.250  -0.0143   0.01811   0.00843  -0.0695   0.8355   0.1372
  -3.000   0.0110   0.01765   0.00805  -0.0690   0.8277   0.1664
  -2.750   0.0358   0.01713   0.00774  -0.0684   0.8196   0.2085
  -2.500   0.0600   0.01658   0.00751  -0.0678   0.8116   0.2876
  -2.250   0.0831   0.01595   0.00732  -0.0669   0.8033   0.3995
  -1.750   0.1275   0.01475   0.00737  -0.0631   0.7869   0.7320
  -1.500   0.1588   0.01465   0.00739  -0.0627   0.7787   0.8112
  -1.250   0.1955   0.01462   0.00731  -0.0635   0.7702   0.8618
  -1.000   0.2354   0.01466   0.00726  -0.0652   0.7607   0.9020
  -0.750   0.2795   0.01466   0.00711  -0.0677   0.7520   0.9384
  -0.500   0.3274   0.01470   0.00702  -0.0714   0.7407   0.9656
  -0.250   0.3804   0.01463   0.00679  -0.0763   0.7305   0.9866
   0.000   0.4283   0.01452   0.00653  -0.0805   0.7201   1.0000
   0.250   0.4495   0.01455   0.00647  -0.0794   0.7085   1.0000
   0.500   0.4717   0.01458   0.00639  -0.0784   0.6981   1.0000
   0.750   0.4944   0.01460   0.00631  -0.0775   0.6880   1.0000
   1.000   0.5160   0.01469   0.00632  -0.0764   0.6765   1.0000
   1.250   0.5386   0.01476   0.00630  -0.0754   0.6662   1.0000
   1.500   0.5616   0.01484   0.00628  -0.0744   0.6561   1.0000
   1.750   0.5836   0.01496   0.00636  -0.0734   0.6449   1.0000
   2.000   0.6066   0.01508   0.00640  -0.0724   0.6348   1.0000
   2.250   0.6296   0.01520   0.00644  -0.0715   0.6247   1.0000
   2.500   0.6519   0.01536   0.00657  -0.0704   0.6138   1.0000
   2.750   0.6751   0.01551   0.00665  -0.0695   0.6039   1.0000
   3.000   0.6980   0.01568   0.00677  -0.0685   0.5938   1.0000
   3.250   0.7205   0.01588   0.00695  -0.0675   0.5837   1.0000
   3.500   0.7442   0.01606   0.00705  -0.0667   0.5747   1.0000
   3.750   0.7664   0.01628   0.00729  -0.0657   0.5641   1.0000
   4.000   0.7894   0.01651   0.00749  -0.0647   0.5545   1.0000
   4.250   0.8129   0.01674   0.00768  -0.0639   0.5454   1.0000
   4.500   0.8354   0.01701   0.00798  -0.0630   0.5359   1.0000
   4.750   0.8597   0.01726   0.00817  -0.0622   0.5276   1.0000
   5.000   0.8812   0.01755   0.00850  -0.0611   0.5162   1.0000
   5.250   0.9034   0.01782   0.00877  -0.0601   0.5049   1.0000
   5.500   0.9259   0.01808   0.00900  -0.0591   0.4935   1.0000
   5.750   0.9472   0.01836   0.00928  -0.0579   0.4811   1.0000
   6.000   0.9681   0.01866   0.00961  -0.0567   0.4684   1.0000
   6.250   0.9892   0.01896   0.00995  -0.0556   0.4563   1.0000
   6.500   1.0101   0.01927   0.01026  -0.0544   0.4439   1.0000
   6.750   1.0310   0.01958   0.01057  -0.0532   0.4320   1.0000
   7.000   1.0509   0.01993   0.01100  -0.0519   0.4194   1.0000
   7.250   1.0704   0.02029   0.01141  -0.0506   0.4063   1.0000
   7.500   1.0891   0.02067   0.01181  -0.0492   0.3921   1.0000
   7.750   1.1069   0.02108   0.01223  -0.0476   0.3771   1.0000
   8.000   1.1237   0.02152   0.01271  -0.0459   0.3613   1.0000
   8.250   1.1396   0.02201   0.01322  -0.0441   0.3450   1.0000
   8.500   1.1546   0.02255   0.01378  -0.0423   0.3282   1.0000
   8.750   1.1679   0.02316   0.01439  -0.0402   0.3102   1.0000
   9.000   1.1791   0.02384   0.01507  -0.0379   0.2907   1.0000
   9.250   1.1882   0.02457   0.01583  -0.0353   0.2681   1.0000
   9.500   1.1944   0.02550   0.01671  -0.0326   0.2430   1.0000
   9.750   1.1983   0.02666   0.01776  -0.0298   0.2172   1.0000
  10.000   1.2013   0.02802   0.01900  -0.0272   0.1957   1.0000
  10.250   1.2038   0.02952   0.02042  -0.0249   0.1803   1.0000
  10.500   1.2069   0.03110   0.02198  -0.0228   0.1685   1.0000
  10.750   1.2090   0.03283   0.02367  -0.0208   0.1591   1.0000
  11.000   1.2137   0.03446   0.02536  -0.0192   0.1501   1.0000
  11.250   1.2169   0.03627   0.02718  -0.0176   0.1430   1.0000
  11.500   1.2221   0.03799   0.02899  -0.0163   0.1360   1.0000
  11.750   1.2242   0.03998   0.03095  -0.0149   0.1307   1.0000
  12.000   1.2314   0.04165   0.03279  -0.0138   0.1251   1.0000
  12.250   1.2360   0.04355   0.03475  -0.0128   0.1202   1.0000
  12.500   1.2399   0.04553   0.03671  -0.0118   0.1161   1.0000
  12.750   1.2469   0.04733   0.03868  -0.0109   0.1115   1.0000
  13.000   1.2523   0.04928   0.04071  -0.0101   0.1075   1.0000
  13.250   1.2572   0.05125   0.04269  -0.0093   0.1042   1.0000
  13.500   1.2641   0.05313   0.04470  -0.0086   0.1009   1.0000
  13.750   1.2690   0.05526   0.04698  -0.0080   0.0971   1.0000
  14.000   1.2712   0.05762   0.04942  -0.0077   0.0934   1.0000
  14.250   1.2728   0.06006   0.05187  -0.0074   0.0899   1.0000
  14.500   1.2728   0.06292   0.05496  -0.0075   0.0858   1.0000
  14.750   1.2712   0.06597   0.05812  -0.0079   0.0820   1.0000
  15.000   1.2694   0.06901   0.06113  -0.0082   0.0787   1.0000
  15.250   1.2675   0.07240   0.06477  -0.0088   0.0751   1.0000
  15.500   1.2657   0.07576   0.06830  -0.0094   0.0720   1.0000
  15.750   1.2629   0.07928   0.07189  -0.0102   0.0693   1.0000
  16.000   1.2605   0.08277   0.07546  -0.0109   0.0668   1.0000
  16.250   1.2566   0.08672   0.07964  -0.0119   0.0639   1.0000
  16.500   1.2517   0.09083   0.08390  -0.0132   0.0612   1.0000
  16.750   1.2464   0.09504   0.08819  -0.0147   0.0589   1.0000
  17.000   1.2408   0.09934   0.09258  -0.0161   0.0565   1.0000
  17.250   1.2324   0.10437   0.09785  -0.0180   0.0540   1.0000
  17.500   1.2242   0.10939   0.10302  -0.0201   0.0516   1.0000
  17.750   1.2170   0.11430   0.10800  -0.0223   0.0496   1.0000
  18.000   1.2090   0.11941   0.11319  -0.0246   0.0476   1.0000
  18.250   1.1958   0.12583   0.11986  -0.0278   0.0455   1.0000
  18.500   1.1837   0.13212   0.12631  -0.0310   0.0437   1.0000
  18.750   1.1747   0.13791   0.13217  -0.0342   0.0419   1.0000
  19.000   1.1683   0.14313   0.13741  -0.0371   0.0404   1.0000
  19.250   1.1475   0.15207   0.14660  -0.0422   0.0392   1.0000
<< Back to BOEING 106 AIRFOIL (boe106-il)

Polar data table (+)

Polar graphs


<< Back to BOEING 106 AIRFOIL (boe106-il)