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

BOEING 103 AIRFOIL (boe103-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: BOEING 103 AIRFOIL (boe103-il)
Reynolds number: 50,000
Max Cl/Cd: 32.58 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-boe103-il-50000.txt
Download as CSV file: xf-boe103-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING 103 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3430   0.10999   0.10301  -0.0267   1.0000   0.2656
  -8.750  -0.3256   0.10581   0.09885  -0.0252   1.0000   0.2758
  -8.500  -0.3641   0.10655   0.09983  -0.0246   1.0000   0.2834
  -8.250  -0.3405   0.10212   0.09538  -0.0227   1.0000   0.2982
  -8.000  -0.3279   0.09855   0.09186  -0.0210   1.0000   0.3088
  -7.750  -0.3434   0.09701   0.09047  -0.0190   1.0000   0.3193
  -7.500  -0.3599   0.09617   0.08977  -0.0160   1.0000   0.3321
  -7.250  -0.3435   0.09256   0.08619  -0.0141   1.0000   0.3451
  -7.000  -0.3464   0.09035   0.08408  -0.0113   1.0000   0.3574
  -6.750  -0.3629   0.08905   0.08292  -0.0074   1.0000   0.3703
  -6.500  -0.4039   0.08974   0.08381  -0.0014   1.0000   0.3823
  -6.250  -0.4077   0.08773   0.08190   0.0023   1.0000   0.3986
  -6.000  -0.4221   0.08639   0.08067   0.0064   1.0000   0.4158
  -5.750  -0.3874   0.08252   0.07676   0.0088   1.0000   0.4443
  -5.500  -0.4151   0.08214   0.07654   0.0146   1.0000   0.4652
  -5.000  -0.4598   0.05374   0.04637  -0.0313   1.0000   0.1679
  -4.750  -0.4432   0.05060   0.04303  -0.0313   1.0000   0.1640
  -4.500  -0.4247   0.04736   0.03950  -0.0317   1.0000   0.1607
  -4.250  -0.4023   0.04382   0.03537  -0.0328   1.0000   0.1552
  -4.000  -0.3768   0.04110   0.03174  -0.0335   1.0000   0.1510
  -3.750  -0.3542   0.03905   0.02935  -0.0335   1.0000   0.1512
  -3.500  -0.3334   0.03745   0.02771  -0.0333   1.0000   0.1548
  -3.250  -0.3113   0.03620   0.02626  -0.0331   1.0000   0.1594
  -3.000  -0.2876   0.03500   0.02469  -0.0329   1.0000   0.1635
  -2.750  -0.2637   0.03388   0.02325  -0.0327   1.0000   0.1680
  -2.500  -0.2414   0.03297   0.02234  -0.0324   1.0000   0.1758
  -2.250  -0.2184   0.03222   0.02152  -0.0321   1.0000   0.1887
  -2.000  -0.1952   0.03150   0.02083  -0.0318   1.0000   0.2083
  -1.750  -0.1682   0.03046   0.01997  -0.0318   1.0000   0.2534
  -1.500  -0.1358   0.02854   0.01947  -0.0329   1.0000   0.4375
  -1.250  -0.1300   0.02670   0.01966  -0.0263   1.0000   1.0000
  -1.000  -0.1090   0.02731   0.01968  -0.0265   1.0000   1.0000
  -0.750  -0.0614   0.02851   0.02033  -0.0316   0.9893   1.0000
  -0.500  -0.0151   0.02969   0.02109  -0.0364   0.9771   1.0000
  -0.250   0.0290   0.03078   0.02184  -0.0407   0.9639   1.0000
   0.000   0.0718   0.03180   0.02258  -0.0446   0.9498   1.0000
   0.250   0.1136   0.03275   0.02329  -0.0482   0.9350   1.0000
   0.500   0.1540   0.03365   0.02399  -0.0514   0.9199   1.0000
   0.750   0.1925   0.03450   0.02468  -0.0542   0.9048   1.0000
   1.000   0.2297   0.03533   0.02535  -0.0566   0.8897   1.0000
   1.250   0.2661   0.03612   0.02602  -0.0588   0.8746   1.0000
   1.500   0.3021   0.03688   0.02668  -0.0608   0.8594   1.0000
   1.750   0.3375   0.03759   0.02731  -0.0625   0.8439   1.0000
   2.000   0.3723   0.03826   0.02792  -0.0641   0.8281   1.0000
   2.250   0.4063   0.03889   0.02850  -0.0653   0.8121   1.0000
   2.500   0.4395   0.03948   0.02906  -0.0663   0.7959   1.0000
   2.750   0.4720   0.04004   0.02960  -0.0671   0.7795   1.0000
   3.000   0.5035   0.04055   0.03010  -0.0676   0.7627   1.0000
   3.250   0.5347   0.04104   0.03060  -0.0680   0.7462   1.0000
   3.500   0.5659   0.04146   0.03103  -0.0682   0.7296   1.0000
   3.750   0.5971   0.04184   0.03143  -0.0683   0.7134   1.0000
   4.000   0.6291   0.04210   0.03173  -0.0683   0.6974   1.0000
   4.250   0.6619   0.04224   0.03190  -0.0683   0.6821   1.0000
   4.500   0.7224   0.04066   0.03039  -0.0702   0.6713   1.0000
   4.750   0.7525   0.04066   0.03045  -0.0695   0.6556   1.0000
   5.000   0.7831   0.04055   0.03040  -0.0687   0.6398   1.0000
   5.250   0.8148   0.04030   0.03022  -0.0678   0.6240   1.0000
   5.500   0.8496   0.03972   0.02969  -0.0670   0.6080   1.0000
   5.750   0.8932   0.03830   0.02831  -0.0666   0.5913   1.0000
   6.000   0.9267   0.03733   0.02739  -0.0652   0.5713   1.0000
   6.250   0.9565   0.03655   0.02662  -0.0635   0.5496   1.0000
   6.500   1.0011   0.03480   0.02482  -0.0631   0.5280   1.0000
   6.750   1.0295   0.03441   0.02441  -0.0617   0.5064   1.0000
   7.000   1.0520   0.03447   0.02448  -0.0599   0.4840   1.0000
   7.250   1.0870   0.03384   0.02371  -0.0592   0.4609   1.0000
   7.500   1.1022   0.03446   0.02437  -0.0567   0.4365   1.0000
   7.750   1.1290   0.03465   0.02439  -0.0554   0.4114   1.0000
   8.000   1.1493   0.03532   0.02495  -0.0536   0.3855   1.0000
   8.250   1.1670   0.03639   0.02595  -0.0517   0.3611   1.0000
   8.500   1.1870   0.03755   0.02699  -0.0502   0.3390   1.0000
   8.750   1.2097   0.03895   0.02826  -0.0493   0.3206   1.0000
   9.000   1.2242   0.04085   0.03026  -0.0476   0.3064   1.0000
   9.250   1.2363   0.04300   0.03260  -0.0458   0.2951   1.0000
   9.500   1.2578   0.04481   0.03440  -0.0451   0.2848   1.0000
   9.750   1.2638   0.04708   0.03691  -0.0427   0.2761   1.0000
  10.000   1.2775   0.04923   0.03916  -0.0413   0.2684   1.0000
  10.250   1.2730   0.05219   0.04245  -0.0382   0.2628   1.0000
  10.500   1.3038   0.05394   0.04416  -0.0387   0.2562   1.0000
  10.750   1.2817   0.05791   0.04853  -0.0344   0.2536   1.0000
  11.000   1.2465   0.06253   0.05350  -0.0296   0.2519   1.0000
  11.250   1.1937   0.06870   0.05991  -0.0252   0.2520   1.0000
  11.500   1.1090   0.08003   0.07138  -0.0245   0.2543   1.0000
<< Back to BOEING 103 AIRFOIL (boe103-il)

Polar data table (+)

Polar graphs


<< Back to BOEING 103 AIRFOIL (boe103-il)