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BOEING HSNLF AIRFOIL (bacnlf-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: BOEING HSNLF AIRFOIL (bacnlf-il)
Reynolds number: 50,000
Max Cl/Cd: 28.29 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-bacnlf-il-50000-n5.txt
Download as CSV file: xf-bacnlf-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING HSNLF 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5467   0.09858   0.09143  -0.0395   1.0014   0.0392
  -9.500  -0.5520   0.09357   0.08648  -0.0418   1.0014   0.0383
  -9.250  -0.5611   0.08846   0.08143  -0.0442   1.0014   0.0373
  -9.000  -0.5747   0.08382   0.07684  -0.0460   1.0014   0.0365
  -8.750  -0.5930   0.07989   0.07295  -0.0464   1.0014   0.0357
  -8.500  -0.6127   0.07639   0.06945  -0.0455   1.0014   0.0349
  -8.250  -0.6294   0.07261   0.06559  -0.0445   1.0014   0.0340
  -8.000  -0.6436   0.06876   0.06154  -0.0430   1.0014   0.0329
  -6.500  -0.6300   0.04902   0.03969  -0.0326   1.0014   0.0310
  -6.250  -0.6171   0.04577   0.03604  -0.0312   1.0014   0.0310
  -5.750  -0.5852   0.03920   0.02879  -0.0289   1.0014   0.0322
  -5.500  -0.5668   0.03686   0.02626  -0.0280   1.0014   0.0351
  -5.250  -0.5455   0.03473   0.02379  -0.0268   1.0014   0.0381
  -5.000  -0.5217   0.03259   0.02121  -0.0254   1.0014   0.0402
  -4.750  -0.4972   0.03080   0.01909  -0.0237   1.0014   0.0424
  -4.500  -0.4760   0.02913   0.01742  -0.0223   1.0014   0.0493
  -4.250  -0.4545   0.02786   0.01586  -0.0204   1.0014   0.0555
  -4.000  -0.4353   0.02636   0.01434  -0.0191   1.0014   0.0644
  -3.750  -0.4144   0.02504   0.01289  -0.0182   1.0014   0.0814
  -3.500  -0.3921   0.02300   0.01126  -0.0184   1.0014   0.1345
  -3.250  -0.3869   0.02083   0.01174  -0.0135   1.0014   0.6490
  -3.000  -0.3886   0.02178   0.01290  -0.0036   1.0014   0.7506
  -2.750  -0.3946   0.02249   0.01370   0.0079   1.0014   0.8192
  -2.500  -0.3790   0.02238   0.01330   0.0114   1.0014   0.8443
  -2.250  -0.3557   0.02209   0.01267   0.0119   1.0014   0.8534
  -2.000  -0.3321   0.02182   0.01210   0.0122   1.0014   0.8623
  -1.750  -0.3083   0.02156   0.01153   0.0122   1.0014   0.8716
  -1.500  -0.2839   0.02134   0.01109   0.0123   1.0014   0.8800
  -1.250  -0.2592   0.02115   0.01071   0.0122   1.0014   0.8890
  -1.000  -0.2343   0.02099   0.01037   0.0120   1.0014   0.8988
  -0.750  -0.2080   0.02087   0.01007   0.0116   1.0014   0.9080
  -0.500  -0.1810   0.02078   0.00987   0.0109   1.0014   0.9183
  -0.250  -0.1528   0.02073   0.00973   0.0100   1.0014   0.9294
   0.000  -0.1230   0.02072   0.00965   0.0086   1.0014   0.9416
   0.250  -0.0914   0.02075   0.00960   0.0068   1.0014   0.9550
   0.500  -0.0596   0.02078   0.00961   0.0047   1.0014   0.9713
   0.750  -0.0350   0.02069   0.00953   0.0038   1.0014   0.9986
   1.000  -0.0149   0.02073   0.00954   0.0036   1.0014   0.9986
   1.250   0.0080   0.02087   0.00966   0.0029   1.0014   0.9986
   1.500   0.0324   0.02109   0.00987   0.0021   1.0014   0.9986
   1.750   0.0617   0.02142   0.01022   0.0003   0.9993   0.9986
   2.000   0.1005   0.02193   0.01077  -0.0033   0.9931   0.9986
   2.250   0.1402   0.02249   0.01140  -0.0069   0.9866   0.9986
   2.500   0.1768   0.02296   0.01200  -0.0099   0.9788   0.9986
   2.750   0.2175   0.02356   0.01273  -0.0136   0.9722   0.9986
   3.000   0.2539   0.02402   0.01336  -0.0164   0.9624   0.9986
   3.250   0.2932   0.02449   0.01409  -0.0195   0.9514   0.9986
   3.500   0.3334   0.02490   0.01477  -0.0227   0.9389   0.9986
   3.750   0.3735   0.02524   0.01542  -0.0256   0.9251   0.9986
   4.000   0.4206   0.02534   0.01596  -0.0293   0.9072   0.9986
   4.250   0.4781   0.02379   0.01494  -0.0319   0.8564   0.9986
   4.500   0.5186   0.02175   0.01334  -0.0302   0.7860   0.9986
   4.750   0.5819   0.02057   0.01083  -0.0296   0.3756   0.9986
   5.000   0.5868   0.02323   0.01205  -0.0266   0.1785   0.9986
   5.250   0.6051   0.02510   0.01355  -0.0254   0.1265   0.9986
   5.500   0.6280   0.02676   0.01516  -0.0247   0.0976   0.9986
   5.750   0.6561   0.02856   0.01697  -0.0248   0.0754   0.9986
   6.000   0.6900   0.03063   0.01922  -0.0253   0.0583   0.9986
   6.250   0.7234   0.03307   0.02188  -0.0259   0.0484   0.9986
   6.500   0.7509   0.03523   0.02427  -0.0259   0.0412   0.9986
   6.750   0.7786   0.03849   0.02782  -0.0257   0.0387   0.9986
   7.000   0.8019   0.04159   0.03156  -0.0245   0.0365   0.9986
   7.250   0.8197   0.04473   0.03525  -0.0228   0.0342   0.9986
   7.500   0.8338   0.04783   0.03881  -0.0210   0.0322   0.9986
   7.750   0.8446   0.05110   0.04250  -0.0191   0.0311   0.9986
   8.000   0.8516   0.05480   0.04663  -0.0168   0.0308   0.9986
   8.250   0.8548   0.05869   0.05094  -0.0144   0.0307   0.9986
   8.500   0.8539   0.06272   0.05535  -0.0120   0.0308   0.9986
   8.750   0.8490   0.06685   0.05980  -0.0096   0.0310   0.9986
   9.000   0.8404   0.07100   0.06423  -0.0075   0.0313   0.9986
   9.250   0.8274   0.07505   0.06849  -0.0055   0.0316   0.9986
   9.500   0.8112   0.07922   0.07281  -0.0041   0.0320   0.9986
   9.750   0.7939   0.08388   0.07758  -0.0040   0.0324   0.9986
  10.000   0.7770   0.08922   0.08300  -0.0056   0.0328   0.9986
  10.250   0.7625   0.09532   0.08914  -0.0089   0.0334   0.9986
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