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BOEING HSNLF AIRFOIL (bacnlf-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: BOEING HSNLF AIRFOIL (bacnlf-il)
Reynolds number: 100,000
Max Cl/Cd: 35.61 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-bacnlf-il-100000.txt
Download as CSV file: xf-bacnlf-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4311   0.10113   0.09646  -0.0349   1.0014   0.1032
  -9.750  -0.4466   0.09715   0.09256  -0.0368   1.0014   0.1073
  -9.500  -0.4762   0.09261   0.08811  -0.0406   1.0014   0.1086
  -9.250  -0.4536   0.08920   0.08469  -0.0362   1.0014   0.1136
  -9.000  -0.4605   0.08552   0.08106  -0.0361   1.0014   0.1183
  -8.750  -0.4889   0.08071   0.07635  -0.0388   1.0014   0.1212
  -8.500  -0.5200   0.07592   0.07163  -0.0408   1.0014   0.1216
  -8.250  -0.5016   0.07280   0.06854  -0.0368   1.0014   0.1284
  -8.000  -0.5253   0.06861   0.06442  -0.0369   1.0014   0.1301
  -7.750  -0.5574   0.06521   0.06107  -0.0355   1.0014   0.1295
  -7.500  -0.5817   0.06180   0.05765  -0.0341   1.0014   0.1315
  -7.250  -0.6114   0.05841   0.05415  -0.0329   1.0014   0.1346
  -5.500  -0.5925   0.03935   0.03136  -0.0263   1.0014   0.0560
  -5.250  -0.5711   0.03402   0.02593  -0.0258   1.0014   0.0486
  -5.000  -0.5448   0.03191   0.02305  -0.0239   1.0014   0.0422
  -4.750  -0.5204   0.02889   0.01973  -0.0232   1.0014   0.0408
  -4.500  -0.4949   0.02688   0.01739  -0.0223   1.0014   0.0414
  -4.250  -0.4686   0.02482   0.01509  -0.0216   1.0014   0.0445
  -4.000  -0.4436   0.02311   0.01341  -0.0211   1.0014   0.0511
  -3.750  -0.4183   0.02141   0.01173  -0.0204   1.0014   0.0610
  -3.500  -0.3932   0.01987   0.01027  -0.0199   1.0014   0.0782
  -3.250  -0.3648   0.01795   0.00860  -0.0204   1.0014   0.1318
  -3.000  -0.3560   0.01657   0.00987  -0.0144   1.0014   0.7211
  -2.750  -0.3495   0.01719   0.01052  -0.0077   1.0014   0.7676
  -2.500  -0.3535   0.01780   0.01124   0.0021   1.0014   0.8174
  -2.250  -0.3599   0.01805   0.01154   0.0121   1.0014   0.8588
  -2.000  -0.3554   0.01795   0.01137   0.0184   1.0014   0.8867
  -1.750  -0.3336   0.01771   0.01090   0.0192   1.0014   0.8972
  -1.500  -0.3101   0.01748   0.01050   0.0194   1.0014   0.9054
  -1.250  -0.2852   0.01729   0.01015   0.0192   1.0014   0.9142
  -1.000  -0.2605   0.01713   0.00987   0.0191   1.0014   0.9228
  -0.750  -0.2343   0.01701   0.00963   0.0185   1.0014   0.9316
  -0.500  -0.2072   0.01693   0.00943   0.0177   1.0014   0.9414
  -0.250  -0.1772   0.01691   0.00933   0.0163   1.0014   0.9511
   0.000  -0.1443   0.01696   0.00932   0.0142   1.0014   0.9611
   0.250  -0.1085   0.01706   0.00937   0.0114   1.0014   0.9718
   0.500  -0.0718   0.01719   0.00945   0.0083   1.0014   0.9834
   0.750  -0.0415   0.01721   0.00947   0.0061   1.0014   0.9986
   1.000  -0.0249   0.01714   0.00940   0.0063   1.0014   0.9986
   1.250  -0.0009   0.01725   0.00951   0.0051   1.0014   0.9986
   1.500   0.0264   0.01748   0.00974   0.0033   1.0014   0.9986
   1.750   0.0550   0.01780   0.01007   0.0014   1.0014   0.9986
   2.000   0.0837   0.01817   0.01047  -0.0006   1.0014   0.9986
   2.250   0.1121   0.01861   0.01094  -0.0024   1.0014   0.9986
   2.500   0.1397   0.01910   0.01150  -0.0040   1.0014   0.9986
   2.750   0.1763   0.01975   0.01223  -0.0074   0.9977   0.9986
   3.000   0.2300   0.02058   0.01319  -0.0138   0.9870   0.9986
   3.250   0.2807   0.02125   0.01403  -0.0195   0.9746   0.9986
   3.500   0.3377   0.02172   0.01478  -0.0258   0.9574   0.9986
   3.750   0.4308   0.02060   0.01412  -0.0360   0.9158   0.9986
   4.000   0.5051   0.01868   0.01267  -0.0417   0.8812   0.9986
   4.250   0.5625   0.01586   0.01029  -0.0423   0.8286   0.9986
   4.500   0.6004   0.01686   0.00822  -0.0385   0.1959   0.9986
   4.750   0.6148   0.01884   0.00963  -0.0363   0.1327   0.9986
   5.000   0.6351   0.02056   0.01115  -0.0350   0.1010   0.9986
   5.250   0.6579   0.02230   0.01275  -0.0343   0.0694   0.9986
   5.500   0.6848   0.02434   0.01469  -0.0341   0.0536   0.9986
   5.750   0.7171   0.02705   0.01760  -0.0343   0.0484   0.9986
   6.000   0.7475   0.02980   0.02076  -0.0340   0.0467   0.9986
   6.250   0.7741   0.03294   0.02436  -0.0330   0.0469   0.9986
   6.500   0.7966   0.03651   0.02840  -0.0315   0.0484   0.9986
   6.750   0.8154   0.04085   0.03309  -0.0300   0.0505   0.9986
   7.000   0.8318   0.04568   0.03878  -0.0264   0.0617   0.9986
   9.750   0.8061   0.09839   0.09419  -0.0107   0.1035   0.9986
  10.000   0.7824   0.10374   0.09958  -0.0145   0.1032   0.9986
  10.250   0.7630   0.11070   0.10649  -0.0218   0.1028   0.9986
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