BOEING HSNLF AIRFOIL (bacnlf-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: BOEING HSNLF AIRFOIL (bacnlf-il) Reynolds number: 50,000 Max Cl/Cd: 26.15 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-bacnlf-il-50000.txt Download as CSV file: xf-bacnlf-il-50000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5147 0.10888 0.10192 -0.0120 1.0014 0.3435
-8.750 -0.4958 0.10479 0.09783 -0.0101 1.0014 0.3674
-6.500 -0.6563 0.05903 0.05197 -0.0332 1.0014 0.1606
-6.250 -0.6460 0.05378 0.04549 -0.0336 1.0014 0.1226
-6.000 -0.6276 0.04897 0.04026 -0.0327 1.0014 0.1079
-5.750 -0.6097 0.04476 0.03575 -0.0317 1.0014 0.1010
-5.500 -0.5888 0.04203 0.03202 -0.0302 1.0014 0.0952
-5.250 -0.5669 0.03902 0.02849 -0.0291 1.0014 0.0953
-5.000 -0.5428 0.03609 0.02514 -0.0280 1.0014 0.0953
-4.750 -0.5167 0.03334 0.02202 -0.0268 1.0014 0.0953
-4.500 -0.4905 0.03056 0.01924 -0.0257 1.0014 0.1002
-4.250 -0.4644 0.02864 0.01721 -0.0241 1.0014 0.1118
-4.000 -0.4392 0.02698 0.01548 -0.0219 1.0014 0.1234
-3.750 -0.4181 0.02519 0.01387 -0.0203 1.0014 0.1516
-3.500 -0.1948 0.02744 0.01779 -0.0257 1.0014 0.9785
-3.250 -0.1382 0.02638 0.01620 -0.0333 1.0014 0.9942
-3.000 -0.1149 0.02573 0.01533 -0.0344 1.0014 0.9986
-2.750 -0.1071 0.02528 0.01476 -0.0324 1.0014 0.9986
-2.500 -0.0996 0.02487 0.01424 -0.0302 1.0014 0.9986
-2.250 -0.0926 0.02448 0.01376 -0.0280 1.0014 0.9986
-2.000 -0.0861 0.02411 0.01327 -0.0256 1.0014 0.9986
-1.750 -0.0801 0.02376 0.01286 -0.0231 1.0014 0.9986
-1.500 -0.0747 0.02342 0.01246 -0.0205 1.0014 0.9986
-1.250 -0.0699 0.02309 0.01209 -0.0178 1.0014 0.9986
-1.000 -0.0659 0.02277 0.01173 -0.0149 1.0014 0.9986
-0.750 -0.0626 0.02244 0.01138 -0.0120 1.0014 0.9986
-0.500 -0.0601 0.02211 0.01104 -0.0089 1.0014 0.9986
-0.250 -0.0585 0.02177 0.01069 -0.0057 1.0014 0.9986
0.000 -0.0573 0.02142 0.01030 -0.0024 1.0014 0.9986
0.250 -0.0559 0.02106 0.00995 0.0008 1.0014 0.9986
0.500 -0.0496 0.02079 0.00966 0.0030 1.0014 0.9986
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 0.0573 0.02137 0.01016 0.0011 1.0014 0.9986
2.000 0.0824 0.02171 0.01052 0.0002 1.0014 0.9986
2.250 0.1072 0.02210 0.01096 -0.0006 1.0014 0.9986
2.500 0.1318 0.02254 0.01146 -0.0014 1.0014 0.9986
2.750 0.1558 0.02303 0.01205 -0.0021 1.0014 0.9986
3.000 0.1794 0.02356 0.01268 -0.0028 1.0014 0.9986
3.250 0.2025 0.02415 0.01339 -0.0034 1.0014 0.9986
3.500 0.2250 0.02479 0.01418 -0.0039 1.0014 0.9986
3.750 0.2469 0.02550 0.01512 -0.0045 1.0014 0.9986
4.000 0.2681 0.02628 0.01609 -0.0050 1.0014 0.9986
4.250 0.2887 0.02714 0.01717 -0.0055 1.0014 0.9986
4.500 0.3084 0.02809 0.01836 -0.0060 1.0014 0.9986
4.750 0.3272 0.02916 0.01969 -0.0066 1.0014 0.9986
5.000 0.5765 0.02205 0.01524 -0.0297 0.7564 0.9986
5.250 0.6272 0.02630 0.01536 -0.0261 0.1844 0.9986
5.500 0.6807 0.02978 0.01871 -0.0291 0.1371 0.9986
5.750 0.7163 0.03305 0.02207 -0.0297 0.1139 0.9986
6.000 0.7465 0.03644 0.02566 -0.0295 0.1027 0.9986
6.250 0.7705 0.03933 0.02922 -0.0280 0.0967 0.9986
6.500 0.7931 0.04301 0.03345 -0.0264 0.0974 0.9986
6.750 0.8109 0.04685 0.03785 -0.0244 0.0992 0.9986
7.000 0.8257 0.05083 0.04229 -0.0223 0.1010 0.9986
7.250 0.8391 0.05511 0.04691 -0.0206 0.1028 0.9986
7.500 0.7807 0.04681 0.03993 -0.0099 0.1140 0.9986
8.500 0.7011 0.07200 0.06653 -0.0046 0.1955 0.9986
8.750 0.6577 0.07938 0.07389 -0.0081 0.2183 0.9986
9.000 0.7177 0.10741 0.10151 -0.0479 0.3811 0.9986
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Polar data table (+)
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