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