BOEING AIRFOIL J (no closed TE) (bacj-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING AIRFOIL J (no closed TE) (bacj-il) Reynolds number: 100,000 Max Cl/Cd: 27.04 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-bacj-il-100000.txt Download as CSV file: xf-bacj-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING AIRFOIL J
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 -0.5502 0.10692 0.10141 -0.0282 1.0000 0.1540
-9.750 -0.5609 0.10297 0.09751 -0.0293 1.0000 0.1558
-9.500 -0.6727 0.09750 0.09231 -0.0380 1.0000 0.1598
-9.250 -0.5756 0.09549 0.09014 -0.0292 1.0000 0.1653
-9.000 -0.5785 0.09244 0.08714 -0.0284 1.0000 0.1705
-8.750 -0.6259 0.08753 0.08239 -0.0316 1.0000 0.1757
-8.500 -0.7797 0.06278 0.05704 -0.0377 1.0000 0.0899
-8.250 -0.8179 0.05410 0.04748 -0.0343 1.0000 0.0785
-8.000 -0.8155 0.05017 0.04334 -0.0324 1.0000 0.0775
-7.750 -0.8121 0.04650 0.03932 -0.0304 1.0000 0.0772
-7.500 -0.8055 0.04313 0.03555 -0.0284 1.0000 0.0773
-7.250 -0.7952 0.04015 0.03216 -0.0266 1.0000 0.0777
-7.000 -0.7820 0.03716 0.02874 -0.0250 1.0000 0.0779
-6.750 -0.7660 0.03465 0.02579 -0.0234 1.0000 0.0783
-6.500 -0.7478 0.03253 0.02326 -0.0220 1.0000 0.0791
-6.250 -0.7282 0.03010 0.02060 -0.0209 1.0000 0.0812
-6.000 -0.7083 0.02885 0.01938 -0.0200 1.0000 0.0865
-5.750 -0.6869 0.02748 0.01782 -0.0188 1.0000 0.0898
-5.500 -0.6647 0.02626 0.01630 -0.0175 1.0000 0.0934
-5.250 -0.6435 0.02470 0.01486 -0.0165 1.0000 0.0990
-5.000 -0.6222 0.02390 0.01395 -0.0153 1.0000 0.1081
-4.750 -0.6017 0.02272 0.01291 -0.0141 1.0000 0.1183
-4.500 -0.5815 0.02177 0.01203 -0.0128 1.0000 0.1339
-4.250 -0.5616 0.02086 0.01123 -0.0115 1.0000 0.1564
-4.000 -0.5431 0.01981 0.01043 -0.0100 1.0000 0.1924
-3.750 -0.5260 0.01869 0.00973 -0.0083 1.0000 0.2461
-3.500 -0.5117 0.01740 0.00912 -0.0063 1.0000 0.3385
-3.250 -0.5069 0.01652 0.00976 -0.0011 1.0000 0.5853
-3.000 -0.4929 0.01715 0.01045 0.0030 1.0000 0.6776
-2.750 -0.4789 0.01792 0.01123 0.0073 1.0000 0.7193
-2.500 -0.4647 0.01851 0.01178 0.0113 1.0000 0.7491
-2.250 -0.4509 0.01893 0.01215 0.0153 1.0000 0.7729
-2.000 -0.4373 0.01917 0.01231 0.0189 1.0000 0.7959
-1.750 -0.4236 0.01923 0.01233 0.0225 1.0000 0.8137
-1.500 -0.4103 0.01920 0.01227 0.0260 1.0000 0.8314
-1.250 -0.3966 0.01908 0.01210 0.0293 1.0000 0.8484
-1.000 -0.3827 0.01891 0.01188 0.0324 1.0000 0.8650
-0.750 -0.3672 0.01870 0.01163 0.0349 1.0000 0.8802
-0.500 -0.3523 0.01846 0.01137 0.0376 1.0000 0.8989
-0.250 -0.3302 0.01825 0.01115 0.0389 1.0000 0.9196
0.000 -0.2983 0.01820 0.01107 0.0380 1.0000 0.9390
0.250 -0.2365 0.01869 0.01155 0.0313 1.0000 0.9613
0.500 -0.1854 0.01907 0.01190 0.0259 1.0000 0.9697
0.750 -0.1408 0.01936 0.01218 0.0215 1.0000 0.9753
1.000 -0.0999 0.01963 0.01244 0.0177 1.0000 0.9823
1.250 -0.0590 0.01989 0.01273 0.0139 1.0000 0.9884
1.500 -0.0279 0.02025 0.01312 0.0117 1.0000 1.0000
1.750 -0.0243 0.01996 0.01285 0.0145 1.0000 1.0000
2.000 -0.0164 0.01983 0.01274 0.0165 1.0000 1.0000
2.250 -0.0033 0.01988 0.01281 0.0177 1.0000 1.0000
2.500 0.0137 0.02009 0.01305 0.0182 1.0000 1.0000
2.750 0.0477 0.02076 0.01379 0.0155 0.9956 1.0000
3.000 0.1249 0.02198 0.01519 0.0052 0.9729 1.0000
3.250 0.1852 0.02250 0.01588 -0.0015 0.9533 1.0000
3.500 0.5212 0.01947 0.00974 -0.0418 0.1217 0.9882
3.750 0.5543 0.02054 0.01071 -0.0432 0.1066 0.9932
4.000 0.5884 0.02176 0.01180 -0.0448 0.0971 0.9983
4.250 0.6174 0.02283 0.01287 -0.0453 0.0910 1.0000
4.500 0.6393 0.02373 0.01381 -0.0443 0.0868 1.0000
4.750 0.6644 0.02493 0.01503 -0.0439 0.0842 1.0000
5.000 0.6936 0.02679 0.01684 -0.0445 0.0808 1.0000
5.250 0.7130 0.02826 0.01852 -0.0430 0.0787 1.0000
5.500 0.7269 0.02943 0.01995 -0.0404 0.0774 1.0000
5.750 0.7409 0.03098 0.02172 -0.0380 0.0777 1.0000
6.000 0.7528 0.03286 0.02386 -0.0351 0.0780 1.0000
6.250 0.7655 0.03520 0.02643 -0.0326 0.0788 1.0000
6.500 0.7824 0.03830 0.02972 -0.0310 0.0801 1.0000
6.750 0.7960 0.04072 0.03243 -0.0287 0.0802 1.0000
7.000 0.8006 0.04279 0.03549 -0.0230 0.0883 1.0000
8.000 0.7774 0.06749 0.06270 -0.0102 0.1932 1.0000
8.250 0.7882 0.07037 0.06555 -0.0090 0.1847 1.0000
8.500 0.7697 0.07423 0.06948 -0.0069 0.1766 1.0000
8.750 0.7490 0.07713 0.07243 -0.0047 0.1708 1.0000
9.000 0.8140 0.08388 0.07892 -0.0071 0.1626 1.0000
9.250 0.7484 0.08516 0.08041 -0.0032 0.1607 1.0000
9.500 0.7004 0.09051 0.08578 -0.0070 0.1565 1.0000
9.750 0.6946 0.09511 0.09035 -0.0094 0.1510 1.0000
10.000 0.7532 0.09787 0.09305 -0.0027 0.1447 1.0000
10.250 0.6872 0.10521 0.10036 -0.0134 0.1426 1.0000
10.500 0.6724 0.11028 0.10537 -0.0181 0.1358 1.0000
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Polar data table (+)
Polar graphs
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