BOEING 103 AIRFOIL (boe103-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 103 AIRFOIL (boe103-il) Reynolds number: 100,000 Max Cl/Cd: 53.11 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-boe103-il-100000.txt Download as CSV file: xf-boe103-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 103 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3449 0.10209 0.09707 -0.0363 1.0000 0.1293
-9.000 -0.3721 0.10028 0.09541 -0.0385 1.0000 0.1336
-8.750 -0.4203 0.09939 0.09475 -0.0393 1.0000 0.1345
-8.500 -0.3928 0.09409 0.08943 -0.0367 1.0000 0.1367
-8.250 -0.3773 0.09146 0.08681 -0.0338 1.0000 0.1397
-8.000 -0.3852 0.08962 0.08505 -0.0313 1.0000 0.1429
-7.750 -0.4085 0.08821 0.08376 -0.0284 1.0000 0.1453
-7.500 -0.4401 0.08677 0.08245 -0.0259 1.0000 0.1471
-7.250 -0.5107 0.08453 0.08015 -0.0314 1.0000 0.1504
-7.000 -0.5022 0.08025 0.07601 -0.0282 1.0000 0.1520
-6.750 -0.4899 0.07841 0.07426 -0.0234 1.0000 0.1545
-6.500 -0.4906 0.07649 0.07238 -0.0209 1.0000 0.1586
-6.250 -0.5188 0.07203 0.06766 -0.0265 1.0000 0.1678
-6.000 -0.5087 0.06966 0.06547 -0.0225 1.0000 0.1703
-5.750 -0.4863 0.06498 0.06051 -0.0299 0.9942 0.1849
-5.500 -0.4545 0.04657 0.04055 -0.0417 0.9881 0.1068
-5.250 -0.4163 0.03977 0.03252 -0.0453 0.9835 0.0917
-5.000 -0.3829 0.03662 0.02900 -0.0474 0.9772 0.0910
-4.750 -0.3434 0.03365 0.02564 -0.0504 0.9723 0.0901
-4.500 -0.3078 0.03149 0.02306 -0.0522 0.9657 0.0901
-4.250 -0.2671 0.02983 0.02096 -0.0548 0.9599 0.0920
-4.000 -0.2313 0.02839 0.01946 -0.0568 0.9533 0.0950
-3.750 -0.1918 0.02725 0.01819 -0.0590 0.9468 0.0976
-3.500 -0.1531 0.02630 0.01710 -0.0610 0.9405 0.1014
-3.250 -0.1168 0.02541 0.01614 -0.0626 0.9331 0.1070
-3.000 -0.0738 0.02472 0.01550 -0.0654 0.9279 0.1178
-2.750 -0.0444 0.02403 0.01494 -0.0658 0.9187 0.1308
-2.500 -0.0016 0.02269 0.01404 -0.0687 0.9142 0.1959
-2.250 0.0245 0.02170 0.01371 -0.0689 0.9049 0.3064
-2.000 0.0585 0.02035 0.01375 -0.0698 0.8997 0.6063
-1.750 0.0770 0.02004 0.01410 -0.0661 0.8911 0.8059
-1.500 0.2093 0.01982 0.01380 -0.0835 0.8980 1.0000
-1.250 0.2300 0.01984 0.01364 -0.0827 0.8853 1.0000
-1.000 0.2659 0.01967 0.01327 -0.0842 0.8749 1.0000
-0.750 0.3162 0.01917 0.01258 -0.0879 0.8673 1.0000
-0.500 0.3430 0.01906 0.01233 -0.0875 0.8550 1.0000
-0.250 0.3900 0.01861 0.01173 -0.0904 0.8493 1.0000
0.000 0.4087 0.01874 0.01176 -0.0887 0.8366 1.0000
0.250 0.4351 0.01873 0.01165 -0.0882 0.8262 1.0000
0.500 0.4714 0.01845 0.01127 -0.0892 0.8183 1.0000
0.750 0.4925 0.01857 0.01131 -0.0877 0.8062 1.0000
1.000 0.5260 0.01835 0.01100 -0.0881 0.7978 1.0000
1.250 0.5522 0.01831 0.01090 -0.0873 0.7868 1.0000
1.500 0.5755 0.01837 0.01091 -0.0861 0.7748 1.0000
1.750 0.6072 0.01818 0.01064 -0.0862 0.7653 1.0000
2.000 0.6346 0.01810 0.01050 -0.0855 0.7537 1.0000
2.250 0.6580 0.01815 0.01052 -0.0842 0.7406 1.0000
2.500 0.6842 0.01813 0.01045 -0.0834 0.7281 1.0000
2.750 0.7131 0.01801 0.01027 -0.0829 0.7159 1.0000
3.000 0.7435 0.01785 0.01003 -0.0826 0.7035 1.0000
3.250 0.7681 0.01788 0.01002 -0.0815 0.6884 1.0000
3.500 0.7935 0.01788 0.00997 -0.0805 0.6726 1.0000
3.750 0.8199 0.01783 0.00985 -0.0795 0.6557 1.0000
4.000 0.8436 0.01785 0.00980 -0.0781 0.6366 1.0000
4.250 0.8658 0.01789 0.00978 -0.0765 0.6153 1.0000
4.500 0.8897 0.01787 0.00965 -0.0751 0.5936 1.0000
4.750 0.9122 0.01793 0.00960 -0.0736 0.5714 1.0000
5.000 0.9328 0.01809 0.00971 -0.0719 0.5489 1.0000
5.250 0.9552 0.01830 0.00985 -0.0706 0.5288 1.0000
5.500 0.9779 0.01856 0.01001 -0.0694 0.5098 1.0000
5.750 0.9981 0.01888 0.01030 -0.0678 0.4894 1.0000
6.000 1.0185 0.01921 0.01060 -0.0663 0.4692 1.0000
6.250 1.0389 0.01956 0.01088 -0.0649 0.4494 1.0000
6.500 1.0592 0.01995 0.01116 -0.0634 0.4300 1.0000
6.750 1.0774 0.02041 0.01159 -0.0617 0.4091 1.0000
7.000 1.0955 0.02091 0.01204 -0.0600 0.3884 1.0000
7.250 1.1135 0.02149 0.01249 -0.0583 0.3679 1.0000
7.500 1.1296 0.02215 0.01309 -0.0563 0.3457 1.0000
7.750 1.1446 0.02291 0.01374 -0.0542 0.3224 1.0000
8.000 1.1583 0.02379 0.01452 -0.0520 0.2983 1.0000
8.250 1.1712 0.02473 0.01534 -0.0498 0.2749 1.0000
8.500 1.1860 0.02577 0.01617 -0.0479 0.2552 1.0000
8.750 1.2011 0.02677 0.01711 -0.0460 0.2388 1.0000
9.000 1.2184 0.02775 0.01802 -0.0446 0.2259 1.0000
9.250 1.2357 0.02875 0.01907 -0.0432 0.2152 1.0000
9.500 1.2571 0.02987 0.02015 -0.0425 0.2071 1.0000
9.750 1.2772 0.03091 0.02122 -0.0416 0.1997 1.0000
10.000 1.2992 0.03208 0.02238 -0.0411 0.1932 1.0000
10.250 1.3150 0.03309 0.02350 -0.0396 0.1865 1.0000
10.500 1.3388 0.03425 0.02453 -0.0395 0.1801 1.0000
10.750 1.3498 0.03535 0.02589 -0.0373 0.1748 1.0000
11.000 1.3691 0.03642 0.02696 -0.0365 0.1695 1.0000
11.250 1.3934 0.03792 0.02847 -0.0366 0.1650 1.0000
11.500 1.4034 0.03929 0.03013 -0.0344 0.1613 1.0000
11.750 1.4179 0.04065 0.03163 -0.0330 0.1575 1.0000
12.000 1.4522 0.04229 0.03307 -0.0347 0.1522 1.0000
12.250 1.4492 0.04369 0.03485 -0.0309 0.1495 1.0000
12.500 1.4500 0.04518 0.03659 -0.0278 0.1459 1.0000
12.750 1.4591 0.04639 0.03787 -0.0260 0.1418 1.0000
13.000 1.4806 0.04804 0.03946 -0.0260 0.1373 1.0000
13.250 1.4705 0.04992 0.04169 -0.0222 0.1350 1.0000
13.500 1.4657 0.05191 0.04393 -0.0194 0.1324 1.0000
13.750 1.4651 0.05376 0.04595 -0.0173 0.1295 1.0000
14.000 1.5004 0.05482 0.04674 -0.0185 0.1244 1.0000
14.250 1.4805 0.05742 0.04971 -0.0150 0.1232 1.0000
14.500 1.4609 0.06051 0.05314 -0.0123 0.1218 1.0000
14.750 1.4417 0.06400 0.05692 -0.0104 0.1203 1.0000
15.000 1.4228 0.06783 0.06101 -0.0091 0.1188 1.0000
15.250 1.4055 0.07181 0.06522 -0.0084 0.1173 1.0000
15.500 1.3868 0.07623 0.06984 -0.0083 0.1160 1.0000
15.750 1.3469 0.08349 0.07740 -0.0094 0.1159 1.0000
16.000 1.3464 0.08631 0.08029 -0.0096 0.1137 1.0000
16.250 1.3262 0.09191 0.08604 -0.0110 0.1127 1.0000
16.500 1.3051 0.09814 0.09240 -0.0129 0.1117 1.0000
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