BOEING 106 AIRFOIL (boe106-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 106 AIRFOIL (boe106-il) Reynolds number: 50,000 Max Cl/Cd: 29.69 at α=10.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-boe106-il-50000.txt Download as CSV file: xf-boe106-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING 106 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3636 0.11880 0.11165 -0.0273 1.0000 0.2571
-9.750 -0.3376 0.11336 0.10619 -0.0262 1.0000 0.2656
-9.500 -0.3662 0.11346 0.10647 -0.0276 1.0000 0.2743
-9.250 -0.3398 0.10857 0.10154 -0.0261 1.0000 0.2882
-9.000 -0.3277 0.10494 0.09794 -0.0252 1.0000 0.3001
-8.750 -0.3290 0.10235 0.09544 -0.0247 1.0000 0.3118
-8.500 -0.3582 0.10254 0.09582 -0.0238 1.0000 0.3235
-8.250 -0.3396 0.09858 0.09186 -0.0222 1.0000 0.3389
-8.000 -0.3246 0.09491 0.08822 -0.0207 1.0000 0.3514
-7.750 -0.3253 0.09254 0.08594 -0.0187 1.0000 0.3645
-7.500 -0.3299 0.09058 0.08408 -0.0161 1.0000 0.3797
-7.250 -0.3758 0.09166 0.08542 -0.0108 1.0000 0.3905
-7.000 -0.3383 0.08696 0.08065 -0.0096 1.0000 0.4135
-6.750 -0.3442 0.08530 0.07909 -0.0059 1.0000 0.4307
-6.500 -0.3512 0.08372 0.07761 -0.0020 1.0000 0.4479
-6.250 -0.3592 0.08218 0.07615 0.0020 1.0000 0.4650
-6.000 -0.3673 0.08066 0.07471 0.0062 1.0000 0.4828
-5.750 -0.3795 0.07944 0.07359 0.0109 1.0000 0.5027
-5.500 -0.5246 0.05871 0.05191 -0.0244 1.0000 0.1948
-5.250 -0.5133 0.05314 0.04563 -0.0262 1.0000 0.1760
-5.000 -0.4992 0.04993 0.04224 -0.0257 1.0000 0.1738
-4.750 -0.4826 0.04666 0.03863 -0.0258 1.0000 0.1701
-4.500 -0.4630 0.04329 0.03468 -0.0261 1.0000 0.1657
-4.250 -0.4428 0.04094 0.03189 -0.0260 1.0000 0.1659
-4.000 -0.4225 0.03916 0.02978 -0.0257 1.0000 0.1692
-3.750 -0.4008 0.03745 0.02766 -0.0255 1.0000 0.1718
-3.500 -0.3783 0.03593 0.02571 -0.0253 1.0000 0.1743
-3.250 -0.3564 0.03463 0.02417 -0.0250 1.0000 0.1796
-3.000 -0.3354 0.03373 0.02318 -0.0247 1.0000 0.1879
-2.750 -0.3130 0.03281 0.02204 -0.0244 1.0000 0.1959
-2.500 -0.2911 0.03216 0.02128 -0.0240 1.0000 0.2080
-2.250 -0.2689 0.03157 0.02070 -0.0237 1.0000 0.2256
-2.000 -0.2216 0.03098 0.02037 -0.0277 0.9921 0.2693
-1.750 -0.1728 0.02940 0.01995 -0.0317 0.9838 0.4259
-1.500 -0.1158 0.02812 0.02036 -0.0338 0.9737 1.0000
-1.250 -0.0743 0.02900 0.02072 -0.0376 0.9611 1.0000
-1.000 -0.0347 0.02986 0.02118 -0.0411 0.9484 1.0000
-0.750 0.0055 0.03075 0.02173 -0.0445 0.9358 1.0000
-0.500 0.0504 0.03168 0.02235 -0.0486 0.9235 1.0000
-0.250 0.0805 0.03239 0.02284 -0.0501 0.9097 1.0000
0.000 0.1104 0.03313 0.02339 -0.0515 0.8961 1.0000
0.250 0.1415 0.03392 0.02399 -0.0530 0.8828 1.0000
0.500 0.1753 0.03471 0.02462 -0.0548 0.8700 1.0000
0.750 0.2167 0.03548 0.02523 -0.0577 0.8580 1.0000
1.000 0.2418 0.03623 0.02587 -0.0581 0.8446 1.0000
1.250 0.2642 0.03705 0.02659 -0.0580 0.8313 1.0000
1.500 0.2902 0.03789 0.02734 -0.0585 0.8185 1.0000
1.750 0.3233 0.03867 0.02804 -0.0599 0.8070 1.0000
2.000 0.3552 0.03942 0.02873 -0.0610 0.7952 1.0000
2.250 0.3709 0.04045 0.02970 -0.0600 0.7822 1.0000
2.500 0.3916 0.04147 0.03068 -0.0597 0.7702 1.0000
2.750 0.4274 0.04220 0.03137 -0.0612 0.7599 1.0000
3.000 0.4454 0.04330 0.03245 -0.0606 0.7480 1.0000
3.250 0.4574 0.04465 0.03378 -0.0595 0.7362 1.0000
3.500 0.4854 0.04562 0.03474 -0.0600 0.7263 1.0000
3.750 0.5052 0.04683 0.03595 -0.0597 0.7159 1.0000
4.000 0.5128 0.04855 0.03767 -0.0584 0.7053 1.0000
4.250 0.5552 0.04907 0.03822 -0.0601 0.6973 1.0000
4.500 0.5468 0.05149 0.04063 -0.0576 0.6865 1.0000
4.750 0.5689 0.05289 0.04206 -0.0577 0.6784 1.0000
5.000 0.5774 0.05479 0.04399 -0.0568 0.6693 1.0000
5.250 0.5911 0.05663 0.04586 -0.0564 0.6614 1.0000
5.500 0.6002 0.05867 0.04793 -0.0558 0.6535 1.0000
5.750 0.6149 0.06055 0.04985 -0.0555 0.6457 1.0000
6.000 0.6221 0.06274 0.05208 -0.0548 0.6378 1.0000
6.250 0.6358 0.06466 0.05407 -0.0544 0.6291 1.0000
6.500 0.6727 0.06518 0.05469 -0.0547 0.6167 1.0000
6.750 0.6853 0.06665 0.05622 -0.0536 0.6031 1.0000
7.000 0.6947 0.06832 0.05796 -0.0524 0.5892 1.0000
7.250 0.7088 0.06979 0.05951 -0.0514 0.5750 1.0000
7.500 0.7251 0.07115 0.06096 -0.0504 0.5608 1.0000
7.750 0.7441 0.07228 0.06219 -0.0494 0.5458 1.0000
8.000 0.7641 0.07325 0.06329 -0.0483 0.5300 1.0000
8.250 0.7855 0.07398 0.06414 -0.0470 0.5131 1.0000
8.500 0.8099 0.07431 0.06461 -0.0456 0.4956 1.0000
8.750 0.8462 0.07320 0.06369 -0.0437 0.4772 1.0000
9.000 0.9007 0.06959 0.06031 -0.0410 0.4594 1.0000
9.250 1.0529 0.05346 0.04466 -0.0366 0.4404 1.0000
9.500 0.9252 0.07105 0.06201 -0.0365 0.4189 1.0000
9.750 1.1932 0.04321 0.03449 -0.0349 0.3741 1.0000
10.000 1.2228 0.04246 0.03358 -0.0331 0.3446 1.0000
10.250 1.2528 0.04220 0.03305 -0.0318 0.3171 1.0000
10.500 1.2545 0.04392 0.03487 -0.0286 0.2985 1.0000
10.750 1.2614 0.04545 0.03644 -0.0260 0.2812 1.0000
11.000 1.2708 0.04702 0.03802 -0.0238 0.2657 1.0000
11.250 1.2828 0.04870 0.03971 -0.0219 0.2522 1.0000
11.500 1.2984 0.05027 0.04125 -0.0204 0.2393 1.0000
11.750 1.3207 0.05185 0.04278 -0.0197 0.2267 1.0000
12.000 1.2958 0.05535 0.04666 -0.0153 0.2228 1.0000
12.250 1.3216 0.05701 0.04826 -0.0149 0.2116 1.0000
12.500 1.2906 0.06115 0.05275 -0.0111 0.2095 1.0000
12.750 1.2546 0.06636 0.05824 -0.0085 0.2082 1.0000
13.000 1.2041 0.07385 0.06598 -0.0077 0.2092 1.0000
13.250 1.1409 0.08476 0.07702 -0.0098 0.2116 1.0000
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Polar data table (+)
Polar graphs
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