BOEING 106 AIRFOIL (boe106-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: BOEING 106 AIRFOIL (boe106-il) Reynolds number: 100,000 Max Cl/Cd: 51.64 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-boe106-il-100000.txt Download as CSV file: xf-boe106-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3586 0.09735 0.09240 -0.0388 1.0000 0.1399
-9.000 -0.4047 0.09539 0.09067 -0.0433 1.0000 0.1438
-8.750 -0.4490 0.09404 0.08951 -0.0422 1.0000 0.1441
-8.500 -0.3882 0.08858 0.08393 -0.0386 1.0000 0.1485
-8.250 -0.3932 0.08675 0.08218 -0.0359 1.0000 0.1516
-8.000 -0.4166 0.08548 0.08103 -0.0325 1.0000 0.1538
-7.750 -0.4500 0.08452 0.08020 -0.0284 1.0000 0.1548
-7.500 -0.4836 0.08274 0.07852 -0.0265 1.0000 0.1570
-7.000 -0.5386 0.07658 0.07239 -0.0255 1.0000 0.1627
-6.750 -0.5323 0.07464 0.07052 -0.0222 1.0000 0.1654
-6.500 -0.5337 0.06966 0.06529 -0.0287 0.9948 0.1784
-6.250 -0.5019 0.06717 0.06289 -0.0281 0.9904 0.1853
-6.000 -0.4909 0.04789 0.04193 -0.0423 0.9810 0.1113
-5.750 -0.4669 0.04094 0.03387 -0.0440 0.9740 0.0983
-5.500 -0.4314 0.03774 0.03011 -0.0466 0.9685 0.0986
-5.250 -0.4020 0.03500 0.02697 -0.0476 0.9611 0.0984
-5.000 -0.3622 0.03260 0.02406 -0.0503 0.9560 0.0995
-4.750 -0.3328 0.03080 0.02220 -0.0514 0.9481 0.1029
-4.500 -0.2927 0.02936 0.02053 -0.0539 0.9422 0.1060
-4.250 -0.2579 0.02815 0.01905 -0.0552 0.9349 0.1093
-4.000 -0.2201 0.02703 0.01768 -0.0570 0.9278 0.1150
-3.750 -0.1735 0.02592 0.01659 -0.0605 0.9236 0.1228
-3.500 -0.1473 0.02520 0.01575 -0.0602 0.9133 0.1298
-3.250 -0.1020 0.02417 0.01481 -0.0634 0.9085 0.1460
-3.000 -0.0762 0.02336 0.01419 -0.0631 0.8986 0.1645
-2.750 -0.0354 0.02202 0.01329 -0.0656 0.8934 0.2254
-2.500 -0.0119 0.02081 0.01299 -0.0651 0.8840 0.3692
-2.250 0.0152 0.01935 0.01311 -0.0633 0.8782 0.7254
-2.000 0.0776 0.01933 0.01335 -0.0662 0.8762 0.9082
-1.750 0.2008 0.01934 0.01301 -0.0823 0.8776 0.9750
-1.500 0.3038 0.01864 0.01201 -0.0964 0.8761 0.9995
-1.250 0.3239 0.01854 0.01179 -0.0953 0.8627 1.0000
-1.000 0.3441 0.01844 0.01157 -0.0941 0.8501 1.0000
-0.750 0.3701 0.01820 0.01119 -0.0935 0.8402 1.0000
-0.500 0.3914 0.01808 0.01097 -0.0922 0.8284 1.0000
-0.250 0.4094 0.01813 0.01092 -0.0905 0.8157 1.0000
0.000 0.4312 0.01807 0.01077 -0.0892 0.8050 1.0000
0.250 0.4559 0.01792 0.01050 -0.0882 0.7954 1.0000
0.500 0.4741 0.01805 0.01056 -0.0864 0.7829 1.0000
0.750 0.4962 0.01806 0.01050 -0.0851 0.7725 1.0000
1.000 0.5215 0.01795 0.01029 -0.0842 0.7631 1.0000
1.250 0.5406 0.01814 0.01042 -0.0825 0.7510 1.0000
1.500 0.5639 0.01817 0.01038 -0.0814 0.7408 1.0000
1.750 0.5891 0.01812 0.01026 -0.0805 0.7312 1.0000
2.000 0.6089 0.01835 0.01045 -0.0789 0.7193 1.0000
2.250 0.6336 0.01839 0.01043 -0.0780 0.7097 1.0000
2.500 0.6582 0.01843 0.01041 -0.0770 0.6993 1.0000
2.750 0.6787 0.01868 0.01064 -0.0756 0.6876 1.0000
3.000 0.7042 0.01875 0.01065 -0.0748 0.6781 1.0000
3.250 0.7281 0.01889 0.01076 -0.0739 0.6676 1.0000
3.500 0.7490 0.01918 0.01106 -0.0725 0.6562 1.0000
3.750 0.7754 0.01926 0.01108 -0.0719 0.6469 1.0000
4.000 0.7986 0.01948 0.01130 -0.0709 0.6360 1.0000
4.250 0.8196 0.01986 0.01170 -0.0696 0.6252 1.0000
4.500 0.8489 0.01992 0.01168 -0.0695 0.6171 1.0000
4.750 0.8686 0.02033 0.01215 -0.0680 0.6052 1.0000
5.000 0.8908 0.02061 0.01245 -0.0669 0.5938 1.0000
5.250 0.9175 0.02068 0.01246 -0.0662 0.5831 1.0000
5.500 0.9426 0.02076 0.01251 -0.0654 0.5710 1.0000
5.750 0.9636 0.02102 0.01281 -0.0640 0.5580 1.0000
6.000 0.9861 0.02125 0.01305 -0.0628 0.5457 1.0000
6.250 1.0111 0.02140 0.01319 -0.0620 0.5343 1.0000
6.500 1.0361 0.02159 0.01336 -0.0613 0.5231 1.0000
6.750 1.0559 0.02196 0.01382 -0.0598 0.5108 1.0000
7.000 1.0780 0.02223 0.01413 -0.0587 0.4985 1.0000
7.250 1.1012 0.02239 0.01430 -0.0576 0.4857 1.0000
7.500 1.1245 0.02251 0.01440 -0.0566 0.4722 1.0000
7.750 1.1466 0.02263 0.01451 -0.0553 0.4579 1.0000
8.000 1.1654 0.02282 0.01478 -0.0536 0.4427 1.0000
8.250 1.1839 0.02301 0.01502 -0.0519 0.4268 1.0000
8.500 1.2001 0.02324 0.01531 -0.0498 0.4094 1.0000
8.750 1.2135 0.02353 0.01569 -0.0473 0.3897 1.0000
9.000 1.2260 0.02383 0.01602 -0.0447 0.3676 1.0000
9.250 1.2348 0.02430 0.01647 -0.0416 0.3420 1.0000
9.500 1.2401 0.02498 0.01705 -0.0381 0.3134 1.0000
9.750 1.2419 0.02589 0.01783 -0.0343 0.2841 1.0000
10.000 1.2411 0.02697 0.01878 -0.0303 0.2579 1.0000
10.250 1.2421 0.02825 0.01994 -0.0269 0.2349 1.0000
10.500 1.2448 0.02965 0.02120 -0.0240 0.2166 1.0000
10.750 1.2502 0.03113 0.02254 -0.0216 0.2014 1.0000
11.000 1.2589 0.03264 0.02396 -0.0197 0.1883 1.0000
11.250 1.2712 0.03417 0.02541 -0.0182 0.1770 1.0000
11.500 1.2884 0.03570 0.02681 -0.0172 0.1667 1.0000
11.750 1.3051 0.03718 0.02829 -0.0162 0.1577 1.0000
12.000 1.3221 0.03887 0.03006 -0.0152 0.1500 1.0000
12.250 1.3542 0.04056 0.03151 -0.0160 0.1406 1.0000
12.500 1.3514 0.04220 0.03351 -0.0132 0.1359 1.0000
12.750 1.3635 0.04367 0.03496 -0.0120 0.1287 1.0000
13.000 1.3673 0.04559 0.03705 -0.0103 0.1229 1.0000
13.250 1.3677 0.04732 0.03892 -0.0084 0.1175 1.0000
13.500 1.3802 0.04916 0.04072 -0.0075 0.1113 1.0000
13.750 1.3729 0.05142 0.04327 -0.0054 0.1072 1.0000
14.000 1.3822 0.05302 0.04481 -0.0045 0.1015 1.0000
14.250 1.3780 0.05575 0.04777 -0.0030 0.0973 1.0000
14.500 1.3686 0.05856 0.05085 -0.0016 0.0935 1.0000
14.750 1.3835 0.06031 0.05240 -0.0010 0.0873 1.0000
15.000 1.3641 0.06400 0.05648 0.0002 0.0850 1.0000
15.250 1.3503 0.06767 0.06042 0.0008 0.0819 1.0000
15.500 1.3572 0.06965 0.06231 0.0013 0.0773 1.0000
15.750 1.3442 0.07394 0.06682 0.0015 0.0748 1.0000
16.000 1.3211 0.07907 0.07229 0.0011 0.0734 1.0000
16.250 1.2992 0.08447 0.07797 0.0001 0.0719 1.0000
16.500 1.2807 0.08981 0.08353 -0.0013 0.0703 1.0000
16.750 1.2647 0.09505 0.08894 -0.0029 0.0687 1.0000
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Polar data table (+)
Polar graphs
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