OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: OAF117 AIRFOIL (oaf117-il) Reynolds number: 100,000 Max Cl/Cd: 46.95 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf117-il-100000.txt Download as CSV file: xf-oaf117-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: OAF117 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.7083 0.06555 0.06037 -0.0283 1.0000 0.0930
-8.750 -0.7122 0.05682 0.05132 -0.0349 1.0000 0.0914
-8.500 -0.7107 0.04819 0.04209 -0.0413 1.0000 0.0913
-8.250 -0.6978 0.04138 0.03441 -0.0459 1.0000 0.0927
-8.000 -0.6747 0.03884 0.03194 -0.0464 1.0000 0.0953
-7.750 -0.6502 0.03610 0.02896 -0.0476 1.0000 0.0980
-7.500 -0.6242 0.03306 0.02540 -0.0492 1.0000 0.1022
-7.250 -0.5968 0.03009 0.02192 -0.0507 1.0000 0.1061
-7.000 -0.5693 0.02859 0.02043 -0.0511 1.0000 0.1106
-6.750 -0.5399 0.02690 0.01831 -0.0520 1.0000 0.1170
-6.500 -0.5117 0.02521 0.01666 -0.0525 1.0000 0.1228
-6.250 -0.4818 0.02398 0.01513 -0.0531 1.0000 0.1314
-6.000 -0.4532 0.02267 0.01394 -0.0535 1.0000 0.1399
-5.750 -0.4237 0.02152 0.01273 -0.0540 1.0000 0.1511
-5.500 -0.3940 0.02057 0.01178 -0.0544 1.0000 0.1650
-5.250 -0.3646 0.01974 0.01101 -0.0549 1.0000 0.1822
-5.000 -0.3350 0.01899 0.01034 -0.0553 1.0000 0.2029
-4.750 -0.3050 0.01837 0.00971 -0.0559 1.0000 0.2263
-4.500 -0.2759 0.01789 0.00940 -0.0562 1.0000 0.2480
-4.250 -0.2463 0.01750 0.00907 -0.0566 1.0000 0.2706
-4.000 -0.2169 0.01714 0.00876 -0.0571 1.0000 0.2930
-3.750 -0.1883 0.01684 0.00862 -0.0574 1.0000 0.3143
-3.500 -0.1603 0.01661 0.00849 -0.0577 1.0000 0.3370
-3.250 -0.1344 0.01643 0.00836 -0.0578 1.0000 0.3601
-3.000 -0.1161 0.01640 0.00849 -0.0570 1.0000 0.3789
-2.750 -0.0933 0.01648 0.00870 -0.0573 0.9968 0.3990
-2.500 -0.0374 0.01612 0.00839 -0.0629 0.9829 0.4312
-2.250 0.0136 0.01581 0.00819 -0.0673 0.9640 0.4613
-2.000 0.0587 0.01558 0.00804 -0.0703 0.9421 0.4890
-1.750 0.0928 0.01553 0.00800 -0.0709 0.9153 0.5156
-1.500 0.1190 0.01554 0.00805 -0.0696 0.8901 0.5384
-1.250 0.1425 0.01555 0.00810 -0.0678 0.8641 0.5622
-1.000 0.1653 0.01556 0.00811 -0.0658 0.8413 0.5882
-0.750 0.1884 0.01553 0.00812 -0.0639 0.8191 0.6150
-0.500 0.2124 0.01547 0.00811 -0.0624 0.7980 0.6431
-0.250 0.2371 0.01541 0.00809 -0.0610 0.7787 0.6734
0.000 0.2614 0.01533 0.00807 -0.0595 0.7605 0.7070
0.250 0.2847 0.01522 0.00805 -0.0577 0.7431 0.7460
0.500 0.3058 0.01504 0.00801 -0.0554 0.7268 0.7984
0.750 0.3209 0.01462 0.00779 -0.0514 0.7121 0.8892
1.000 0.3558 0.01454 0.00756 -0.0528 0.6949 1.0000
1.250 0.3876 0.01481 0.00760 -0.0536 0.6788 1.0000
1.500 0.4178 0.01508 0.00769 -0.0538 0.6633 1.0000
1.750 0.4474 0.01535 0.00780 -0.0540 0.6477 1.0000
2.000 0.4770 0.01562 0.00796 -0.0541 0.6315 1.0000
2.250 0.5063 0.01590 0.00814 -0.0541 0.6155 1.0000
2.500 0.5354 0.01618 0.00833 -0.0541 0.5994 1.0000
2.750 0.5644 0.01647 0.00853 -0.0541 0.5835 1.0000
3.000 0.5932 0.01676 0.00875 -0.0540 0.5679 1.0000
3.250 0.6220 0.01706 0.00897 -0.0539 0.5527 1.0000
3.500 0.6507 0.01737 0.00921 -0.0537 0.5378 1.0000
3.750 0.6792 0.01767 0.00944 -0.0536 0.5231 1.0000
4.000 0.7076 0.01798 0.00965 -0.0533 0.5088 1.0000
4.250 0.7362 0.01829 0.00994 -0.0532 0.4931 1.0000
4.500 0.7646 0.01862 0.01027 -0.0531 0.4773 1.0000
4.750 0.7929 0.01895 0.01059 -0.0530 0.4615 1.0000
5.000 0.8210 0.01928 0.01090 -0.0528 0.4454 1.0000
5.250 0.8490 0.01962 0.01123 -0.0526 0.4289 1.0000
5.500 0.8767 0.01994 0.01153 -0.0524 0.4119 1.0000
5.750 0.9042 0.02026 0.01182 -0.0521 0.3943 1.0000
6.000 0.9315 0.02058 0.01207 -0.0518 0.3760 1.0000
6.250 0.9585 0.02093 0.01237 -0.0515 0.3567 1.0000
6.500 0.9851 0.02129 0.01280 -0.0512 0.3355 1.0000
6.750 1.0115 0.02167 0.01315 -0.0509 0.3146 1.0000
7.000 1.0374 0.02210 0.01349 -0.0505 0.2944 1.0000
7.250 1.0629 0.02264 0.01404 -0.0502 0.2728 1.0000
7.500 1.0878 0.02326 0.01465 -0.0499 0.2507 1.0000
8.000 1.1355 0.02476 0.01608 -0.0491 0.2058 1.0000
8.250 1.1576 0.02584 0.01697 -0.0486 0.1847 1.0000
8.500 1.1790 0.02714 0.01828 -0.0479 0.1637 1.0000
8.750 1.1999 0.02854 0.01962 -0.0473 0.1466 1.0000
9.000 1.2205 0.03004 0.02110 -0.0465 0.1331 1.0000
9.250 1.2410 0.03162 0.02265 -0.0458 0.1226 1.0000
9.500 1.2619 0.03317 0.02405 -0.0451 0.1140 1.0000
9.750 1.2815 0.03489 0.02603 -0.0442 0.1067 1.0000
10.000 1.3032 0.03675 0.02765 -0.0437 0.1007 1.0000
10.250 1.3203 0.03861 0.02992 -0.0426 0.0958 1.0000
10.500 1.3396 0.04039 0.03175 -0.0418 0.0916 1.0000
10.750 1.3586 0.04284 0.03420 -0.0412 0.0880 1.0000
11.000 1.3702 0.04518 0.03696 -0.0398 0.0848 1.0000
11.250 1.3832 0.04755 0.03956 -0.0387 0.0822 1.0000
11.500 1.3973 0.04991 0.04202 -0.0377 0.0799 1.0000
11.750 1.4123 0.05330 0.04542 -0.0372 0.0778 1.0000
12.000 1.4065 0.05655 0.04918 -0.0351 0.0768 1.0000
12.250 1.3954 0.06014 0.05319 -0.0330 0.0759 1.0000
12.500 1.3768 0.06380 0.05719 -0.0306 0.0754 1.0000
12.750 1.3537 0.06819 0.06188 -0.0296 0.0751 1.0000
13.000 1.3269 0.07371 0.06770 -0.0305 0.0752 1.0000
13.250 1.2952 0.08063 0.07491 -0.0332 0.0755 1.0000
13.500 1.2587 0.08922 0.08375 -0.0378 0.0761 1.0000
13.750 1.2187 0.09952 0.09427 -0.0441 0.0769 1.0000
14.000 1.1802 0.11088 0.10578 -0.0509 0.0778 1.0000
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Polar data table (+)
Polar graphs
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