GOE 681 AIRFOIL (goe681-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 681 AIRFOIL (goe681-il) Reynolds number: 200,000 Max Cl/Cd: 55.42 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe681-il-200000-n5.txt Download as CSV file: xf-goe681-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 681 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.750 -0.5842 0.07305 0.06808 -0.0801 1.0000 0.0369
-13.500 -0.6327 0.06290 0.05770 -0.0870 1.0000 0.0368
-13.250 -0.6646 0.05620 0.05083 -0.0912 1.0000 0.0367
-13.000 -0.6933 0.05109 0.04558 -0.0936 1.0000 0.0367
-12.750 -0.7148 0.04691 0.04122 -0.0958 0.9984 0.0367
-12.500 -0.7128 0.04259 0.03661 -0.1013 0.9887 0.0368
-12.250 -0.7061 0.03963 0.03340 -0.1043 0.9790 0.0369
-12.000 -0.6963 0.03731 0.03083 -0.1058 0.9692 0.0370
-11.750 -0.6836 0.03552 0.02893 -0.1059 0.9583 0.0372
-11.500 -0.6651 0.03390 0.02722 -0.1064 0.9495 0.0374
-11.250 -0.6416 0.03236 0.02560 -0.1076 0.9417 0.0377
-11.000 -0.6188 0.03096 0.02411 -0.1084 0.9332 0.0379
-10.750 -0.5935 0.02964 0.02269 -0.1095 0.9252 0.0382
-10.500 -0.5693 0.02844 0.02139 -0.1101 0.9166 0.0385
-10.250 -0.5434 0.02728 0.02012 -0.1108 0.9079 0.0389
-10.000 -0.5186 0.02623 0.01897 -0.1112 0.8989 0.0393
-9.750 -0.4933 0.02525 0.01788 -0.1116 0.8897 0.0396
-9.500 -0.4690 0.02436 0.01687 -0.1116 0.8805 0.0400
-9.250 -0.4454 0.02357 0.01595 -0.1113 0.8708 0.0406
-9.000 -0.4220 0.02285 0.01511 -0.1109 0.8620 0.0411
-8.750 -0.4006 0.02215 0.01432 -0.1101 0.8525 0.0416
-8.500 -0.3783 0.02146 0.01359 -0.1094 0.8443 0.0422
-8.250 -0.3585 0.02090 0.01299 -0.1082 0.8349 0.0427
-8.000 -0.3355 0.02034 0.01234 -0.1075 0.8275 0.0433
-7.750 -0.3156 0.01985 0.01181 -0.1062 0.8190 0.0439
-7.500 -0.2938 0.01937 0.01125 -0.1051 0.8109 0.0445
-7.250 -0.2711 0.01892 0.01070 -0.1042 0.8040 0.0452
-7.000 -0.2503 0.01853 0.01025 -0.1029 0.7960 0.0459
-6.750 -0.2277 0.01810 0.00976 -0.1020 0.7888 0.0468
-6.500 -0.2053 0.01771 0.00932 -0.1009 0.7819 0.0478
-6.250 -0.1833 0.01737 0.00894 -0.0998 0.7744 0.0491
-6.000 -0.1590 0.01704 0.00853 -0.0990 0.7679 0.0507
-5.750 -0.1359 0.01671 0.00818 -0.0981 0.7612 0.0528
-5.500 -0.1128 0.01641 0.00785 -0.0971 0.7538 0.0562
-5.250 -0.0882 0.01610 0.00752 -0.0964 0.7476 0.0608
-5.000 -0.0643 0.01580 0.00722 -0.0955 0.7413 0.0681
-4.750 -0.0409 0.01550 0.00693 -0.0945 0.7340 0.0785
-4.500 -0.0161 0.01519 0.00661 -0.0938 0.7275 0.0912
-4.250 0.0074 0.01491 0.00637 -0.0929 0.7198 0.1052
-4.000 0.0310 0.01465 0.00616 -0.0920 0.7109 0.1229
-3.750 0.0554 0.01445 0.00602 -0.0911 0.7018 0.1443
-3.500 0.0797 0.01434 0.00592 -0.0902 0.6913 0.1654
-3.250 0.1049 0.01427 0.00582 -0.0894 0.6817 0.1816
-3.000 0.1296 0.01421 0.00573 -0.0886 0.6720 0.1937
-2.750 0.1551 0.01415 0.00563 -0.0879 0.6646 0.2033
-2.500 0.1800 0.01411 0.00556 -0.0871 0.6561 0.2133
-2.250 0.2054 0.01404 0.00544 -0.0864 0.6483 0.2219
-2.000 0.2299 0.01399 0.00539 -0.0855 0.6399 0.2315
-1.750 0.2548 0.01393 0.00530 -0.0848 0.6314 0.2409
-1.500 0.2792 0.01387 0.00524 -0.0839 0.6229 0.2498
-1.250 0.3040 0.01384 0.00515 -0.0831 0.6138 0.2576
-1.000 0.3279 0.01375 0.00508 -0.0821 0.6050 0.2667
-0.750 0.3519 0.01371 0.00502 -0.0812 0.5952 0.2769
-0.500 0.3756 0.01365 0.00497 -0.0802 0.5859 0.2867
-0.250 0.3990 0.01361 0.00492 -0.0791 0.5756 0.2978
0.000 0.4221 0.01355 0.00488 -0.0780 0.5656 0.3094
0.750 0.4883 0.01340 0.00483 -0.0742 0.5322 0.3733
1.000 0.5084 0.01329 0.00487 -0.0726 0.5194 0.4221
1.250 0.5273 0.01316 0.00491 -0.0707 0.5067 0.4868
1.500 0.5442 0.01299 0.00496 -0.0684 0.4938 0.5649
1.750 0.5591 0.01272 0.00509 -0.0655 0.4803 0.6885
2.000 0.5956 0.01267 0.00538 -0.0667 0.4631 0.8416
2.250 0.6612 0.01305 0.00569 -0.0743 0.4391 0.9235
2.500 0.7041 0.01340 0.00589 -0.0774 0.4181 0.9554
2.750 0.7405 0.01376 0.00610 -0.0793 0.3991 0.9748
3.000 0.7741 0.01413 0.00632 -0.0806 0.3831 0.9902
3.500 0.8185 0.01477 0.00674 -0.0789 0.3597 1.0000
3.750 0.8297 0.01507 0.00693 -0.0758 0.3514 1.0000
4.000 0.8438 0.01531 0.00711 -0.0731 0.3437 1.0000
4.250 0.8554 0.01561 0.00733 -0.0701 0.3363 1.0000
4.500 0.8666 0.01588 0.00753 -0.0670 0.3302 1.0000
4.750 0.8786 0.01614 0.00775 -0.0640 0.3239 1.0000
5.000 0.8897 0.01646 0.00799 -0.0609 0.3184 1.0000
5.250 0.9025 0.01678 0.00826 -0.0582 0.3135 1.0000
5.500 0.9178 0.01706 0.00852 -0.0560 0.3088 1.0000
5.750 0.9326 0.01740 0.00882 -0.0537 0.3043 1.0000
6.000 0.9469 0.01778 0.00915 -0.0514 0.3002 1.0000
6.250 0.9621 0.01817 0.00949 -0.0494 0.2965 1.0000
6.500 0.9796 0.01850 0.00983 -0.0477 0.2928 1.0000
6.750 0.9965 0.01887 0.01020 -0.0460 0.2891 1.0000
7.000 1.0126 0.01928 0.01058 -0.0442 0.2855 1.0000
7.250 1.0282 0.01974 0.01100 -0.0424 0.2821 1.0000
7.500 1.0447 0.02021 0.01143 -0.0408 0.2790 1.0000
7.750 1.0626 0.02060 0.01186 -0.0395 0.2759 1.0000
8.000 1.0802 0.02103 0.01231 -0.0381 0.2728 1.0000
8.250 1.0972 0.02150 0.01278 -0.0367 0.2700 1.0000
8.500 1.1140 0.02199 0.01327 -0.0353 0.2673 1.0000
8.750 1.1308 0.02253 0.01378 -0.0339 0.2649 1.0000
9.000 1.1483 0.02309 0.01429 -0.0327 0.2626 1.0000
9.250 1.1660 0.02358 0.01485 -0.0316 0.2604 1.0000
9.500 1.1829 0.02410 0.01543 -0.0303 0.2578 1.0000
9.750 1.1996 0.02466 0.01603 -0.0292 0.2553 1.0000
10.000 1.2161 0.02524 0.01665 -0.0280 0.2529 1.0000
10.250 1.2323 0.02585 0.01727 -0.0268 0.2508 1.0000
10.500 1.2484 0.02649 0.01790 -0.0257 0.2485 1.0000
10.750 1.2655 0.02713 0.01851 -0.0247 0.2463 1.0000
11.000 1.2805 0.02782 0.01927 -0.0235 0.2440 1.0000
11.250 1.2949 0.02854 0.02008 -0.0224 0.2415 1.0000
11.500 1.3080 0.02932 0.02093 -0.0211 0.2387 1.0000
11.750 1.3211 0.03012 0.02178 -0.0200 0.2360 1.0000
12.000 1.3338 0.03095 0.02264 -0.0188 0.2333 1.0000
12.250 1.3478 0.03177 0.02344 -0.0178 0.2310 1.0000
12.500 1.3621 0.03260 0.02430 -0.0169 0.2288 1.0000
12.750 1.3734 0.03359 0.02541 -0.0158 0.2266 1.0000
13.000 1.3843 0.03462 0.02655 -0.0148 0.2241 1.0000
13.250 1.3946 0.03571 0.02772 -0.0138 0.2215 1.0000
13.500 1.4048 0.03681 0.02887 -0.0128 0.2190 1.0000
13.750 1.4149 0.03792 0.03001 -0.0119 0.2165 1.0000
14.000 1.4265 0.03898 0.03104 -0.0110 0.2140 1.0000
14.250 1.4327 0.04047 0.03270 -0.0102 0.2113 1.0000
14.500 1.4384 0.04204 0.03439 -0.0093 0.2081 1.0000
14.750 1.4445 0.04362 0.03606 -0.0086 0.2051 1.0000
15.000 1.4501 0.04524 0.03771 -0.0079 0.2020 1.0000
15.250 1.4577 0.04673 0.03919 -0.0073 0.1994 1.0000
15.500 1.4613 0.04872 0.04136 -0.0068 0.1965 1.0000
15.750 1.4642 0.05082 0.04360 -0.0064 0.1932 1.0000
16.000 1.4679 0.05286 0.04574 -0.0061 0.1904 1.0000
16.250 1.4712 0.05497 0.04789 -0.0058 0.1875 1.0000
16.500 1.4744 0.05713 0.05010 -0.0056 0.1849 1.0000
16.750 1.4749 0.05973 0.05287 -0.0056 0.1821 1.0000
17.000 1.4740 0.06252 0.05580 -0.0057 0.1787 1.0000
17.250 1.4729 0.06535 0.05873 -0.0059 0.1756 1.0000
17.500 1.4712 0.06827 0.06167 -0.0061 0.1724 1.0000
17.750 1.4658 0.07184 0.06543 -0.0066 0.1691 1.0000
18.000 1.4570 0.07587 0.06960 -0.0073 0.1648 1.0000
18.250 1.4484 0.07991 0.07370 -0.0082 0.1605 1.0000
18.500 1.4347 0.08483 0.07879 -0.0093 0.1562 1.0000
18.750 1.4199 0.08996 0.08406 -0.0107 0.1524 1.0000
19.000 1.4045 0.09522 0.08939 -0.0122 0.1484 1.0000
19.250 1.3774 0.10244 0.09681 -0.0146 0.1442 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 681 AIRFOIL (goe681-il)