GOE 367 AIRFOIL (goe367-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 367 AIRFOIL (goe367-il) Reynolds number: 500,000 Max Cl/Cd: 93.25 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe367-il-500000.txt Download as CSV file: xf-goe367-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 367 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.2459 0.10779 0.10530 -0.0769 0.9824 0.0472
-11.750 -0.5435 0.04235 0.03909 -0.1201 0.9487 0.0402
-11.500 -0.5852 0.03774 0.03405 -0.1154 0.9297 0.0402
-11.250 -0.6025 0.03422 0.03008 -0.1118 0.9178 0.0405
-11.000 -0.6046 0.03179 0.02727 -0.1090 0.9074 0.0408
-10.750 -0.5979 0.03015 0.02528 -0.1068 0.8990 0.0412
-10.500 -0.5908 0.02778 0.02259 -0.1048 0.8906 0.0415
-10.250 -0.5757 0.02604 0.02070 -0.1036 0.8840 0.0419
-10.000 -0.5564 0.02499 0.01956 -0.1027 0.8767 0.0423
-9.750 -0.5360 0.02411 0.01856 -0.1019 0.8704 0.0428
-9.500 -0.5148 0.02328 0.01763 -0.1011 0.8637 0.0432
-9.250 -0.4932 0.02246 0.01669 -0.1003 0.8564 0.0437
-9.000 -0.4714 0.02163 0.01568 -0.0995 0.8498 0.0443
-8.750 -0.4489 0.02083 0.01475 -0.0988 0.8427 0.0449
-8.500 -0.4257 0.02015 0.01390 -0.0982 0.8367 0.0457
-8.250 -0.4017 0.01960 0.01315 -0.0976 0.8314 0.0463
-8.000 -0.3776 0.01893 0.01234 -0.0971 0.8255 0.0468
-7.750 -0.3536 0.01773 0.01109 -0.0967 0.8201 0.0476
-7.500 -0.3281 0.01716 0.01045 -0.0964 0.8152 0.0483
-7.250 -0.3024 0.01667 0.00994 -0.0962 0.8100 0.0491
-7.000 -0.2765 0.01621 0.00942 -0.0959 0.8047 0.0499
-6.750 -0.2503 0.01578 0.00890 -0.0956 0.7998 0.0509
-6.500 -0.2237 0.01542 0.00845 -0.0954 0.7951 0.0518
-6.250 -0.1970 0.01517 0.00812 -0.0952 0.7900 0.0527
-6.000 -0.1721 0.01435 0.00731 -0.0949 0.7850 0.0541
-5.750 -0.1457 0.01398 0.00691 -0.0946 0.7804 0.0553
-5.500 -0.1191 0.01367 0.00657 -0.0944 0.7756 0.0566
-5.250 -0.0925 0.01336 0.00623 -0.0942 0.7705 0.0580
-5.000 -0.0655 0.01312 0.00592 -0.0940 0.7655 0.0592
-4.750 -0.0400 0.01266 0.00543 -0.0936 0.7608 0.0612
-4.500 -0.0137 0.01239 0.00520 -0.0934 0.7557 0.0633
-4.250 0.0132 0.01217 0.00495 -0.0932 0.7502 0.0656
-4.000 0.0397 0.01190 0.00463 -0.0929 0.7451 0.0680
-3.750 0.0661 0.01166 0.00442 -0.0926 0.7393 0.0711
-3.500 0.0928 0.01147 0.00421 -0.0924 0.7324 0.0745
-3.250 0.1196 0.01128 0.00401 -0.0921 0.7261 0.0796
-3.000 0.1463 0.01111 0.00388 -0.0919 0.7197 0.0864
-2.750 0.1734 0.01096 0.00377 -0.0917 0.7135 0.0963
-2.500 0.2009 0.01088 0.00368 -0.0916 0.7078 0.1092
-2.250 0.2284 0.01081 0.00365 -0.0916 0.7014 0.1204
-2.000 0.2555 0.01072 0.00358 -0.0914 0.6949 0.1303
-1.750 0.2830 0.01067 0.00349 -0.0914 0.6890 0.1377
-1.500 0.3100 0.01056 0.00343 -0.0912 0.6820 0.1453
-1.250 0.3373 0.01053 0.00335 -0.0911 0.6749 0.1513
-1.000 0.3639 0.01040 0.00326 -0.0909 0.6676 0.1595
-0.750 0.3908 0.01034 0.00318 -0.0907 0.6594 0.1663
-0.500 0.4172 0.01024 0.00310 -0.0904 0.6517 0.1765
-0.250 0.4436 0.01014 0.00304 -0.0902 0.6429 0.1905
0.000 0.4692 0.00999 0.00298 -0.0898 0.6347 0.2230
0.250 0.4881 0.00917 0.00295 -0.0886 0.6256 0.4692
0.500 0.4959 0.00811 0.00315 -0.0841 0.6175 0.8418
0.750 0.5172 0.00825 0.00334 -0.0822 0.6077 0.9028
1.000 0.5386 0.00846 0.00352 -0.0802 0.5983 0.9298
1.250 0.5614 0.00868 0.00369 -0.0786 0.5881 0.9491
1.500 0.5964 0.00894 0.00389 -0.0798 0.5776 0.9623
1.750 0.6436 0.00924 0.00408 -0.0837 0.5666 0.9705
2.000 0.6952 0.00949 0.00423 -0.0888 0.5546 0.9753
2.250 0.7318 0.00968 0.00435 -0.0908 0.5440 0.9811
2.500 0.7846 0.00990 0.00446 -0.0963 0.5319 0.9849
2.750 0.8287 0.01010 0.00456 -0.1000 0.5192 0.9897
3.000 0.8700 0.01025 0.00466 -0.1031 0.5072 0.9938
3.250 0.9158 0.01040 0.00472 -0.1072 0.4963 0.9975
3.500 0.9536 0.01053 0.00480 -0.1097 0.4864 1.0000
3.750 0.9743 0.01068 0.00490 -0.1085 0.4786 1.0000
4.000 0.9947 0.01081 0.00502 -0.1072 0.4703 1.0000
4.250 1.0138 0.01099 0.00514 -0.1057 0.4626 1.0000
4.500 1.0333 0.01112 0.00527 -0.1042 0.4545 1.0000
4.750 1.0503 0.01133 0.00542 -0.1023 0.4468 1.0000
5.000 1.0686 0.01146 0.00555 -0.1006 0.4391 1.0000
5.250 1.0835 0.01167 0.00571 -0.0983 0.4317 1.0000
5.500 1.0993 0.01182 0.00586 -0.0961 0.4246 1.0000
5.750 1.1124 0.01200 0.00603 -0.0934 0.4176 1.0000
6.000 1.1233 0.01220 0.00620 -0.0902 0.4115 1.0000
6.250 1.1343 0.01233 0.00634 -0.0871 0.4054 1.0000
6.500 1.1382 0.01250 0.00649 -0.0825 0.3994 1.0000
6.750 1.1381 0.01263 0.00662 -0.0771 0.3943 1.0000
7.000 1.1467 0.01278 0.00677 -0.0735 0.3887 1.0000
7.250 1.1596 0.01303 0.00701 -0.0709 0.3831 1.0000
7.500 1.1761 0.01333 0.00729 -0.0691 0.3782 1.0000
7.750 1.1958 0.01357 0.00757 -0.0680 0.3734 1.0000
8.000 1.2148 0.01388 0.00788 -0.0667 0.3685 1.0000
8.250 1.2324 0.01428 0.00825 -0.0653 0.3637 1.0000
8.500 1.2531 0.01458 0.00859 -0.0645 0.3595 1.0000
8.750 1.2730 0.01491 0.00895 -0.0635 0.3544 1.0000
9.000 1.2908 0.01535 0.00937 -0.0623 0.3491 1.0000
9.250 1.3096 0.01577 0.00981 -0.0613 0.3444 1.0000
9.500 1.3298 0.01613 0.01023 -0.0605 0.3396 1.0000
9.750 1.3473 0.01662 0.01072 -0.0593 0.3338 1.0000
10.000 1.3640 0.01716 0.01127 -0.0582 0.3276 1.0000
10.250 1.3820 0.01766 0.01181 -0.0572 0.3203 1.0000
10.500 1.3960 0.01837 0.01249 -0.0557 0.3128 1.0000
10.750 1.4134 0.01893 0.01310 -0.0548 0.3039 1.0000
11.000 1.4270 0.01972 0.01387 -0.0535 0.2957 1.0000
11.250 1.4425 0.02043 0.01462 -0.0524 0.2858 1.0000
11.500 1.4556 0.02130 0.01549 -0.0511 0.2741 1.0000
11.750 1.4650 0.02243 0.01657 -0.0495 0.2570 1.0000
12.000 1.4646 0.02425 0.01824 -0.0471 0.2245 1.0000
12.250 1.4462 0.02741 0.02110 -0.0433 0.1796 1.0000
12.500 1.4410 0.02981 0.02341 -0.0409 0.1620 1.0000
12.750 1.4400 0.03198 0.02556 -0.0390 0.1500 1.0000
13.000 1.4407 0.03410 0.02766 -0.0374 0.1377 1.0000
13.250 1.4416 0.03628 0.02981 -0.0360 0.1213 1.0000
13.750 1.4331 0.04178 0.03518 -0.0331 0.0947 1.0000
14.000 1.4315 0.04444 0.03786 -0.0320 0.0898 1.0000
14.250 1.4316 0.04701 0.04049 -0.0312 0.0869 1.0000
14.500 1.4288 0.04994 0.04346 -0.0304 0.0843 1.0000
14.750 1.4241 0.05316 0.04673 -0.0296 0.0818 1.0000
15.000 1.4219 0.05620 0.04984 -0.0291 0.0797 1.0000
15.250 1.4235 0.05888 0.05262 -0.0288 0.0778 1.0000
15.500 1.4216 0.06198 0.05580 -0.0285 0.0757 1.0000
15.750 1.4159 0.06556 0.05944 -0.0282 0.0736 1.0000
16.000 1.4061 0.06969 0.06363 -0.0280 0.0715 1.0000
16.250 1.4087 0.07241 0.06645 -0.0280 0.0695 1.0000
16.500 1.4109 0.07519 0.06931 -0.0281 0.0670 1.0000
16.750 1.4081 0.07861 0.07278 -0.0282 0.0647 1.0000
17.000 1.3989 0.08284 0.07705 -0.0284 0.0625 1.0000
17.250 1.4060 0.08509 0.07941 -0.0287 0.0595 1.0000
17.500 1.4053 0.08835 0.08271 -0.0290 0.0565 1.0000
17.750 1.4027 0.09185 0.08625 -0.0295 0.0539 1.0000
18.000 1.4035 0.09493 0.08939 -0.0299 0.0507 1.0000
18.250 1.3989 0.09875 0.09323 -0.0305 0.0483 1.0000
18.500 1.3984 0.10201 0.09655 -0.0310 0.0457 1.0000
18.750 1.3947 0.10578 0.10035 -0.0318 0.0436 1.0000
19.000 1.3894 0.10976 0.10436 -0.0326 0.0418 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 367 AIRFOIL (goe367-il)