GOE 546 AIRFOIL (goe546-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: GOE 546 AIRFOIL (goe546-il) Reynolds number: 100,000 Max Cl/Cd: 59.41 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe546-il-100000.txt Download as CSV file: xf-goe546-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 546 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3718 0.09672 0.09170 -0.0366 1.0000 0.1168
-8.500 -0.3911 0.09495 0.09006 -0.0377 1.0000 0.1208
-8.250 -0.4247 0.09413 0.08945 -0.0374 1.0000 0.1218
-8.000 -0.4610 0.09258 0.08805 -0.0389 1.0000 0.1223
-7.750 -0.4068 0.08753 0.08289 -0.0319 1.0000 0.1282
-7.500 -0.4215 0.08611 0.08156 -0.0289 1.0000 0.1308
-7.250 -0.4424 0.08450 0.08005 -0.0272 1.0000 0.1338
-7.000 -0.4919 0.08105 0.07658 -0.0378 1.0000 0.1376
-6.750 -0.4754 0.07820 0.07389 -0.0294 1.0000 0.1400
-6.500 -0.4683 0.07682 0.07254 -0.0247 1.0000 0.1453
-6.250 -0.4867 0.07160 0.06722 -0.0331 1.0000 0.1541
-6.000 -0.4763 0.07071 0.06646 -0.0257 1.0000 0.1585
-5.750 -0.4792 0.06652 0.06219 -0.0300 1.0000 0.1705
-5.500 -0.4719 0.06506 0.06079 -0.0258 1.0000 0.1759
-5.250 -0.4674 0.06169 0.05736 -0.0274 1.0000 0.1880
-5.000 -0.4581 0.05927 0.05460 -0.0327 1.0000 0.2144
-4.750 -0.4511 0.05651 0.05207 -0.0279 1.0000 0.2173
-4.500 -0.3918 0.03738 0.03036 -0.0458 1.0000 0.1082
-4.250 -0.3654 0.03409 0.02697 -0.0465 0.9985 0.1026
-4.000 -0.3244 0.03096 0.02318 -0.0496 0.9937 0.0988
-3.750 -0.2851 0.02899 0.02074 -0.0522 0.9876 0.0994
-3.500 -0.2448 0.02732 0.01861 -0.0546 0.9820 0.0993
-3.250 -0.2065 0.02622 0.01720 -0.0567 0.9752 0.1022
-3.000 -0.1676 0.02540 0.01608 -0.0588 0.9685 0.1061
-2.750 -0.1293 0.02432 0.01490 -0.0608 0.9617 0.1091
-2.500 -0.0932 0.02352 0.01417 -0.0626 0.9546 0.1155
-2.250 -0.0546 0.02287 0.01353 -0.0647 0.9474 0.1255
-2.000 -0.0205 0.02235 0.01306 -0.0660 0.9391 0.1377
-1.750 0.0190 0.02151 0.01250 -0.0684 0.9323 0.1726
-1.500 0.0443 0.01986 0.01262 -0.0681 0.9241 0.5481
-1.250 0.0788 0.01936 0.01269 -0.0684 0.9175 0.7097
-1.000 0.1455 0.01881 0.01253 -0.0751 0.9136 1.0000
-0.750 0.1801 0.01898 0.01244 -0.0767 0.9039 1.0000
-0.500 0.2172 0.01915 0.01239 -0.0787 0.8956 1.0000
-0.250 0.2563 0.01925 0.01232 -0.0810 0.8877 1.0000
0.000 0.2894 0.01943 0.01236 -0.0821 0.8786 1.0000
0.250 0.3314 0.01943 0.01224 -0.0847 0.8718 1.0000
0.500 0.3621 0.01961 0.01232 -0.0853 0.8622 1.0000
0.750 0.4071 0.01950 0.01213 -0.0883 0.8562 1.0000
1.000 0.4361 0.01968 0.01226 -0.0886 0.8463 1.0000
1.250 0.4794 0.01952 0.01206 -0.0911 0.8402 1.0000
1.500 0.5054 0.01973 0.01224 -0.0907 0.8295 1.0000
1.750 0.5489 0.01948 0.01197 -0.0930 0.8238 1.0000
2.000 0.5729 0.01971 0.01219 -0.0922 0.8123 1.0000
2.250 0.6183 0.01928 0.01177 -0.0945 0.8069 1.0000
2.500 0.6438 0.01927 0.01177 -0.0936 0.7939 1.0000
2.750 0.6740 0.01901 0.01153 -0.0931 0.7811 1.0000
3.000 0.7058 0.01869 0.01122 -0.0929 0.7687 1.0000
3.250 0.7433 0.01827 0.01082 -0.0936 0.7595 1.0000
3.500 0.7705 0.01824 0.01084 -0.0930 0.7473 1.0000
3.750 0.7966 0.01826 0.01090 -0.0921 0.7343 1.0000
4.000 0.8241 0.01825 0.01095 -0.0915 0.7217 1.0000
4.250 0.8530 0.01819 0.01095 -0.0911 0.7092 1.0000
4.500 0.8819 0.01802 0.01082 -0.0904 0.6940 1.0000
4.750 0.9103 0.01776 0.01057 -0.0895 0.6757 1.0000
5.000 0.9413 0.01735 0.01010 -0.0887 0.6542 1.0000
5.250 0.9628 0.01726 0.01000 -0.0866 0.6275 1.0000
5.500 0.9850 0.01724 0.00994 -0.0847 0.6004 1.0000
5.750 1.0062 0.01729 0.00996 -0.0827 0.5730 1.0000
6.000 1.0252 0.01740 0.01008 -0.0805 0.5435 1.0000
6.250 1.0425 0.01757 0.01025 -0.0781 0.5112 1.0000
6.500 1.0586 0.01782 0.01046 -0.0755 0.4747 1.0000
6.750 1.0723 0.01822 0.01074 -0.0725 0.4309 1.0000
7.000 1.0819 0.01888 0.01119 -0.0691 0.3728 1.0000
7.250 1.0804 0.02019 0.01191 -0.0640 0.2730 1.0000
7.500 1.0691 0.02264 0.01340 -0.0582 0.1585 1.0000
7.750 1.0682 0.02451 0.01483 -0.0538 0.1255 1.0000
8.000 1.0721 0.02604 0.01620 -0.0502 0.1125 1.0000
8.250 1.0814 0.02735 0.01750 -0.0475 0.1036 1.0000
8.500 1.0937 0.02886 0.01895 -0.0452 0.0978 1.0000
8.750 1.1127 0.03025 0.02038 -0.0439 0.0926 1.0000
9.000 1.1448 0.03242 0.02232 -0.0449 0.0869 1.0000
9.250 1.1684 0.03378 0.02389 -0.0442 0.0829 1.0000
9.500 1.2007 0.03572 0.02596 -0.0450 0.0795 1.0000
9.750 1.2362 0.03829 0.02844 -0.0466 0.0755 1.0000
10.000 1.2608 0.04079 0.03121 -0.0463 0.0729 1.0000
10.250 1.2799 0.04315 0.03393 -0.0450 0.0715 1.0000
10.500 1.2940 0.04557 0.03669 -0.0433 0.0697 1.0000
10.750 1.3058 0.04797 0.03940 -0.0414 0.0678 1.0000
11.000 1.3183 0.05060 0.04224 -0.0398 0.0661 1.0000
11.250 1.3257 0.05377 0.04573 -0.0377 0.0655 1.0000
11.500 1.3293 0.05704 0.04930 -0.0353 0.0650 1.0000
11.750 1.3301 0.06042 0.05293 -0.0330 0.0643 1.0000
12.000 1.3275 0.06427 0.05699 -0.0306 0.0636 1.0000
12.250 1.3176 0.06836 0.06133 -0.0277 0.0632 1.0000
12.500 1.3013 0.07158 0.06481 -0.0242 0.0632 1.0000
12.750 1.2791 0.07441 0.06793 -0.0208 0.0635 1.0000
13.000 1.0716 0.07980 0.07448 -0.0137 0.0702 1.0000
13.250 1.0346 0.08693 0.08182 -0.0155 0.0710 1.0000
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Polar data table (+)
Polar graphs
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