GOE 498 AIRFOIL (goe498-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 498 AIRFOIL (goe498-il) Reynolds number: 100,000 Max Cl/Cd: 53.48 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe498-il-100000-n5.txt Download as CSV file: xf-goe498-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 498 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.3139 0.10120 0.09561 -0.0624 0.9868 0.0659
-10.750 -0.3100 0.09532 0.08971 -0.0675 0.9826 0.0666
-10.250 -0.5126 0.04781 0.04140 -0.1122 0.9451 0.0687
-10.000 -0.5046 0.04354 0.03678 -0.1174 0.9362 0.0702
-9.750 -0.4943 0.03993 0.03272 -0.1206 0.9270 0.0718
-9.500 -0.4707 0.03688 0.02922 -0.1242 0.9220 0.0736
-9.250 -0.4531 0.03576 0.02809 -0.1236 0.9126 0.0748
-9.000 -0.4229 0.03430 0.02650 -0.1255 0.9081 0.0768
-8.750 -0.3897 0.03243 0.02433 -0.1284 0.9049 0.0796
-8.500 -0.3753 0.03117 0.02282 -0.1272 0.8941 0.0814
-8.250 -0.3427 0.03013 0.02183 -0.1286 0.8898 0.0835
-8.000 -0.3066 0.02893 0.02051 -0.1309 0.8866 0.0865
-7.750 -0.2893 0.02803 0.01939 -0.1296 0.8764 0.0890
-7.500 -0.2581 0.02712 0.01850 -0.1307 0.8712 0.0918
-7.000 -0.2027 0.02556 0.01672 -0.1313 0.8582 0.0986
-6.750 -0.1731 0.02487 0.01608 -0.1319 0.8522 0.1018
-6.500 -0.1382 0.02414 0.01528 -0.1333 0.8482 0.1064
-6.250 -0.1169 0.02366 0.01475 -0.1323 0.8391 0.1103
-6.000 -0.0872 0.02313 0.01421 -0.1327 0.8328 0.1149
-5.750 -0.0520 0.02248 0.01348 -0.1341 0.8286 0.1211
-5.500 -0.0310 0.02217 0.01317 -0.1329 0.8193 0.1263
-5.250 -0.0003 0.02166 0.01260 -0.1334 0.8133 0.1336
-5.000 0.0356 0.02110 0.01198 -0.1347 0.8091 0.1434
-4.750 0.0558 0.02083 0.01174 -0.1334 0.7994 0.1515
-4.500 0.0881 0.02035 0.01124 -0.1341 0.7937 0.1641
-4.250 0.1195 0.01995 0.01081 -0.1347 0.7881 0.1784
-4.000 0.1444 0.01973 0.01057 -0.1341 0.7797 0.1926
-3.750 0.1779 0.01939 0.01024 -0.1350 0.7741 0.2086
-3.500 0.2060 0.01923 0.01009 -0.1349 0.7671 0.2233
-3.250 0.2340 0.01909 0.00996 -0.1348 0.7597 0.2378
-3.000 0.2682 0.01887 0.00969 -0.1356 0.7543 0.2535
-2.750 0.2939 0.01883 0.00962 -0.1351 0.7463 0.2676
-2.500 0.3234 0.01875 0.00949 -0.1352 0.7391 0.2844
-2.250 0.3579 0.01859 0.00928 -0.1361 0.7338 0.3019
-2.000 0.3805 0.01864 0.00939 -0.1350 0.7249 0.3164
-1.750 0.4106 0.01856 0.00929 -0.1351 0.7178 0.3333
-1.500 0.4412 0.01849 0.00916 -0.1353 0.7111 0.3493
-1.250 0.4652 0.01851 0.00919 -0.1344 0.7017 0.3629
-1.000 0.4971 0.01839 0.00902 -0.1348 0.6945 0.3770
-0.750 0.5204 0.01844 0.00906 -0.1338 0.6850 0.3904
-0.500 0.5485 0.01839 0.00901 -0.1335 0.6768 0.4034
-0.250 0.5764 0.01838 0.00899 -0.1333 0.6693 0.4167
0.000 0.6003 0.01845 0.00907 -0.1324 0.6606 0.4307
0.250 0.6299 0.01842 0.00903 -0.1324 0.6539 0.4458
0.500 0.6526 0.01852 0.00919 -0.1313 0.6456 0.4609
0.750 0.6785 0.01858 0.00925 -0.1306 0.6378 0.4788
1.000 0.7069 0.01860 0.00928 -0.1305 0.6314 0.4998
1.250 0.7283 0.01875 0.00952 -0.1291 0.6228 0.5211
1.500 0.7557 0.01879 0.00957 -0.1287 0.6158 0.5438
1.750 0.7799 0.01889 0.00972 -0.1279 0.6079 0.5655
2.000 0.8041 0.01896 0.00982 -0.1270 0.5993 0.5871
2.250 0.8307 0.01898 0.00985 -0.1265 0.5917 0.6104
2.500 0.8522 0.01908 0.01004 -0.1251 0.5824 0.6347
2.750 0.8800 0.01904 0.01001 -0.1248 0.5752 0.6642
3.000 0.8987 0.01914 0.01027 -0.1229 0.5657 0.6982
3.250 0.9225 0.01904 0.01029 -0.1217 0.5578 0.7508
3.500 0.9478 0.01886 0.01042 -0.1209 0.5481 1.0000
3.750 0.9735 0.01906 0.01049 -0.1204 0.5393 1.0000
4.000 0.9952 0.01935 0.01072 -0.1193 0.5298 1.0000
4.250 1.0191 0.01960 0.01089 -0.1186 0.5212 1.0000
4.500 1.0416 0.01990 0.01113 -0.1176 0.5129 1.0000
4.750 1.0633 0.02021 0.01141 -0.1165 0.5046 1.0000
5.000 1.0881 0.02049 0.01160 -0.1159 0.4972 1.0000
5.250 1.1063 0.02087 0.01201 -0.1142 0.4880 1.0000
5.500 1.1306 0.02114 0.01217 -0.1135 0.4804 1.0000
5.750 1.1470 0.02156 0.01264 -0.1116 0.4712 1.0000
6.000 1.1681 0.02188 0.01290 -0.1104 0.4632 1.0000
6.250 1.1845 0.02229 0.01333 -0.1085 0.4548 1.0000
6.500 1.2019 0.02267 0.01368 -0.1067 0.4462 1.0000
6.750 1.2180 0.02310 0.01410 -0.1048 0.4375 1.0000
7.000 1.2334 0.02354 0.01454 -0.1028 0.4284 1.0000
7.250 1.2495 0.02401 0.01499 -0.1009 0.4201 1.0000
7.500 1.2643 0.02453 0.01554 -0.0989 0.4119 1.0000
7.750 1.2820 0.02500 0.01598 -0.0974 0.4048 1.0000
8.000 1.2948 0.02563 0.01668 -0.0953 0.3967 1.0000
8.250 1.3116 0.02615 0.01716 -0.0937 0.3896 1.0000
8.500 1.3237 0.02685 0.01794 -0.0916 0.3816 1.0000
8.750 1.3377 0.02749 0.01858 -0.0898 0.3740 1.0000
9.000 1.3510 0.02820 0.01931 -0.0879 0.3666 1.0000
9.250 1.3620 0.02900 0.02018 -0.0859 0.3585 1.0000
9.500 1.3758 0.02972 0.02086 -0.0842 0.3511 1.0000
9.750 1.3830 0.03073 0.02197 -0.0819 0.3421 1.0000
10.000 1.3943 0.03157 0.02275 -0.0800 0.3336 1.0000
10.250 1.3984 0.03281 0.02409 -0.0777 0.3234 1.0000
10.500 1.4055 0.03392 0.02518 -0.0756 0.3137 1.0000
10.750 1.4083 0.03534 0.02666 -0.0734 0.3030 1.0000
11.000 1.4126 0.03677 0.02811 -0.0714 0.2931 1.0000
11.250 1.4159 0.03830 0.02965 -0.0695 0.2830 1.0000
11.500 1.4182 0.04003 0.03147 -0.0678 0.2731 1.0000
11.750 1.4212 0.04173 0.03314 -0.0662 0.2637 1.0000
12.000 1.4213 0.04380 0.03531 -0.0647 0.2535 1.0000
12.250 1.4216 0.04591 0.03745 -0.0633 0.2438 1.0000
12.500 1.4212 0.04817 0.03970 -0.0621 0.2343 1.0000
12.750 1.4209 0.05058 0.04217 -0.0610 0.2251 1.0000
13.000 1.4204 0.05303 0.04461 -0.0601 0.2170 1.0000
13.250 1.4197 0.05566 0.04730 -0.0594 0.2089 1.0000
13.500 1.4193 0.05830 0.04993 -0.0588 0.2019 1.0000
13.750 1.4189 0.06105 0.05275 -0.0583 0.1952 1.0000
14.000 1.4186 0.06378 0.05543 -0.0578 0.1893 1.0000
14.250 1.4183 0.06672 0.05850 -0.0576 0.1833 1.0000
14.500 1.4175 0.06970 0.06148 -0.0575 0.1776 1.0000
14.750 1.4173 0.07266 0.06448 -0.0574 0.1721 1.0000
15.000 1.4154 0.07598 0.06790 -0.0575 0.1663 1.0000
15.250 1.4153 0.07896 0.07084 -0.0576 0.1607 1.0000
15.500 1.4123 0.08263 0.07466 -0.0580 0.1552 1.0000
15.750 1.4087 0.08636 0.07849 -0.0586 0.1493 1.0000
16.000 1.4062 0.08993 0.08210 -0.0591 0.1436 1.0000
16.250 1.3991 0.09446 0.08683 -0.0602 0.1371 1.0000
16.500 1.3941 0.09858 0.09099 -0.0613 0.1312 1.0000
16.750 1.3857 0.10352 0.09613 -0.0628 0.1241 1.0000
17.000 1.3790 0.10816 0.10085 -0.0643 0.1178 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 498 AIRFOIL (goe498-il)