GOE 630 AIRFOIL (goe630-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 630 AIRFOIL (goe630-il) Reynolds number: 100,000 Max Cl/Cd: 48.43 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe630-il-100000.txt Download as CSV file: xf-goe630-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 630 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.0307 0.10349 0.09896 -0.0807 0.9315 0.0792
-8.000 -0.0155 0.10146 0.09692 -0.0879 0.9260 0.0824
-7.750 -0.0297 0.10233 0.09782 -0.0902 0.9124 0.0832
-7.500 -0.0052 0.09692 0.09240 -0.0956 0.9091 0.0844
-7.250 0.0417 0.09035 0.08577 -0.0976 0.9090 0.0874
-7.000 0.0409 0.08900 0.08444 -0.0959 0.8974 0.0896
-6.750 0.0636 0.08602 0.08141 -0.1029 0.8925 0.0951
-6.500 0.0531 0.08747 0.08283 -0.1108 0.8768 0.0972
-6.250 0.0478 0.08566 0.08104 -0.1100 0.8648 0.0978
-6.000 0.0783 0.07960 0.07501 -0.1067 0.8633 0.1000
-5.750 0.1108 0.07570 0.07106 -0.1097 0.8601 0.1041
-5.500 0.1078 0.07471 0.07008 -0.1082 0.8482 0.1067
-5.250 0.1219 0.07482 0.06994 -0.1209 0.8354 0.1131
-5.000 0.1358 0.07000 0.06524 -0.1168 0.8305 0.1148
-4.750 0.1654 0.06647 0.06169 -0.1170 0.8271 0.1199
-4.500 0.1704 0.06542 0.06061 -0.1171 0.8156 0.1257
-4.250 0.2042 0.06166 0.05670 -0.1236 0.8106 0.1320
-4.000 0.2094 0.06021 0.05527 -0.1207 0.8004 0.1351
-3.750 0.2500 0.05729 0.05207 -0.1286 0.7945 0.1477
-3.500 0.2683 0.05465 0.04948 -0.1270 0.7884 0.1515
-3.250 0.2942 0.05296 0.04756 -0.1305 0.7788 0.1649
-3.000 0.3253 0.04989 0.04449 -0.1309 0.7753 0.1738
-2.750 0.3355 0.04871 0.04323 -0.1301 0.7634 0.1846
-2.500 0.3733 0.04597 0.04034 -0.1329 0.7592 0.2018
-2.250 0.3875 0.04499 0.03926 -0.1320 0.7480 0.2169
-2.000 0.4209 0.04248 0.03667 -0.1333 0.7431 0.2357
-1.750 0.4358 0.04134 0.03550 -0.1318 0.7329 0.2536
-1.500 0.4668 0.03939 0.03343 -0.1325 0.7270 0.2866
-1.000 0.5648 0.03265 0.02464 -0.1399 0.7118 0.1288
-0.750 0.6045 0.02996 0.02171 -0.1415 0.7075 0.1224
-0.500 0.6211 0.02923 0.02073 -0.1393 0.6959 0.1191
-0.250 0.6616 0.02771 0.01885 -0.1405 0.6910 0.1216
0.000 0.6785 0.02734 0.01833 -0.1384 0.6795 0.1221
0.250 0.7188 0.02600 0.01666 -0.1395 0.6743 0.1241
0.500 0.7359 0.02590 0.01643 -0.1373 0.6627 0.1281
0.750 0.7749 0.02470 0.01514 -0.1385 0.6573 0.1369
1.000 0.7915 0.02469 0.01512 -0.1364 0.6459 0.1433
1.250 0.8294 0.02379 0.01416 -0.1374 0.6403 0.1610
1.500 0.8466 0.02378 0.01422 -0.1354 0.6291 0.1798
1.750 0.9024 0.02135 0.01339 -0.1404 0.6215 1.0000
2.000 0.9262 0.02154 0.01334 -0.1394 0.6120 1.0000
2.250 0.9520 0.02173 0.01331 -0.1388 0.6038 1.0000
2.500 0.9776 0.02191 0.01332 -0.1382 0.5955 1.0000
2.750 1.0025 0.02217 0.01343 -0.1375 0.5875 1.0000
3.000 1.0279 0.02237 0.01349 -0.1369 0.5793 1.0000
3.250 1.0528 0.02267 0.01366 -0.1363 0.5716 1.0000
3.500 1.0766 0.02297 0.01387 -0.1356 0.5636 1.0000
3.750 1.1048 0.02323 0.01397 -0.1355 0.5571 1.0000
4.000 1.1231 0.02381 0.01455 -0.1341 0.5491 1.0000
4.250 1.1571 0.02389 0.01444 -0.1349 0.5437 1.0000
4.500 1.1693 0.02474 0.01536 -0.1326 0.5356 1.0000
4.750 1.1979 0.02498 0.01548 -0.1327 0.5296 1.0000
5.000 1.2193 0.02558 0.01607 -0.1318 0.5233 1.0000
5.250 1.2377 0.02626 0.01676 -0.1304 0.5167 1.0000
5.500 1.2695 0.02653 0.01691 -0.1311 0.5121 1.0000
5.750 1.2828 0.02751 0.01799 -0.1291 0.5062 1.0000
6.000 1.3008 0.02826 0.01878 -0.1278 0.5004 1.0000
6.250 1.3328 0.02854 0.01896 -0.1285 0.4960 1.0000
6.500 1.3451 0.02960 0.02011 -0.1264 0.4905 1.0000
6.750 1.3574 0.03062 0.02123 -0.1244 0.4851 1.0000
7.000 1.3851 0.03112 0.02171 -0.1245 0.4810 1.0000
7.250 1.4154 0.03170 0.02223 -0.1252 0.4775 1.0000
7.500 1.4054 0.03365 0.02444 -0.1201 0.4719 1.0000
7.750 1.4203 0.03464 0.02551 -0.1186 0.4673 1.0000
8.000 1.4544 0.03492 0.02575 -0.1197 0.4637 1.0000
8.250 1.4640 0.03635 0.02728 -0.1175 0.4598 1.0000
8.500 1.4318 0.03946 0.03065 -0.1101 0.4546 1.0000
8.750 1.4396 0.04085 0.03211 -0.1079 0.4506 1.0000
9.000 1.4830 0.04077 0.03204 -0.1101 0.4475 1.0000
9.250 1.1829 0.06768 0.05934 -0.0883 0.4294 1.0000
9.500 1.2393 0.06401 0.05570 -0.0880 0.4294 1.0000
9.750 1.3052 0.05981 0.05152 -0.0886 0.4293 1.0000
10.000 0.8251 0.12867 0.12085 -0.0990 0.4539 1.0000
10.250 0.8437 0.13075 0.12294 -0.0990 0.4493 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 630 AIRFOIL (goe630-il)