GOE 394 AIRFOIL (goe394-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 394 AIRFOIL (goe394-il) Reynolds number: 100,000 Max Cl/Cd: 64.42 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe394-il-100000.txt Download as CSV file: xf-goe394-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 394 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2844 0.09706 0.09251 -0.0246 1.0000 0.0701
-7.250 -0.2845 0.09542 0.09098 -0.0249 1.0000 0.0723
-7.000 -0.3024 0.09558 0.09128 -0.0214 1.0000 0.0731
-6.750 -0.3253 0.09614 0.09194 -0.0169 1.0000 0.0733
-6.500 -0.3398 0.09632 0.09220 -0.0155 1.0000 0.0743
-6.250 -0.3460 0.09728 0.09319 -0.0194 1.0000 0.0753
-6.000 -0.3313 0.09404 0.08995 -0.0255 0.9976 0.0762
-5.750 -0.3240 0.08860 0.08453 -0.0209 0.9956 0.0776
-5.500 -0.3055 0.08488 0.08079 -0.0230 0.9921 0.0800
-5.250 -0.2807 0.08144 0.07731 -0.0282 0.9875 0.0835
-5.000 -0.2144 0.07891 0.07454 -0.0513 0.9813 0.0884
-4.750 -0.2137 0.07410 0.06981 -0.0460 0.9781 0.0897
-4.500 -0.1955 0.07060 0.06629 -0.0464 0.9742 0.0927
-4.250 -0.1076 0.06814 0.06336 -0.0707 0.9678 0.1012
-4.000 -0.0969 0.06279 0.05814 -0.0685 0.9617 0.1023
-3.750 -0.0664 0.05875 0.05411 -0.0705 0.9567 0.1060
-3.500 -0.0049 0.05553 0.05054 -0.0839 0.9489 0.1152
-3.250 0.0208 0.05184 0.04689 -0.0849 0.9440 0.1177
-3.000 0.0864 0.04880 0.04346 -0.0968 0.9397 0.1288
-2.750 0.1014 0.04603 0.04078 -0.0957 0.9329 0.1314
-2.500 0.1506 0.04323 0.03777 -0.1023 0.9286 0.1449
-2.250 0.1985 0.04082 0.03513 -0.1082 0.9243 0.1589
-2.000 0.2215 0.03885 0.03319 -0.1084 0.9171 0.1653
-1.750 0.2739 0.03655 0.03064 -0.1146 0.9130 0.1897
-1.500 0.3041 0.03485 0.02893 -0.1159 0.9066 0.2080
-1.250 0.3377 0.03344 0.02746 -0.1178 0.9000 0.2496
-1.000 0.4191 0.02753 0.02013 -0.1270 0.8972 0.1233
-0.750 0.4517 0.02588 0.01808 -0.1277 0.8894 0.1218
-0.500 0.4905 0.02447 0.01636 -0.1292 0.8832 0.1328
-0.250 0.5216 0.02343 0.01510 -0.1295 0.8750 0.1480
0.000 0.5561 0.02256 0.01397 -0.1299 0.8678 0.1729
0.250 0.5821 0.02223 0.01376 -0.1295 0.8579 0.1975
0.500 0.6149 0.02164 0.01310 -0.1294 0.8508 0.2146
0.750 0.6406 0.02132 0.01275 -0.1285 0.8394 0.2214
1.000 0.6680 0.02097 0.01238 -0.1276 0.8291 0.2256
1.250 0.6973 0.02050 0.01184 -0.1266 0.8205 0.2304
1.500 0.7226 0.02024 0.01164 -0.1254 0.8079 0.2378
1.750 0.7486 0.02002 0.01141 -0.1241 0.7954 0.2459
2.000 0.7746 0.01973 0.01118 -0.1228 0.7832 0.2535
2.250 0.8007 0.01943 0.01091 -0.1213 0.7713 0.2649
2.500 0.8269 0.01912 0.01065 -0.1198 0.7593 0.2849
2.750 0.8514 0.01857 0.01073 -0.1186 0.7449 0.4413
3.000 0.8779 0.01785 0.01051 -0.1172 0.7302 1.0000
3.250 0.9031 0.01796 0.01049 -0.1159 0.7152 1.0000
3.500 0.9283 0.01808 0.01052 -0.1146 0.7001 1.0000
3.750 0.9535 0.01819 0.01060 -0.1135 0.6849 1.0000
4.000 0.9788 0.01829 0.01067 -0.1125 0.6698 1.0000
4.250 1.0042 0.01833 0.01069 -0.1114 0.6546 1.0000
4.500 1.0298 0.01827 0.01061 -0.1104 0.6392 1.0000
4.750 1.0553 0.01815 0.01051 -0.1092 0.6231 1.0000
5.000 1.0802 0.01811 0.01050 -0.1082 0.6051 1.0000
5.250 1.1056 0.01806 0.01045 -0.1072 0.5882 1.0000
5.500 1.1311 0.01804 0.01041 -0.1061 0.5712 1.0000
5.750 1.1564 0.01811 0.01044 -0.1050 0.5538 1.0000
6.000 1.1815 0.01834 0.01060 -0.1040 0.5368 1.0000
6.250 1.2061 0.01875 0.01101 -0.1031 0.5198 1.0000
6.500 1.2304 0.01924 0.01156 -0.1023 0.5027 1.0000
6.750 1.2542 0.01975 0.01213 -0.1014 0.4850 1.0000
7.000 1.2775 0.02026 0.01268 -0.1003 0.4658 1.0000
7.250 1.2993 0.02071 0.01316 -0.0990 0.4423 1.0000
7.500 1.3174 0.02101 0.01349 -0.0971 0.4084 1.0000
7.750 1.3336 0.02134 0.01383 -0.0951 0.3679 1.0000
8.000 1.3488 0.02189 0.01440 -0.0931 0.3204 1.0000
8.250 1.3601 0.02297 0.01521 -0.0908 0.2540 1.0000
8.500 1.3650 0.02496 0.01666 -0.0881 0.1833 1.0000
8.750 1.3637 0.02773 0.01893 -0.0848 0.1319 1.0000
9.000 1.3595 0.03070 0.02157 -0.0811 0.0999 1.0000
9.250 1.3604 0.03316 0.02385 -0.0778 0.0841 1.0000
9.500 1.3673 0.03494 0.02569 -0.0751 0.0738 1.0000
9.750 1.3764 0.03684 0.02764 -0.0727 0.0664 1.0000
10.000 1.3875 0.03880 0.02951 -0.0707 0.0612 1.0000
10.250 1.4093 0.04117 0.03198 -0.0694 0.0574 1.0000
10.500 1.4272 0.04330 0.03427 -0.0680 0.0538 1.0000
10.750 1.4471 0.04580 0.03670 -0.0673 0.0501 1.0000
11.000 1.4692 0.04931 0.04052 -0.0666 0.0482 1.0000
11.250 1.4809 0.05251 0.04410 -0.0648 0.0474 1.0000
11.500 1.4859 0.05592 0.04791 -0.0625 0.0469 1.0000
11.750 1.4846 0.05946 0.05183 -0.0598 0.0466 1.0000
12.000 1.4774 0.06301 0.05573 -0.0571 0.0464 1.0000
12.250 1.4658 0.06670 0.05975 -0.0546 0.0462 1.0000
12.500 1.4512 0.07067 0.06403 -0.0526 0.0461 1.0000
12.750 1.4341 0.07500 0.06867 -0.0513 0.0461 1.0000
13.000 1.4147 0.07980 0.07374 -0.0507 0.0462 1.0000
13.250 1.3935 0.08506 0.07927 -0.0508 0.0464 1.0000
13.500 1.3708 0.09085 0.08530 -0.0518 0.0467 1.0000
13.750 1.3471 0.09719 0.09186 -0.0537 0.0470 1.0000
14.000 1.3233 0.10410 0.09896 -0.0563 0.0475 1.0000
14.250 1.2998 0.11155 0.10657 -0.0598 0.0480 1.0000
14.500 1.2774 0.11952 0.11468 -0.0640 0.0486 1.0000
14.750 1.2585 0.12772 0.12298 -0.0682 0.0491 1.0000
15.000 1.1147 0.18748 0.18280 -0.1116 0.0895 1.0000
15.250 1.1475 0.18399 0.17947 -0.1024 0.0846 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 394 AIRFOIL (goe394-il)