GOE 711 AIRFOIL (goe711-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 711 AIRFOIL (goe711-il) Reynolds number: 100,000 Max Cl/Cd: 48.31 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe711-il-100000.txt Download as CSV file: xf-goe711-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 711 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3438 0.13443 0.13030 -0.0287 1.0000 0.0649
-8.750 -0.3634 0.13378 0.12974 -0.0278 1.0000 0.0650
-8.500 -0.3846 0.13308 0.12913 -0.0265 1.0000 0.0651
-8.250 -0.3440 0.12445 0.12048 -0.0228 1.0000 0.0670
-8.000 -0.3460 0.12227 0.11835 -0.0207 1.0000 0.0681
-7.750 -0.3481 0.12006 0.11618 -0.0200 0.9992 0.0694
-7.500 -0.3316 0.11606 0.11216 -0.0249 0.9938 0.0721
-7.250 -0.3238 0.11252 0.10862 -0.0301 0.9861 0.0749
-7.000 -0.3251 0.10919 0.10528 -0.0433 0.9738 0.0776
-6.750 -0.3125 0.10317 0.09923 -0.0514 0.9643 0.0790
-6.500 -0.2887 0.09939 0.09549 -0.0468 0.9606 0.0819
-6.250 -0.2681 0.09545 0.09152 -0.0508 0.9528 0.0856
-6.000 -0.2428 0.09083 0.08655 -0.0694 0.9388 0.0927
-5.750 -0.2277 0.08497 0.08077 -0.0699 0.9313 0.0944
-5.500 -0.1991 0.08108 0.07693 -0.0704 0.9274 0.0980
-5.250 -0.1806 0.07778 0.07354 -0.0736 0.9161 0.1047
-5.000 -0.1529 0.07281 0.06839 -0.0803 0.9080 0.1121
-4.750 -0.1248 0.06959 0.06511 -0.0830 0.8999 0.1207
-4.250 -0.0591 0.06156 0.05663 -0.0933 0.8842 0.1443
-3.750 0.0050 0.05518 0.05003 -0.0989 0.8691 0.1787
-3.500 0.0285 0.05266 0.04750 -0.0994 0.8589 0.1977
-3.250 0.0686 0.04952 0.04429 -0.1029 0.8542 0.2313
-3.000 0.0916 0.04739 0.04216 -0.1027 0.8439 0.2653
-2.750 0.1270 0.04465 0.03950 -0.1039 0.8395 0.3182
-2.000 0.3150 0.03345 0.02558 -0.1219 0.8238 0.1383
-1.750 0.3524 0.03145 0.02312 -0.1222 0.8142 0.1211
-1.500 0.4007 0.02912 0.02046 -0.1249 0.8091 0.1161
-1.250 0.4389 0.02797 0.01885 -0.1254 0.7994 0.1097
-1.000 0.4852 0.02618 0.01688 -0.1277 0.7928 0.1081
-0.750 0.5189 0.02497 0.01554 -0.1280 0.7807 0.1080
-0.500 0.5562 0.02355 0.01410 -0.1291 0.7698 0.1115
-0.250 0.5970 0.02223 0.01276 -0.1307 0.7597 0.1153
0.000 0.6277 0.02150 0.01199 -0.1306 0.7448 0.1189
0.250 0.6585 0.02095 0.01131 -0.1305 0.7294 0.1242
0.500 0.6887 0.02043 0.01070 -0.1305 0.7135 0.1376
0.750 0.7190 0.01998 0.01022 -0.1306 0.6972 0.1668
1.000 0.7487 0.01783 0.00995 -0.1297 0.6811 1.0000
1.250 0.7765 0.01798 0.00971 -0.1293 0.6642 1.0000
1.500 0.8023 0.01821 0.00966 -0.1287 0.6468 1.0000
1.750 0.8281 0.01847 0.00967 -0.1283 0.6300 1.0000
2.000 0.8541 0.01875 0.00971 -0.1278 0.6139 1.0000
2.250 0.8803 0.01906 0.00980 -0.1275 0.5988 1.0000
2.500 0.9072 0.01939 0.00989 -0.1274 0.5848 1.0000
2.750 0.9316 0.01978 0.01013 -0.1269 0.5704 1.0000
3.000 0.9559 0.02019 0.01041 -0.1263 0.5570 1.0000
3.250 0.9806 0.02062 0.01071 -0.1259 0.5446 1.0000
3.500 1.0074 0.02106 0.01098 -0.1259 0.5339 1.0000
3.750 1.0331 0.02150 0.01130 -0.1257 0.5233 1.0000
4.000 1.0567 0.02202 0.01176 -0.1252 0.5129 1.0000
4.250 1.0855 0.02247 0.01202 -0.1256 0.5044 1.0000
4.500 1.1071 0.02300 0.01258 -0.1248 0.4946 1.0000
4.750 1.1347 0.02352 0.01297 -0.1250 0.4870 1.0000
5.000 1.1572 0.02410 0.01358 -0.1244 0.4788 1.0000
5.250 1.1866 0.02464 0.01396 -0.1251 0.4724 1.0000
5.500 1.2064 0.02528 0.01471 -0.1240 0.4645 1.0000
5.750 1.2352 0.02579 0.01511 -0.1245 0.4583 1.0000
6.000 1.2556 0.02650 0.01592 -0.1236 0.4515 1.0000
6.250 1.2800 0.02711 0.01652 -0.1234 0.4454 1.0000
6.500 1.3098 0.02776 0.01708 -0.1242 0.4405 1.0000
6.750 1.3263 0.02859 0.01810 -0.1227 0.4345 1.0000
7.000 1.3510 0.02923 0.01876 -0.1226 0.4292 1.0000
7.250 1.3804 0.02992 0.01936 -0.1233 0.4246 1.0000
7.500 1.3941 0.03085 0.02054 -0.1214 0.4189 1.0000
7.750 1.4171 0.03159 0.02133 -0.1211 0.4140 1.0000
8.000 1.4478 0.03230 0.02197 -0.1221 0.4102 1.0000
8.250 1.4605 0.03344 0.02336 -0.1201 0.4056 1.0000
8.500 1.4766 0.03437 0.02444 -0.1186 0.4004 1.0000
8.750 1.5049 0.03490 0.02493 -0.1192 0.3956 1.0000
9.000 1.5209 0.03593 0.02610 -0.1178 0.3908 1.0000
9.250 1.5313 0.03703 0.02742 -0.1154 0.3856 1.0000
9.500 1.5568 0.03758 0.02798 -0.1155 0.3808 1.0000
9.750 1.5790 0.03844 0.02889 -0.1152 0.3763 1.0000
10.000 1.5796 0.03983 0.03056 -0.1113 0.3713 1.0000
10.250 1.6057 0.03996 0.03069 -0.1114 0.3650 1.0000
10.500 1.6173 0.04065 0.03147 -0.1092 0.3586 1.0000
10.750 1.6274 0.04095 0.03185 -0.1066 0.3512 1.0000
11.000 1.6470 0.04093 0.03179 -0.1055 0.3436 1.0000
11.250 1.6447 0.04177 0.03283 -0.1010 0.3374 1.0000
11.500 1.6804 0.04127 0.03217 -0.1025 0.3301 1.0000
11.750 1.6602 0.04287 0.03408 -0.0955 0.3261 1.0000
12.000 1.6611 0.04376 0.03512 -0.0918 0.3210 1.0000
12.250 1.6867 0.04342 0.03469 -0.0917 0.3140 1.0000
12.500 1.6655 0.04538 0.03696 -0.0854 0.3097 1.0000
12.750 1.6774 0.04542 0.03701 -0.0834 0.3028 1.0000
13.000 1.6763 0.04654 0.03825 -0.0802 0.2974 1.0000
13.250 1.6596 0.04873 0.04070 -0.0757 0.2927 1.0000
13.500 1.6733 0.04831 0.04018 -0.0740 0.2839 1.0000
13.750 1.6483 0.05146 0.04365 -0.0695 0.2793 1.0000
14.000 1.6567 0.05148 0.04358 -0.0677 0.2701 1.0000
14.250 1.6294 0.05551 0.04797 -0.0642 0.2658 1.0000
14.500 1.6179 0.05798 0.05057 -0.0621 0.2580 1.0000
14.750 1.5994 0.06174 0.05452 -0.0602 0.2516 1.0000
15.000 1.5810 0.06590 0.05886 -0.0591 0.2441 1.0000
15.250 1.5601 0.07087 0.06402 -0.0585 0.2372 1.0000
15.500 1.5353 0.07684 0.07018 -0.0585 0.2307 1.0000
15.750 1.5115 0.08305 0.07655 -0.0590 0.2239 1.0000
16.000 1.4776 0.09128 0.08502 -0.0602 0.2218 1.0000
16.250 1.4364 0.10105 0.09499 -0.0620 0.2226 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 711 AIRFOIL (goe711-il)