GOE 458 AIRFOIL (goe458-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 458 AIRFOIL (goe458-il) Reynolds number: 500,000 Max Cl/Cd: 115.29 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe458-il-500000-n5.txt Download as CSV file: xf-goe458-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 458 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.1732 0.10012 0.09800 -0.0500 0.9850 0.0045
-9.750 -0.2934 0.11551 0.11329 -0.0342 1.0000 0.0056
-9.500 -0.2878 0.11187 0.10966 -0.0356 0.9993 0.0050
-9.250 -0.2745 0.10783 0.10562 -0.0390 0.9971 0.0044
-9.000 -0.2617 0.10387 0.10166 -0.0423 0.9931 0.0040
-8.750 -0.2484 0.09982 0.09762 -0.0458 0.9889 0.0045
-8.500 -0.2358 0.09611 0.09391 -0.0490 0.9838 0.0041
-8.250 -0.2224 0.09228 0.09008 -0.0524 0.9785 0.0045
-8.000 -0.2099 0.08859 0.08640 -0.0557 0.9723 0.0042
-7.750 -0.1899 0.08475 0.08256 -0.0611 0.9670 0.0057
-7.500 -0.1658 0.08188 0.07968 -0.0670 0.9594 0.0064
-7.250 -0.1446 0.07758 0.07537 -0.0730 0.9529 0.0064
-7.000 -0.1238 0.07347 0.07124 -0.0788 0.9435 0.0065
-6.750 -0.1023 0.06906 0.06680 -0.0848 0.9336 0.0064
-6.500 -0.0856 0.06417 0.06187 -0.0902 0.9223 0.0061
-6.250 -0.0687 0.05979 0.05745 -0.0955 0.9099 0.0058
-6.000 -0.0510 0.05584 0.05341 -0.1003 0.8980 0.0056
-5.750 -0.0324 0.05161 0.04911 -0.1051 0.8866 0.0055
-5.500 -0.0119 0.04746 0.04488 -0.1097 0.8762 0.0053
-5.250 0.0109 0.04298 0.04029 -0.1144 0.8666 0.0052
-5.000 0.0364 0.03795 0.03510 -0.1191 0.8578 0.0050
-4.750 0.0637 0.03255 0.02950 -0.1234 0.8494 0.0049
-4.500 0.0916 0.02550 0.02207 -0.1269 0.8413 0.0047
-4.250 0.1175 0.02054 0.01655 -0.1283 0.8342 0.0046
-4.000 0.1429 0.01723 0.01268 -0.1284 0.8266 0.0046
-3.750 0.1689 0.01469 0.00958 -0.1280 0.8196 0.0047
-3.500 0.1953 0.01270 0.00709 -0.1274 0.8116 0.0055
-3.250 0.2225 0.01226 0.00642 -0.1271 0.8037 0.0064
-3.000 0.2480 0.01115 0.00516 -0.1268 0.7955 0.0076
-2.750 0.2746 0.01111 0.00510 -0.1268 0.7869 0.0106
-2.500 0.3012 0.01044 0.00420 -0.1264 0.7784 0.0113
-2.250 0.3277 0.00983 0.00344 -0.1259 0.7685 0.0119
-2.000 0.3542 0.00941 0.00290 -0.1255 0.7582 0.0135
-1.750 0.3810 0.00937 0.00278 -0.1253 0.7472 0.0152
-1.500 0.4071 0.00890 0.00218 -0.1249 0.7363 0.0175
-1.250 0.4334 0.00871 0.00189 -0.1246 0.7244 0.0194
-0.750 0.4861 0.00853 0.00142 -0.1239 0.6985 0.0212
-0.500 0.5125 0.00852 0.00130 -0.1236 0.6860 0.0210
-0.250 0.5389 0.00852 0.00122 -0.1233 0.6743 0.0208
0.000 0.5650 0.00856 0.00117 -0.1229 0.6611 0.0206
0.250 0.5907 0.00863 0.00115 -0.1224 0.6456 0.0206
0.500 0.6164 0.00871 0.00116 -0.1220 0.6309 0.0208
0.750 0.6424 0.00878 0.00117 -0.1217 0.6182 0.0213
1.250 0.6941 0.00872 0.00123 -0.1211 0.5961 0.1184
1.500 0.7200 0.00873 0.00132 -0.1208 0.5862 0.1551
1.750 0.7456 0.00863 0.00147 -0.1205 0.5761 0.2469
2.000 0.7765 0.00725 0.00176 -0.1218 0.5648 1.0000
2.250 0.8020 0.00739 0.00184 -0.1214 0.5525 1.0000
2.500 0.8276 0.00753 0.00194 -0.1210 0.5412 1.0000
2.750 0.8532 0.00767 0.00204 -0.1206 0.5321 1.0000
3.000 0.8792 0.00779 0.00215 -0.1203 0.5231 1.0000
3.250 0.9049 0.00792 0.00228 -0.1200 0.5143 1.0000
3.500 0.9304 0.00807 0.00248 -0.1196 0.5045 1.0000
3.750 0.9525 0.00840 0.00265 -0.1186 0.4641 1.0000
4.000 0.9641 0.00956 0.00312 -0.1158 0.3283 1.0000
4.250 0.9752 0.01098 0.00385 -0.1132 0.2047 1.0000
4.500 0.9778 0.01322 0.00507 -0.1093 0.0054 1.0000
4.750 1.0001 0.01367 0.00563 -0.1082 0.0030 1.0000
5.000 1.0223 0.01410 0.00614 -0.1072 0.0027 1.0000
5.250 1.0431 0.01465 0.00680 -0.1059 0.0025 1.0000
5.500 1.0622 0.01536 0.00762 -0.1043 0.0024 1.0000
5.750 1.0799 0.01617 0.00852 -0.1025 0.0023 1.0000
6.000 1.0959 0.01708 0.00953 -0.1004 0.0023 1.0000
6.250 1.1099 0.01810 0.01065 -0.0980 0.0023 1.0000
6.500 1.1222 0.01923 0.01187 -0.0953 0.0023 1.0000
6.750 1.1324 0.02051 0.01336 -0.0923 0.0024 1.0000
7.000 1.1432 0.02174 0.01467 -0.0894 0.0024 1.0000
7.250 1.1558 0.02316 0.01617 -0.0869 0.0024 1.0000
7.500 1.1728 0.02483 0.01792 -0.0851 0.0025 1.0000
7.750 1.1956 0.02675 0.01994 -0.0844 0.0025 1.0000
8.000 1.2233 0.02924 0.02257 -0.0845 0.0026 1.0000
8.250 1.2489 0.03222 0.02574 -0.0842 0.0028 1.0000
9.000 1.3058 0.03804 0.03202 -0.0804 0.0034 1.0000
9.250 1.3197 0.04045 0.03468 -0.0783 0.0035 1.0000
9.500 1.3295 0.04312 0.03760 -0.0758 0.0036 1.0000
9.750 1.3340 0.04576 0.04048 -0.0728 0.0035 1.0000
10.000 1.3333 0.04835 0.04331 -0.0692 0.0035 1.0000
10.250 1.3280 0.05084 0.04607 -0.0652 0.0035 1.0000
10.500 1.3196 0.05343 0.04886 -0.0613 0.0035 1.0000
10.750 1.3100 0.05620 0.05184 -0.0577 0.0035 1.0000
11.000 1.2983 0.05913 0.05496 -0.0547 0.0035 1.0000
11.250 1.2851 0.06235 0.05837 -0.0522 0.0034 1.0000
11.500 1.2711 0.06590 0.06211 -0.0502 0.0034 1.0000
11.750 1.2558 0.06977 0.06616 -0.0490 0.0034 1.0000
12.000 1.2397 0.07405 0.07062 -0.0483 0.0034 1.0000
12.250 1.2237 0.07861 0.07535 -0.0485 0.0034 1.0000
12.500 1.2065 0.08373 0.08064 -0.0494 0.0034 1.0000
12.750 1.1897 0.08918 0.08624 -0.0510 0.0035 1.0000
13.000 1.1724 0.09521 0.09243 -0.0533 0.0035 1.0000
13.250 1.1549 0.10175 0.09912 -0.0565 0.0035 1.0000
13.500 1.1370 0.10909 0.10662 -0.0605 0.0035 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 458 AIRFOIL (goe458-il)