GOE 510 AIRFOIL (goe510-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 510 AIRFOIL (goe510-il) Reynolds number: 100,000 Max Cl/Cd: 48.44 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe510-il-100000.txt Download as CSV file: xf-goe510-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 510 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 -0.3354 0.12078 0.11647 -0.0186 1.0000 0.1066
-8.000 -0.3592 0.12061 0.11639 -0.0211 0.9979 0.1086
-7.750 -0.3892 0.12018 0.11599 -0.0324 0.9855 0.1095
-7.500 -0.3111 0.11123 0.10697 -0.0280 0.9866 0.1130
-7.250 -0.2870 0.10779 0.10350 -0.0322 0.9797 0.1177
-7.000 -0.2951 0.10588 0.10160 -0.0394 0.9684 0.1226
-6.750 -0.3077 0.10275 0.09844 -0.0475 0.9536 0.1246
-6.500 -0.2685 0.09775 0.09345 -0.0456 0.9478 0.1270
-6.250 -0.2364 0.09392 0.08958 -0.0496 0.9417 0.1316
-6.000 -0.2549 0.09300 0.08846 -0.0587 0.9215 0.1391
-5.750 -0.2356 0.08764 0.08318 -0.0581 0.9154 0.1407
-5.500 -0.2024 0.08381 0.07936 -0.0588 0.9118 0.1442
-5.250 -0.1939 0.08154 0.07708 -0.0588 0.9005 0.1497
-4.750 -0.1636 0.07513 0.07055 -0.0632 0.8828 0.1605
-4.500 -0.1478 0.07290 0.06823 -0.0645 0.8734 0.1682
-4.250 -0.1276 0.06934 0.06454 -0.0677 0.8656 0.1746
-3.750 -0.0856 0.06424 0.05920 -0.0705 0.8489 0.1905
-3.500 -0.0732 0.06174 0.05675 -0.0686 0.8389 0.1936
-3.250 -0.0417 0.05922 0.05400 -0.0716 0.8326 0.2079
-3.000 -0.0280 0.05690 0.05173 -0.0697 0.8226 0.2123
-2.750 0.0050 0.05435 0.04895 -0.0722 0.8164 0.2259
-2.500 0.0459 0.05099 0.04553 -0.0748 0.8137 0.2344
-2.000 0.0999 0.03674 0.02897 -0.0766 0.7979 0.1233
-1.750 0.1162 0.03514 0.02718 -0.0744 0.7861 0.1221
-1.500 0.1605 0.03232 0.02397 -0.0764 0.7827 0.1206
-1.250 0.2135 0.02973 0.02094 -0.0797 0.7804 0.1212
-1.000 0.2342 0.02893 0.01987 -0.0777 0.7679 0.1234
-0.750 0.2927 0.02672 0.01755 -0.0823 0.7648 0.1274
-0.500 0.3215 0.02588 0.01666 -0.0818 0.7536 0.1305
-0.250 0.3847 0.02429 0.01489 -0.0873 0.7486 0.1383
0.000 0.4197 0.02345 0.01415 -0.0881 0.7371 0.1471
0.250 0.4815 0.02204 0.01275 -0.0935 0.7299 0.1630
0.500 0.5130 0.02153 0.01231 -0.0936 0.7161 0.1884
0.750 0.5462 0.02113 0.01221 -0.0939 0.7031 0.2777
1.000 0.5846 0.02077 0.01178 -0.0953 0.6907 0.3442
1.250 0.6284 0.02027 0.01123 -0.0981 0.6783 0.3924
1.500 0.6519 0.01998 0.01104 -0.0971 0.6638 0.4297
1.750 0.6708 0.01963 0.01088 -0.0952 0.6507 0.4797
2.000 0.8939 0.01912 0.01090 -0.1336 0.6276 1.0000
2.250 0.9097 0.01948 0.01114 -0.1312 0.6148 1.0000
2.500 0.9314 0.01978 0.01127 -0.1299 0.6037 1.0000
2.750 0.9556 0.02003 0.01134 -0.1291 0.5933 1.0000
3.000 0.9689 0.02044 0.01169 -0.1263 0.5826 1.0000
3.250 0.9974 0.02068 0.01171 -0.1263 0.5738 1.0000
3.500 1.0065 0.02112 0.01217 -0.1228 0.5639 1.0000
3.750 1.0328 0.02139 0.01227 -0.1225 0.5558 1.0000
4.000 1.0428 0.02183 0.01274 -0.1191 0.5467 1.0000
4.250 1.0709 0.02211 0.01284 -0.1191 0.5393 1.0000
4.500 1.0781 0.02260 0.01340 -0.1152 0.5308 1.0000
4.750 1.1061 0.02287 0.01351 -0.1152 0.5237 1.0000
5.000 1.1137 0.02341 0.01413 -0.1115 0.5160 1.0000
5.250 1.1361 0.02374 0.01436 -0.1105 0.5087 1.0000
5.500 1.1511 0.02424 0.01486 -0.1081 0.5016 1.0000
5.750 1.1660 0.02469 0.01532 -0.1058 0.4944 1.0000
6.000 1.1946 0.02507 0.01558 -0.1060 0.4881 1.0000
6.250 1.1985 0.02572 0.01634 -0.1016 0.4809 1.0000
6.500 1.2248 0.02608 0.01661 -0.1014 0.4743 1.0000
6.750 1.2355 0.02673 0.01732 -0.0984 0.4676 1.0000
7.000 1.2491 0.02729 0.01791 -0.0959 0.4608 1.0000
7.250 1.2843 0.02774 0.01821 -0.0974 0.4548 1.0000
7.500 1.2797 0.02858 0.01925 -0.0917 0.4482 1.0000
7.750 1.3008 0.02907 0.01972 -0.0906 0.4417 1.0000
8.000 1.3214 0.02974 0.02038 -0.0895 0.4356 1.0000
8.250 1.3224 0.03058 0.02136 -0.0849 0.4291 1.0000
8.500 1.3518 0.03103 0.02175 -0.0853 0.4230 1.0000
8.750 1.3579 0.03196 0.02278 -0.0818 0.4171 1.0000
9.000 1.3610 0.03279 0.02373 -0.0777 0.4108 1.0000
9.250 1.3992 0.03317 0.02400 -0.0797 0.4047 1.0000
9.500 1.3891 0.03433 0.02535 -0.0735 0.3990 1.0000
9.750 1.3930 0.03513 0.02626 -0.0696 0.3928 1.0000
10.000 1.4341 0.03523 0.02623 -0.0719 0.3856 1.0000
10.250 1.4093 0.03646 0.02769 -0.0634 0.3799 1.0000
10.500 1.4630 0.03558 0.02655 -0.0673 0.3695 1.0000
10.750 1.4298 0.03693 0.02815 -0.0575 0.3649 1.0000
11.000 1.4334 0.03704 0.02829 -0.0535 0.3571 1.0000
11.250 1.4444 0.03711 0.02833 -0.0507 0.3492 1.0000
11.500 1.4333 0.03804 0.02939 -0.0451 0.3430 1.0000
11.750 1.4648 0.03741 0.02862 -0.0453 0.3348 1.0000
12.000 1.4411 0.03906 0.03051 -0.0386 0.3297 1.0000
12.250 1.4600 0.03877 0.03017 -0.0371 0.3221 1.0000
12.500 1.4518 0.03989 0.03141 -0.0328 0.3159 1.0000
12.750 1.4435 0.04105 0.03269 -0.0287 0.3096 1.0000
13.000 1.4636 0.04081 0.03238 -0.0276 0.3022 1.0000
13.250 1.4413 0.04312 0.03492 -0.0228 0.2970 1.0000
13.500 1.4638 0.04247 0.03415 -0.0218 0.2887 1.0000
13.750 1.4408 0.04519 0.03710 -0.0177 0.2835 1.0000
14.000 1.4365 0.04653 0.03853 -0.0151 0.2763 1.0000
14.250 1.4373 0.04779 0.03983 -0.0132 0.2698 1.0000
14.500 1.4197 0.05079 0.04303 -0.0107 0.2642 1.0000
14.750 1.4399 0.05051 0.04264 -0.0098 0.2570 1.0000
15.000 1.4089 0.05511 0.04752 -0.0075 0.2519 1.0000
15.250 1.4039 0.05729 0.04975 -0.0063 0.2446 1.0000
15.500 1.3922 0.06048 0.05304 -0.0053 0.2383 1.0000
15.750 1.3732 0.06485 0.05758 -0.0047 0.2327 1.0000
16.000 1.3774 0.06643 0.05915 -0.0042 0.2254 1.0000
16.250 1.3442 0.07310 0.06604 -0.0045 0.2203 1.0000
16.500 1.3429 0.07570 0.06867 -0.0044 0.2136 1.0000
16.750 1.3215 0.08131 0.07444 -0.0052 0.2085 1.0000
17.000 1.2838 0.08963 0.08295 -0.0069 0.2038 1.0000
17.250 1.2946 0.09067 0.08394 -0.0068 0.1957 1.0000
17.500 1.2385 0.10248 0.09599 -0.0101 0.1926 1.0000
17.750 0.9232 0.17238 0.16568 -0.0388 0.1941 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 510 AIRFOIL (goe510-il)