GOE 496 AIRFOIL (goe496-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 496 AIRFOIL (goe496-il) Reynolds number: 100,000 Max Cl/Cd: 62.88 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe496-il-100000.txt Download as CSV file: xf-goe496-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 496 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 -0.3356 0.10455 0.09977 -0.0329 1.0000 0.1183
-8.000 -0.3666 0.10533 0.10072 -0.0308 1.0000 0.1190
-7.750 -0.3968 0.10580 0.10134 -0.0287 1.0000 0.1193
-7.500 -0.4053 0.10271 0.09834 -0.0274 1.0000 0.1201
-7.250 -0.3646 0.09658 0.09211 -0.0234 1.0000 0.1234
-7.000 -0.3726 0.09520 0.09080 -0.0204 1.0000 0.1251
-6.750 -0.3859 0.09416 0.08982 -0.0174 1.0000 0.1269
-6.500 -0.3981 0.09285 0.08860 -0.0156 1.0000 0.1292
-6.250 -0.4124 0.09140 0.08723 -0.0162 1.0000 0.1325
-6.000 -0.4268 0.08809 0.08397 -0.0289 1.0000 0.1356
-5.750 -0.4096 0.08467 0.08056 -0.0232 0.9968 0.1375
-5.500 -0.3827 0.08172 0.07757 -0.0241 0.9915 0.1430
-5.250 -0.3438 0.07519 0.07093 -0.0409 0.9827 0.1529
-5.000 -0.2983 0.06947 0.06498 -0.0582 0.9735 0.1669
-4.750 -0.2801 0.06657 0.06215 -0.0552 0.9681 0.1687
-4.500 -0.1919 0.04084 0.03479 -0.0929 0.9655 0.1057
-4.250 -0.1479 0.03631 0.02973 -0.0987 0.9607 0.1027
-4.000 -0.0978 0.03276 0.02533 -0.1049 0.9571 0.1042
-3.750 -0.0642 0.03032 0.02259 -0.1073 0.9500 0.1066
-3.500 -0.0221 0.02868 0.02069 -0.1106 0.9446 0.1096
-3.250 0.0153 0.02744 0.01913 -0.1127 0.9381 0.1142
-3.000 0.0542 0.02630 0.01771 -0.1151 0.9313 0.1225
-2.750 0.0980 0.02539 0.01660 -0.1181 0.9265 0.1334
-2.500 0.1283 0.02456 0.01578 -0.1187 0.9176 0.1467
-2.250 0.1730 0.02361 0.01488 -0.1219 0.9131 0.1829
-2.000 0.2012 0.02330 0.01470 -0.1222 0.9035 0.2156
-1.750 0.2446 0.02302 0.01448 -0.1250 0.8983 0.2481
-1.500 0.2722 0.02302 0.01453 -0.1251 0.8884 0.2725
-1.250 0.3162 0.02275 0.01431 -0.1280 0.8834 0.3094
-1.000 0.3434 0.02262 0.01431 -0.1279 0.8736 0.3389
-0.750 0.3870 0.02203 0.01394 -0.1305 0.8688 0.3917
-0.500 0.4153 0.02187 0.01397 -0.1304 0.8598 0.4490
-0.250 0.4569 0.02136 0.01364 -0.1324 0.8545 0.5163
0.000 0.4954 0.02090 0.01341 -0.1339 0.8490 0.5802
0.250 0.5241 0.02063 0.01339 -0.1336 0.8402 0.6526
0.500 0.5645 0.01936 0.01276 -0.1344 0.8365 1.0000
0.750 0.5913 0.01968 0.01287 -0.1345 0.8256 1.0000
1.000 0.6355 0.01939 0.01241 -0.1369 0.8211 1.0000
1.250 0.6596 0.01975 0.01266 -0.1363 0.8102 1.0000
1.500 0.7015 0.01946 0.01228 -0.1381 0.8054 1.0000
1.750 0.7243 0.01987 0.01262 -0.1372 0.7943 1.0000
2.000 0.7663 0.01949 0.01217 -0.1389 0.7892 1.0000
2.250 0.7906 0.01968 0.01232 -0.1379 0.7768 1.0000
2.500 0.8202 0.01956 0.01216 -0.1374 0.7651 1.0000
2.750 0.8555 0.01913 0.01167 -0.1376 0.7544 1.0000
3.000 0.8880 0.01888 0.01137 -0.1375 0.7435 1.0000
3.250 0.9124 0.01910 0.01160 -0.1365 0.7316 1.0000
3.500 0.9409 0.01917 0.01166 -0.1361 0.7214 1.0000
3.750 0.9728 0.01908 0.01155 -0.1361 0.7119 1.0000
4.000 0.9967 0.01932 0.01183 -0.1351 0.6996 1.0000
4.250 1.0231 0.01945 0.01198 -0.1344 0.6880 1.0000
4.500 1.0528 0.01943 0.01197 -0.1340 0.6771 1.0000
4.750 1.0813 0.01942 0.01197 -0.1335 0.6650 1.0000
5.000 1.1059 0.01957 0.01217 -0.1324 0.6511 1.0000
5.250 1.1312 0.01969 0.01235 -0.1314 0.6370 1.0000
5.500 1.1569 0.01976 0.01246 -0.1305 0.6219 1.0000
5.750 1.1825 0.01982 0.01254 -0.1294 0.6058 1.0000
6.000 1.2083 0.01987 0.01263 -0.1284 0.5886 1.0000
6.250 1.2330 0.01998 0.01273 -0.1272 0.5694 1.0000
6.500 1.2541 0.02018 0.01299 -0.1254 0.5467 1.0000
6.750 1.2780 0.02035 0.01310 -0.1240 0.5232 1.0000
7.000 1.2972 0.02063 0.01337 -0.1220 0.4950 1.0000
7.250 1.3141 0.02091 0.01358 -0.1195 0.4634 1.0000
7.500 1.3292 0.02130 0.01385 -0.1169 0.4311 1.0000
7.750 1.3434 0.02183 0.01429 -0.1143 0.4004 1.0000
8.000 1.3573 0.02248 0.01489 -0.1118 0.3720 1.0000
8.250 1.3700 0.02321 0.01555 -0.1092 0.3447 1.0000
8.500 1.3815 0.02402 0.01628 -0.1066 0.3189 1.0000
8.750 1.3913 0.02493 0.01707 -0.1037 0.2948 1.0000
9.000 1.3995 0.02587 0.01802 -0.1008 0.2702 1.0000
9.250 1.4036 0.02693 0.01902 -0.0972 0.2467 1.0000
9.500 1.4030 0.02817 0.02020 -0.0931 0.2197 1.0000
9.750 1.3978 0.02972 0.02171 -0.0888 0.1827 1.0000
10.000 1.3870 0.03198 0.02369 -0.0843 0.1387 1.0000
10.250 1.3784 0.03459 0.02599 -0.0806 0.1108 1.0000
10.500 1.3733 0.03718 0.02840 -0.0775 0.0959 1.0000
10.750 1.3726 0.03967 0.03086 -0.0749 0.0858 1.0000
11.000 1.3743 0.04218 0.03331 -0.0727 0.0786 1.0000
11.250 1.3798 0.04443 0.03556 -0.0709 0.0723 1.0000
11.500 1.3927 0.04672 0.03784 -0.0693 0.0674 1.0000
11.750 1.4085 0.04875 0.03998 -0.0680 0.0633 1.0000
12.000 1.4412 0.05122 0.04223 -0.0680 0.0584 1.0000
12.250 1.4498 0.05347 0.04480 -0.0664 0.0562 1.0000
12.500 1.4651 0.05602 0.04758 -0.0653 0.0540 1.0000
12.750 1.4805 0.05882 0.05062 -0.0643 0.0524 1.0000
13.000 1.4921 0.06185 0.05384 -0.0633 0.0511 1.0000
13.250 1.5114 0.06560 0.05766 -0.0630 0.0493 1.0000
13.500 1.5112 0.07023 0.06257 -0.0619 0.0484 1.0000
13.750 1.4967 0.07379 0.06647 -0.0600 0.0481 1.0000
14.000 1.4810 0.07782 0.07083 -0.0588 0.0479 1.0000
14.250 1.4642 0.08233 0.07566 -0.0581 0.0477 1.0000
14.500 1.4461 0.08731 0.08094 -0.0581 0.0477 1.0000
14.750 1.4266 0.09279 0.08669 -0.0588 0.0477 1.0000
15.000 1.4062 0.09871 0.09288 -0.0601 0.0478 1.0000
15.250 1.3852 0.10509 0.09950 -0.0622 0.0480 1.0000
15.500 1.3639 0.11195 0.10658 -0.0649 0.0482 1.0000
15.750 1.3438 0.11925 0.11407 -0.0683 0.0485 1.0000
16.000 1.3186 0.12686 0.12194 -0.0732 0.0489 1.0000
16.250 1.1653 0.17221 0.16783 -0.1076 0.0599 1.0000
16.500 1.1755 0.17364 0.16927 -0.1071 0.0588 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 496 AIRFOIL (goe496-il)