GOE 676 (= M 12) AIRFOIL (goe676-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 676 (= M 12) AIRFOIL (goe676-il) Reynolds number: 200,000 Max Cl/Cd: 62.77 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe676-il-200000-n5.txt Download as CSV file: xf-goe676-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 676 (= M 12) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.7667 0.05268 0.04864 -0.0486 1.0000 0.0293
-10.500 -0.8104 0.04539 0.04085 -0.0457 1.0000 0.0298
-10.250 -0.8344 0.03929 0.03405 -0.0424 1.0000 0.0308
-10.000 -0.8335 0.03627 0.03065 -0.0401 1.0000 0.0316
-9.750 -0.8172 0.03554 0.02990 -0.0390 1.0000 0.0322
-9.500 -0.8012 0.03482 0.02913 -0.0378 1.0000 0.0330
-9.250 -0.7876 0.03371 0.02789 -0.0361 1.0000 0.0339
-9.000 -0.7765 0.03218 0.02614 -0.0341 1.0000 0.0348
-8.750 -0.7660 0.03058 0.02428 -0.0317 1.0000 0.0359
-8.500 -0.7555 0.02900 0.02239 -0.0292 1.0000 0.0371
-8.250 -0.7441 0.02766 0.02072 -0.0267 1.0000 0.0380
-8.000 -0.7199 0.02603 0.01890 -0.0269 0.9966 0.0390
-7.750 -0.6884 0.02503 0.01783 -0.0284 0.9922 0.0400
-7.500 -0.6565 0.02423 0.01692 -0.0298 0.9872 0.0413
-7.250 -0.6238 0.02324 0.01576 -0.0313 0.9830 0.0427
-7.000 -0.5931 0.02215 0.01447 -0.0323 0.9770 0.0439
-6.750 -0.5601 0.02111 0.01325 -0.0336 0.9725 0.0451
-6.500 -0.5301 0.02030 0.01226 -0.0343 0.9651 0.0462
-6.250 -0.4979 0.01933 0.01123 -0.0355 0.9599 0.0478
-6.000 -0.4694 0.01866 0.01054 -0.0359 0.9515 0.0491
-5.750 -0.4377 0.01801 0.00985 -0.0368 0.9454 0.0506
-5.500 -0.4104 0.01743 0.00921 -0.0368 0.9356 0.0521
-5.250 -0.3814 0.01686 0.00857 -0.0371 0.9277 0.0538
-5.000 -0.3543 0.01635 0.00799 -0.0369 0.9181 0.0556
-4.750 -0.3278 0.01582 0.00742 -0.0367 0.9085 0.0578
-4.500 -0.3013 0.01534 0.00695 -0.0364 0.8993 0.0607
-4.250 -0.2757 0.01497 0.00653 -0.0359 0.8887 0.0639
-4.000 -0.2497 0.01461 0.00610 -0.0355 0.8788 0.0671
-3.750 -0.2243 0.01419 0.00565 -0.0349 0.8688 0.0708
-3.500 -0.1992 0.01383 0.00527 -0.0343 0.8578 0.0756
-3.250 -0.1734 0.01357 0.00493 -0.0337 0.8472 0.0812
-3.000 -0.1482 0.01323 0.00461 -0.0331 0.8365 0.0893
-2.750 -0.1228 0.01296 0.00435 -0.0325 0.8245 0.1001
-2.500 -0.0973 0.01270 0.00415 -0.0320 0.8124 0.1194
-2.250 -0.0717 0.01244 0.00394 -0.0315 0.8005 0.1489
-2.000 -0.0465 0.01214 0.00372 -0.0310 0.7883 0.1831
-1.750 -0.0229 0.01162 0.00350 -0.0303 0.7753 0.2609
-1.500 -0.0025 0.01082 0.00332 -0.0291 0.7620 0.4304
-1.250 0.0175 0.01015 0.00320 -0.0276 0.7491 0.5727
-1.000 0.0374 0.00962 0.00327 -0.0253 0.7366 0.7282
-0.750 0.0638 0.00960 0.00351 -0.0239 0.7244 0.8360
-0.250 0.1218 0.00985 0.00367 -0.0231 0.6991 0.9047
0.000 0.1523 0.00999 0.00373 -0.0231 0.6867 0.9253
0.250 0.1862 0.01015 0.00379 -0.0239 0.6745 0.9415
0.500 0.2240 0.01030 0.00384 -0.0257 0.6624 0.9523
0.750 0.2573 0.01041 0.00384 -0.0267 0.6507 0.9605
1.000 0.2982 0.01052 0.00387 -0.0294 0.6385 0.9666
1.250 0.3375 0.01064 0.00390 -0.0317 0.6262 0.9744
1.500 0.3781 0.01075 0.00392 -0.0344 0.6120 0.9814
1.750 0.4175 0.01084 0.00393 -0.0370 0.5972 0.9870
2.000 0.4534 0.01091 0.00393 -0.0388 0.5823 0.9913
2.250 0.4863 0.01099 0.00395 -0.0400 0.5685 0.9945
2.500 0.5198 0.01105 0.00396 -0.0414 0.5545 0.9971
2.750 0.5529 0.01112 0.00399 -0.0427 0.5371 1.0000
3.000 0.5770 0.01121 0.00402 -0.0421 0.5185 1.0000
3.250 0.6006 0.01131 0.00407 -0.0414 0.5003 1.0000
3.500 0.6241 0.01143 0.00415 -0.0407 0.4840 1.0000
3.750 0.6474 0.01157 0.00424 -0.0400 0.4689 1.0000
4.000 0.6704 0.01173 0.00435 -0.0392 0.4536 1.0000
4.250 0.6929 0.01191 0.00449 -0.0383 0.4371 1.0000
4.500 0.7153 0.01211 0.00464 -0.0373 0.4217 1.0000
4.750 0.7376 0.01231 0.00482 -0.0364 0.4090 1.0000
5.000 0.7597 0.01253 0.00502 -0.0354 0.3963 1.0000
5.250 0.7813 0.01277 0.00523 -0.0344 0.3821 1.0000
5.500 0.8027 0.01302 0.00546 -0.0333 0.3673 1.0000
5.750 0.8238 0.01329 0.00570 -0.0321 0.3522 1.0000
6.000 0.8451 0.01355 0.00597 -0.0310 0.3393 1.0000
6.250 0.8663 0.01382 0.00625 -0.0299 0.3258 1.0000
6.500 0.8869 0.01413 0.00654 -0.0286 0.3093 1.0000
6.750 0.9068 0.01448 0.00685 -0.0273 0.2878 1.0000
7.000 0.9254 0.01491 0.00721 -0.0259 0.2583 1.0000
7.250 0.9425 0.01548 0.00761 -0.0242 0.2187 1.0000
7.500 0.9578 0.01623 0.00813 -0.0224 0.1760 1.0000
7.750 0.9737 0.01698 0.00872 -0.0207 0.1498 1.0000
8.000 0.9909 0.01765 0.00934 -0.0192 0.1356 1.0000
8.250 1.0087 0.01827 0.00995 -0.0178 0.1250 1.0000
8.500 1.0257 0.01894 0.01060 -0.0164 0.1154 1.0000
8.750 1.0442 0.01950 0.01121 -0.0151 0.1061 1.0000
9.000 1.0624 0.02009 0.01184 -0.0138 0.0964 1.0000
9.250 1.0802 0.02068 0.01247 -0.0126 0.0837 1.0000
9.500 1.0960 0.02141 0.01313 -0.0111 0.0658 1.0000
9.750 1.1084 0.02234 0.01396 -0.0092 0.0529 1.0000
10.000 1.1185 0.02332 0.01491 -0.0070 0.0455 1.0000
10.250 1.1280 0.02426 0.01590 -0.0046 0.0409 1.0000
10.500 1.1354 0.02540 0.01706 -0.0023 0.0372 1.0000
10.750 1.1442 0.02652 0.01828 -0.0003 0.0345 1.0000
11.000 1.1527 0.02772 0.01958 0.0014 0.0322 1.0000
11.250 1.1583 0.02921 0.02112 0.0032 0.0301 1.0000
11.500 1.1644 0.03074 0.02275 0.0047 0.0285 1.0000
11.750 1.1717 0.03225 0.02438 0.0059 0.0268 1.0000
12.000 1.1774 0.03397 0.02620 0.0070 0.0255 1.0000
12.250 1.1812 0.03593 0.02825 0.0080 0.0245 1.0000
12.500 1.1813 0.03833 0.03073 0.0088 0.0236 1.0000
12.750 1.1849 0.04049 0.03302 0.0094 0.0228 1.0000
13.000 1.1886 0.04272 0.03539 0.0098 0.0219 1.0000
13.250 1.1912 0.04513 0.03792 0.0100 0.0210 1.0000
13.500 1.1933 0.04768 0.04057 0.0101 0.0202 1.0000
13.750 1.1941 0.05044 0.04343 0.0100 0.0196 1.0000
14.000 1.1934 0.05342 0.04650 0.0097 0.0191 1.0000
14.250 1.1904 0.05672 0.04988 0.0093 0.0186 1.0000
14.500 1.1879 0.06001 0.05326 0.0090 0.0183 1.0000
14.750 1.1869 0.06322 0.05664 0.0085 0.0179 1.0000
15.000 1.1851 0.06661 0.06018 0.0079 0.0175 1.0000
15.250 1.1828 0.07018 0.06389 0.0072 0.0171 1.0000
15.500 1.1799 0.07392 0.06776 0.0062 0.0167 1.0000
15.750 1.1763 0.07784 0.07182 0.0050 0.0164 1.0000
16.000 1.1723 0.08194 0.07605 0.0037 0.0160 1.0000
16.250 1.1678 0.08624 0.08047 0.0021 0.0157 1.0000
16.500 1.1626 0.09071 0.08506 0.0003 0.0154 1.0000
16.750 1.1569 0.09536 0.08982 -0.0016 0.0152 1.0000
17.000 1.1509 0.10015 0.09471 -0.0037 0.0149 1.0000
17.250 1.1446 0.10503 0.09969 -0.0059 0.0147 1.0000
17.500 1.1379 0.11003 0.10478 -0.0081 0.0145 1.0000
18.000 1.1203 0.12122 0.11624 -0.0135 0.0142 1.0000
18.250 1.1091 0.12766 0.12285 -0.0169 0.0141 1.0000
18.500 1.0965 0.13458 0.12994 -0.0207 0.0140 1.0000
18.750 1.0824 0.14207 0.13761 -0.0250 0.0139 1.0000
19.000 1.0661 0.15046 0.14617 -0.0300 0.0139 1.0000
19.250 1.0465 0.16020 0.15609 -0.0359 0.0139 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 676 (= M 12) AIRFOIL (goe676-il)