GOE 675 AIRFOIL (goe675-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 675 AIRFOIL (goe675-il) Reynolds number: 1,000,000 Max Cl/Cd: 117.81 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe675-il-1000000.txt Download as CSV file: xf-goe675-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 675 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.000 -0.6613 0.06696 0.06432 -0.1048 0.9571 0.0288
-14.750 -0.7286 0.05393 0.05103 -0.1145 0.9394 0.0287
-14.500 -0.7835 0.04056 0.03738 -0.1274 0.9243 0.0286
-14.250 -0.8010 0.03097 0.02742 -0.1412 0.9132 0.0287
-14.000 -0.7920 0.02895 0.02521 -0.1422 0.9071 0.0288
-13.750 -0.7782 0.02744 0.02355 -0.1427 0.9019 0.0289
-13.500 -0.7713 0.02500 0.02094 -0.1429 0.8970 0.0291
-13.250 -0.7571 0.02338 0.01919 -0.1428 0.8925 0.0294
-13.000 -0.7389 0.02220 0.01790 -0.1426 0.8883 0.0296
-12.750 -0.7184 0.02125 0.01686 -0.1425 0.8841 0.0299
-12.500 -0.6963 0.02039 0.01594 -0.1424 0.8810 0.0301
-12.250 -0.6735 0.01961 0.01509 -0.1422 0.8774 0.0304
-12.000 -0.6501 0.01890 0.01429 -0.1421 0.8739 0.0306
-11.750 -0.6261 0.01824 0.01355 -0.1419 0.8704 0.0309
-11.500 -0.6017 0.01764 0.01285 -0.1418 0.8664 0.0312
-11.250 -0.5766 0.01702 0.01217 -0.1417 0.8637 0.0315
-11.000 -0.5511 0.01646 0.01153 -0.1416 0.8606 0.0318
-10.750 -0.5253 0.01593 0.01093 -0.1415 0.8573 0.0321
-10.500 -0.4992 0.01545 0.01037 -0.1414 0.8541 0.0324
-10.250 -0.4726 0.01506 0.00989 -0.1413 0.8508 0.0326
-10.000 -0.4470 0.01447 0.00921 -0.1412 0.8474 0.0330
-9.750 -0.4212 0.01383 0.00854 -0.1411 0.8447 0.0336
-9.500 -0.3945 0.01336 0.00802 -0.1410 0.8415 0.0341
-9.250 -0.3673 0.01295 0.00757 -0.1410 0.8382 0.0345
-9.000 -0.3399 0.01259 0.00716 -0.1410 0.8350 0.0351
-8.750 -0.3122 0.01228 0.00677 -0.1409 0.8318 0.0356
-8.500 -0.2843 0.01198 0.00642 -0.1410 0.8285 0.0362
-8.250 -0.2562 0.01168 0.00608 -0.1410 0.8249 0.0367
-8.000 -0.2286 0.01130 0.00566 -0.1409 0.8207 0.0375
-7.750 -0.2008 0.01095 0.00528 -0.1409 0.8166 0.0385
-7.500 -0.1727 0.01070 0.00497 -0.1409 0.8123 0.0396
-7.250 -0.1443 0.01045 0.00469 -0.1409 0.8088 0.0408
-7.000 -0.1159 0.01016 0.00440 -0.1410 0.8050 0.0425
-6.750 -0.0875 0.00990 0.00414 -0.1410 0.8010 0.0452
-6.500 -0.0591 0.00967 0.00390 -0.1411 0.7966 0.0494
-6.250 -0.0303 0.00952 0.00377 -0.1411 0.7924 0.0559
-6.000 -0.0010 0.00943 0.00372 -0.1413 0.7887 0.0617
-5.750 0.0282 0.00937 0.00366 -0.1414 0.7844 0.0656
-5.500 0.0574 0.00935 0.00362 -0.1415 0.7795 0.0686
-5.250 0.0867 0.00937 0.00359 -0.1416 0.7747 0.0707
-5.000 0.1156 0.00920 0.00344 -0.1417 0.7702 0.0731
-4.750 0.1446 0.00914 0.00338 -0.1418 0.7650 0.0751
-4.500 0.1736 0.00912 0.00331 -0.1419 0.7598 0.0770
-4.250 0.2029 0.00909 0.00326 -0.1420 0.7552 0.0784
-4.000 0.2316 0.00892 0.00308 -0.1421 0.7499 0.0803
-3.750 0.2602 0.00881 0.00296 -0.1422 0.7447 0.0822
-3.500 0.2890 0.00875 0.00288 -0.1422 0.7398 0.0838
-3.250 0.3181 0.00867 0.00281 -0.1424 0.7345 0.0854
-3.000 0.3470 0.00862 0.00273 -0.1424 0.7285 0.0869
-2.750 0.3756 0.00861 0.00267 -0.1425 0.7224 0.0878
-2.500 0.4042 0.00842 0.00250 -0.1426 0.7164 0.0905
-2.250 0.4327 0.00835 0.00242 -0.1426 0.7099 0.0926
-2.000 0.4612 0.00831 0.00236 -0.1426 0.7034 0.0944
-1.750 0.4899 0.00826 0.00231 -0.1427 0.6958 0.0964
-1.500 0.5179 0.00827 0.00225 -0.1426 0.6878 0.0976
-1.250 0.5465 0.00817 0.00216 -0.1427 0.6800 0.0999
-1.000 0.5742 0.00814 0.00211 -0.1425 0.6714 0.1029
-0.750 0.6026 0.00811 0.00208 -0.1426 0.6624 0.1057
-0.500 0.6300 0.00814 0.00206 -0.1424 0.6521 0.1083
-0.250 0.6578 0.00812 0.00203 -0.1423 0.6409 0.1134
0.000 0.6851 0.00813 0.00204 -0.1421 0.6298 0.1204
0.250 0.7118 0.00816 0.00206 -0.1418 0.6169 0.1325
0.500 0.7384 0.00820 0.00209 -0.1416 0.6020 0.1495
0.750 0.7645 0.00825 0.00214 -0.1412 0.5851 0.1703
1.000 0.7896 0.00830 0.00222 -0.1407 0.5659 0.2084
1.250 0.8145 0.00836 0.00234 -0.1402 0.5447 0.2619
1.500 0.8386 0.00851 0.00247 -0.1395 0.5199 0.3001
1.750 0.8615 0.00871 0.00263 -0.1386 0.4947 0.3374
2.000 0.8850 0.00882 0.00280 -0.1379 0.4738 0.3987
2.250 0.9085 0.00878 0.00301 -0.1373 0.4579 0.5342
2.500 0.9330 0.00806 0.00330 -0.1365 0.4462 1.0000
2.750 0.9577 0.00826 0.00344 -0.1359 0.4366 1.0000
3.000 0.9820 0.00847 0.00358 -0.1353 0.4278 1.0000
3.250 1.0068 0.00866 0.00372 -0.1347 0.4206 1.0000
3.500 1.0319 0.00883 0.00385 -0.1342 0.4144 1.0000
3.750 1.0556 0.00906 0.00402 -0.1334 0.4079 1.0000
4.000 1.0807 0.00922 0.00416 -0.1329 0.4029 1.0000
4.250 1.1056 0.00939 0.00430 -0.1324 0.3975 1.0000
4.500 1.1283 0.00963 0.00449 -0.1315 0.3907 1.0000
4.750 1.1523 0.00981 0.00465 -0.1308 0.3846 1.0000
5.000 1.1757 0.00998 0.00480 -0.1300 0.3784 1.0000
5.250 1.1961 0.01024 0.00501 -0.1286 0.3721 1.0000
5.500 1.2189 0.01043 0.00519 -0.1277 0.3675 1.0000
5.750 1.2423 0.01060 0.00537 -0.1269 0.3634 1.0000
6.000 1.2643 0.01084 0.00558 -0.1259 0.3586 1.0000
6.250 1.2844 0.01116 0.00586 -0.1246 0.3525 1.0000
6.500 1.3084 0.01133 0.00605 -0.1241 0.3483 1.0000
6.750 1.3307 0.01157 0.00629 -0.1232 0.3434 1.0000
7.000 1.3512 0.01189 0.00658 -0.1221 0.3376 1.0000
7.250 1.3732 0.01216 0.00685 -0.1212 0.3320 1.0000
7.500 1.3948 0.01245 0.00713 -0.1203 0.3256 1.0000
7.750 1.4137 0.01287 0.00750 -0.1190 0.3183 1.0000
8.000 1.4361 0.01313 0.00778 -0.1183 0.3129 1.0000
8.250 1.4557 0.01354 0.00816 -0.1172 0.3054 1.0000
8.500 1.4755 0.01393 0.00855 -0.1161 0.2990 1.0000
8.750 1.4951 0.01435 0.00895 -0.1150 0.2913 1.0000
9.000 1.5133 0.01484 0.00941 -0.1138 0.2830 1.0000
9.250 1.5307 0.01539 0.00992 -0.1124 0.2730 1.0000
9.500 1.5479 0.01595 0.01045 -0.1111 0.2621 1.0000
9.750 1.5628 0.01666 0.01110 -0.1095 0.2498 1.0000
10.000 1.5765 0.01744 0.01181 -0.1078 0.2358 1.0000
10.250 1.5884 0.01834 0.01263 -0.1059 0.2199 1.0000
10.500 1.5990 0.01933 0.01353 -0.1039 0.2027 1.0000
10.750 1.6066 0.02053 0.01462 -0.1016 0.1840 1.0000
11.000 1.6129 0.02184 0.01582 -0.0992 0.1654 1.0000
11.250 1.6177 0.02328 0.01715 -0.0967 0.1459 1.0000
11.500 1.6164 0.02515 0.01886 -0.0937 0.1206 1.0000
11.750 1.6083 0.02758 0.02108 -0.0901 0.0905 1.0000
12.000 1.6014 0.03003 0.02340 -0.0869 0.0684 1.0000
12.250 1.5915 0.03286 0.02607 -0.0838 0.0437 1.0000
12.500 1.5949 0.03476 0.02796 -0.0819 0.0384 1.0000
12.750 1.6016 0.03642 0.02965 -0.0805 0.0362 1.0000
13.000 1.6084 0.03811 0.03138 -0.0792 0.0350 1.0000
13.250 1.6146 0.03991 0.03321 -0.0779 0.0339 1.0000
13.500 1.6191 0.04192 0.03527 -0.0766 0.0329 1.0000
13.750 1.6254 0.04380 0.03721 -0.0755 0.0324 1.0000
14.000 1.6319 0.04568 0.03915 -0.0745 0.0319 1.0000
14.250 1.6368 0.04777 0.04130 -0.0735 0.0315 1.0000
14.500 1.6412 0.04995 0.04354 -0.0726 0.0310 1.0000
14.750 1.6442 0.05233 0.04599 -0.0717 0.0305 1.0000
15.000 1.6452 0.05498 0.04871 -0.0709 0.0300 1.0000
15.250 1.6445 0.05785 0.05164 -0.0701 0.0296 1.0000
15.500 1.6423 0.06096 0.05483 -0.0694 0.0292 1.0000
15.750 1.6377 0.06444 0.05840 -0.0688 0.0288 1.0000
16.000 1.6314 0.06820 0.06225 -0.0683 0.0285 1.0000
16.250 1.6310 0.07127 0.06540 -0.0680 0.0283 1.0000
16.500 1.6289 0.07460 0.06882 -0.0678 0.0281 1.0000
16.750 1.6255 0.07817 0.07247 -0.0677 0.0279 1.0000
17.000 1.6212 0.08190 0.07629 -0.0678 0.0277 1.0000
17.250 1.6160 0.08581 0.08029 -0.0680 0.0275 1.0000
17.500 1.6101 0.08983 0.08441 -0.0682 0.0272 1.0000
17.750 1.6032 0.09402 0.08868 -0.0686 0.0270 1.0000
18.000 1.5961 0.09830 0.09305 -0.0691 0.0267 1.0000
18.250 1.5886 0.10267 0.09750 -0.0698 0.0265 1.0000
18.500 1.5811 0.10707 0.10198 -0.0705 0.0263 1.0000
18.750 1.5733 0.11150 0.10650 -0.0714 0.0261 1.0000
19.000 1.5649 0.11606 0.11114 -0.0724 0.0260 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 675 AIRFOIL (goe675-il)