GOE 533 AIRFOIL (goe533-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 533 AIRFOIL (goe533-il) Reynolds number: 200,000 Max Cl/Cd: 75.5 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe533-il-200000-n5.txt Download as CSV file: xf-goe533-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 533 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.2755 0.11928 0.11548 -0.0380 1.0000 0.0387
-11.250 -0.2722 0.11645 0.11269 -0.0389 1.0000 0.0407
-11.000 -0.2887 0.11126 0.10754 -0.0438 1.0000 0.0424
-10.250 -0.2831 0.09969 0.09605 -0.0448 1.0000 0.0361
-10.000 -0.2791 0.09725 0.09365 -0.0444 1.0000 0.0356
-9.750 -0.2810 0.09402 0.09047 -0.0449 1.0000 0.0356
-9.500 -0.2756 0.08905 0.08553 -0.0488 0.9904 0.0357
-9.250 -0.2660 0.08354 0.08002 -0.0545 0.9777 0.0357
-9.000 -0.2528 0.07836 0.07482 -0.0603 0.9633 0.0353
-8.750 -0.2407 0.07203 0.06847 -0.0680 0.9447 0.0350
-8.500 -0.2315 0.06371 0.06010 -0.0790 0.9197 0.0347
-8.250 -0.2539 0.04686 0.04283 -0.1012 0.8778 0.0342
-8.000 -0.2578 0.03920 0.03456 -0.1071 0.8472 0.0347
-7.750 -0.2503 0.03447 0.02924 -0.1090 0.8207 0.0349
-7.250 -0.2188 0.02863 0.02236 -0.1095 0.7778 0.0355
-7.000 -0.1988 0.02662 0.01988 -0.1093 0.7601 0.0358
-6.750 -0.1770 0.02510 0.01783 -0.1090 0.7445 0.0364
-6.500 -0.1541 0.02375 0.01628 -0.1087 0.7306 0.0370
-6.250 -0.1299 0.02276 0.01511 -0.1084 0.7184 0.0375
-6.000 -0.1053 0.02180 0.01393 -0.1081 0.7078 0.0379
-5.750 -0.0801 0.02091 0.01284 -0.1078 0.6977 0.0382
-5.500 -0.0545 0.02012 0.01186 -0.1075 0.6891 0.0387
-5.250 -0.0286 0.01939 0.01097 -0.1071 0.6804 0.0392
-5.000 -0.0025 0.01873 0.01014 -0.1068 0.6730 0.0397
-4.750 0.0238 0.01812 0.00941 -0.1065 0.6651 0.0403
-4.500 0.0502 0.01763 0.00877 -0.1061 0.6581 0.0412
-4.250 0.0769 0.01719 0.00821 -0.1058 0.6511 0.0421
-4.000 0.1034 0.01670 0.00765 -0.1055 0.6441 0.0428
-3.750 0.1295 0.01619 0.00710 -0.1052 0.6379 0.0435
-3.500 0.1560 0.01576 0.00667 -0.1049 0.6311 0.0443
-3.250 0.1825 0.01542 0.00628 -0.1046 0.6246 0.0452
-3.000 0.2093 0.01512 0.00593 -0.1044 0.6187 0.0463
-2.750 0.2363 0.01483 0.00562 -0.1042 0.6122 0.0475
-2.500 0.2634 0.01462 0.00533 -0.1039 0.6062 0.0490
-2.250 0.2905 0.01437 0.00504 -0.1038 0.6007 0.0513
-2.000 0.3178 0.01415 0.00482 -0.1036 0.5943 0.0539
-1.750 0.3451 0.01400 0.00459 -0.1034 0.5881 0.0569
-1.500 0.3725 0.01384 0.00440 -0.1032 0.5823 0.0615
-1.250 0.3999 0.01364 0.00428 -0.1031 0.5758 0.0731
-1.000 0.4270 0.01337 0.00416 -0.1030 0.5698 0.1256
-0.750 0.4529 0.01272 0.00407 -0.1031 0.5638 0.2920
-0.500 0.4784 0.01233 0.00419 -0.1028 0.5568 0.4530
-0.250 0.5048 0.01227 0.00422 -0.1023 0.5506 0.5154
0.000 0.5308 0.01219 0.00433 -0.1018 0.5438 0.5711
0.250 0.5560 0.01218 0.00449 -0.1009 0.5370 0.6348
0.500 0.5818 0.01223 0.00456 -0.1003 0.5309 0.6763
0.750 0.6085 0.01225 0.00460 -0.0998 0.5232 0.6956
1.000 0.6349 0.01229 0.00459 -0.0994 0.5163 0.7121
1.250 0.6613 0.01231 0.00464 -0.0989 0.5085 0.7301
1.500 0.6872 0.01232 0.00465 -0.0984 0.5009 0.7487
1.750 0.7128 0.01233 0.00469 -0.0978 0.4927 0.7685
2.000 0.7379 0.01235 0.00471 -0.0970 0.4843 0.7906
2.250 0.7636 0.01235 0.00476 -0.0964 0.4757 0.8192
2.500 0.7910 0.01230 0.00478 -0.0960 0.4671 0.8792
2.750 0.8263 0.01237 0.00484 -0.0976 0.4581 1.0000
3.000 0.8526 0.01256 0.00494 -0.0974 0.4504 1.0000
3.250 0.8794 0.01275 0.00508 -0.0972 0.4432 1.0000
3.500 0.9056 0.01296 0.00521 -0.0969 0.4359 1.0000
3.750 0.9317 0.01317 0.00537 -0.0966 0.4290 1.0000
4.000 0.9577 0.01339 0.00555 -0.0963 0.4219 1.0000
4.250 0.9832 0.01364 0.00572 -0.0959 0.4158 1.0000
4.500 1.0093 0.01385 0.00594 -0.0957 0.4092 1.0000
4.750 1.0347 0.01410 0.00615 -0.0953 0.4032 1.0000
5.000 1.0599 0.01436 0.00638 -0.0949 0.3973 1.0000
5.250 1.0851 0.01460 0.00663 -0.0944 0.3905 1.0000
5.500 1.1095 0.01489 0.00686 -0.0939 0.3842 1.0000
5.750 1.1344 0.01514 0.00714 -0.0935 0.3778 1.0000
6.000 1.1586 0.01542 0.00742 -0.0929 0.3713 1.0000
6.250 1.1824 0.01573 0.00770 -0.0923 0.3660 1.0000
6.500 1.2069 0.01600 0.00802 -0.0918 0.3604 1.0000
6.750 1.2306 0.01630 0.00835 -0.0912 0.3549 1.0000
7.000 1.2533 0.01665 0.00867 -0.0905 0.3495 1.0000
7.250 1.2766 0.01693 0.00902 -0.0898 0.3426 1.0000
7.500 1.2981 0.01729 0.00937 -0.0889 0.3352 1.0000
7.750 1.3197 0.01762 0.00975 -0.0880 0.3279 1.0000
8.000 1.3402 0.01799 0.01014 -0.0869 0.3200 1.0000
8.250 1.3604 0.01838 0.01055 -0.0858 0.3125 1.0000
8.500 1.3800 0.01877 0.01099 -0.0846 0.3050 1.0000
8.750 1.3985 0.01920 0.01144 -0.0833 0.2981 1.0000
9.000 1.4165 0.01962 0.01192 -0.0819 0.2897 1.0000
9.250 1.4313 0.02010 0.01242 -0.0800 0.2814 1.0000
9.500 1.4446 0.02061 0.01296 -0.0779 0.2718 1.0000
9.750 1.4584 0.02115 0.01355 -0.0759 0.2635 1.0000
10.000 1.4697 0.02179 0.01421 -0.0737 0.2551 1.0000
10.250 1.4822 0.02244 0.01492 -0.0718 0.2463 1.0000
10.500 1.4916 0.02324 0.01575 -0.0696 0.2377 1.0000
10.750 1.5003 0.02412 0.01666 -0.0675 0.2282 1.0000
11.000 1.5080 0.02513 0.01769 -0.0654 0.2192 1.0000
11.250 1.5126 0.02636 0.01893 -0.0633 0.2100 1.0000
11.500 1.5171 0.02770 0.02029 -0.0613 0.2013 1.0000
11.750 1.5215 0.02916 0.02179 -0.0596 0.1950 1.0000
12.000 1.5267 0.03066 0.02336 -0.0581 0.1894 1.0000
12.250 1.5285 0.03250 0.02523 -0.0567 0.1841 1.0000
12.500 1.5333 0.03422 0.02703 -0.0556 0.1791 1.0000
12.750 1.5355 0.03624 0.02911 -0.0545 0.1734 1.0000
13.000 1.5333 0.03873 0.03163 -0.0536 0.1679 1.0000
13.250 1.5362 0.04089 0.03389 -0.0529 0.1609 1.0000
13.500 1.5311 0.04389 0.03691 -0.0524 0.1538 1.0000
13.750 1.5334 0.04626 0.03940 -0.0520 0.1460 1.0000
14.000 1.5289 0.04940 0.04258 -0.0517 0.1378 1.0000
14.250 1.5262 0.05248 0.04573 -0.0517 0.1247 1.0000
14.500 1.5165 0.05644 0.04968 -0.0518 0.1051 1.0000
15.000 1.4853 0.06635 0.05950 -0.0531 0.0855 1.0000
15.250 1.4728 0.07121 0.06438 -0.0540 0.0807 1.0000
15.500 1.4638 0.07574 0.06900 -0.0549 0.0768 1.0000
15.750 1.4540 0.08052 0.07385 -0.0560 0.0738 1.0000
16.000 1.4434 0.08556 0.07897 -0.0574 0.0713 1.0000
16.250 1.4388 0.08981 0.08333 -0.0585 0.0690 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 533 AIRFOIL (goe533-il)