GOE 515 AIRFOIL (goe515-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 515 AIRFOIL (goe515-il) Reynolds number: 200,000 Max Cl/Cd: 71.04 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe515-il-200000-n5.txt Download as CSV file: xf-goe515-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 515 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.6407 0.05837 0.05502 -0.0506 1.0000 0.0205
-9.500 -0.7232 0.03651 0.03237 -0.0587 1.0000 0.0195
-9.250 -0.7290 0.03235 0.02769 -0.0566 1.0000 0.0199
-9.000 -0.7231 0.03044 0.02560 -0.0544 1.0000 0.0208
-8.750 -0.7102 0.02972 0.02483 -0.0525 1.0000 0.0218
-8.500 -0.6957 0.02874 0.02371 -0.0511 0.9996 0.0233
-8.250 -0.6685 0.02635 0.02086 -0.0528 0.9953 0.0256
-8.000 -0.6413 0.02433 0.01833 -0.0539 0.9909 0.0275
-7.750 -0.6119 0.02318 0.01712 -0.0553 0.9869 0.0298
-7.500 -0.5815 0.02234 0.01610 -0.0564 0.9830 0.0324
-7.250 -0.5519 0.02123 0.01464 -0.0574 0.9784 0.0349
-7.000 -0.5207 0.02000 0.01318 -0.0587 0.9751 0.0375
-6.750 -0.4920 0.01933 0.01242 -0.0593 0.9697 0.0400
-6.500 -0.4604 0.01862 0.01156 -0.0603 0.9657 0.0428
-6.250 -0.4270 0.01781 0.01052 -0.0617 0.9629 0.0451
-6.000 -0.3993 0.01701 0.00956 -0.0619 0.9572 0.0475
-5.750 -0.3682 0.01637 0.00886 -0.0628 0.9530 0.0503
-5.500 -0.3351 0.01578 0.00816 -0.0639 0.9499 0.0529
-5.250 -0.3061 0.01525 0.00751 -0.0642 0.9446 0.0556
-5.000 -0.2767 0.01472 0.00689 -0.0646 0.9393 0.0591
-4.750 -0.2442 0.01432 0.00650 -0.0656 0.9356 0.0647
-4.500 -0.2149 0.01400 0.00614 -0.0659 0.9302 0.0721
-4.250 -0.1853 0.01377 0.00586 -0.0662 0.9244 0.0807
-4.000 -0.1532 0.01350 0.00556 -0.0670 0.9203 0.0889
-3.750 -0.1251 0.01336 0.00529 -0.0670 0.9139 0.0965
-3.500 -0.0959 0.01308 0.00503 -0.0672 0.9082 0.1033
-3.250 -0.0647 0.01288 0.00474 -0.0678 0.9038 0.1098
-3.000 -0.0385 0.01265 0.00453 -0.0675 0.8964 0.1159
-2.750 -0.0084 0.01249 0.00433 -0.0678 0.8913 0.1250
-2.500 0.0185 0.01228 0.00416 -0.0676 0.8844 0.1336
-2.250 0.0469 0.01211 0.00394 -0.0676 0.8783 0.1404
-2.000 0.0745 0.01189 0.00376 -0.0675 0.8722 0.1467
-1.750 0.1016 0.01172 0.00358 -0.0672 0.8651 0.1535
-1.500 0.1296 0.01154 0.00342 -0.0671 0.8591 0.1612
-1.250 0.1559 0.01140 0.00330 -0.0666 0.8513 0.1722
-1.000 0.1827 0.01122 0.00317 -0.0662 0.8418 0.1892
-0.750 0.2092 0.01102 0.00302 -0.0657 0.8298 0.2138
-0.500 0.2354 0.01083 0.00289 -0.0651 0.8177 0.2446
-0.250 0.2602 0.01065 0.00284 -0.0644 0.8071 0.2840
0.000 0.2850 0.01039 0.00279 -0.0637 0.7985 0.3478
0.250 0.3068 0.00988 0.00277 -0.0625 0.7897 0.4980
0.500 0.3999 0.00880 0.00294 -0.0757 0.7837 0.9789
0.750 0.4479 0.00883 0.00291 -0.0799 0.7746 1.0000
1.000 0.4720 0.00888 0.00292 -0.0790 0.7640 1.0000
1.250 0.4963 0.00893 0.00293 -0.0781 0.7534 1.0000
1.500 0.5206 0.00899 0.00294 -0.0772 0.7417 1.0000
1.750 0.5443 0.00906 0.00295 -0.0762 0.7270 1.0000
2.000 0.5677 0.00913 0.00299 -0.0750 0.7093 1.0000
2.250 0.5909 0.00921 0.00300 -0.0738 0.6881 1.0000
2.500 0.6140 0.00932 0.00302 -0.0726 0.6663 1.0000
2.750 0.6371 0.00945 0.00309 -0.0714 0.6449 1.0000
3.000 0.6602 0.00961 0.00318 -0.0702 0.6238 1.0000
3.250 0.6829 0.00979 0.00329 -0.0690 0.6001 1.0000
3.500 0.7052 0.00999 0.00342 -0.0678 0.5745 1.0000
3.750 0.7267 0.01023 0.00358 -0.0663 0.5427 1.0000
4.000 0.7471 0.01054 0.00374 -0.0647 0.5049 1.0000
4.250 0.7665 0.01092 0.00395 -0.0629 0.4637 1.0000
4.500 0.7855 0.01134 0.00420 -0.0612 0.4253 1.0000
4.750 0.8040 0.01182 0.00452 -0.0594 0.3871 1.0000
5.000 0.8229 0.01229 0.00484 -0.0577 0.3494 1.0000
5.250 0.8418 0.01278 0.00519 -0.0560 0.3141 1.0000
5.500 0.8610 0.01325 0.00555 -0.0544 0.2763 1.0000
5.750 0.8782 0.01389 0.00598 -0.0526 0.2237 1.0000
6.000 0.8943 0.01464 0.00648 -0.0507 0.1771 1.0000
6.250 0.9118 0.01533 0.00703 -0.0489 0.1497 1.0000
6.750 0.9485 0.01655 0.00808 -0.0458 0.1052 1.0000
7.000 0.9672 0.01712 0.00862 -0.0443 0.0849 1.0000
7.250 0.9850 0.01778 0.00918 -0.0427 0.0651 1.0000
7.500 1.0020 0.01849 0.00985 -0.0410 0.0520 1.0000
7.750 1.0186 0.01925 0.01060 -0.0392 0.0412 1.0000
8.000 1.0346 0.02005 0.01140 -0.0373 0.0328 1.0000
8.250 1.0494 0.02094 0.01228 -0.0353 0.0281 1.0000
8.500 1.0654 0.02169 0.01317 -0.0334 0.0247 1.0000
8.750 1.0793 0.02260 0.01410 -0.0313 0.0218 1.0000
9.000 1.0917 0.02356 0.01515 -0.0290 0.0200 1.0000
9.250 1.1030 0.02448 0.01619 -0.0264 0.0190 1.0000
9.500 1.1128 0.02549 0.01731 -0.0237 0.0179 1.0000
9.750 1.1225 0.02654 0.01844 -0.0212 0.0169 1.0000
10.000 1.1317 0.02765 0.01963 -0.0188 0.0159 1.0000
10.250 1.1376 0.02908 0.02114 -0.0161 0.0151 1.0000
10.500 1.1443 0.03061 0.02277 -0.0137 0.0144 1.0000
10.750 1.1533 0.03199 0.02434 -0.0116 0.0139 1.0000
11.000 1.1613 0.03352 0.02602 -0.0096 0.0134 1.0000
11.250 1.1684 0.03523 0.02790 -0.0076 0.0130 1.0000
11.500 1.1746 0.03702 0.02986 -0.0057 0.0126 1.0000
11.750 1.1796 0.03897 0.03198 -0.0040 0.0123 1.0000
12.000 1.1834 0.04103 0.03422 -0.0023 0.0120 1.0000
12.250 1.1852 0.04324 0.03659 -0.0007 0.0117 1.0000
12.500 1.1863 0.04540 0.03890 0.0005 0.0113 1.0000
12.750 1.1852 0.04798 0.04165 0.0016 0.0111 1.0000
13.000 1.1830 0.05060 0.04440 0.0024 0.0108 1.0000
13.250 1.1796 0.05357 0.04753 0.0029 0.0107 1.0000
13.500 1.1706 0.05739 0.05151 0.0031 0.0104 1.0000
13.750 1.1592 0.06174 0.05604 0.0027 0.0102 1.0000
14.000 1.1480 0.06624 0.06078 0.0017 0.0101 1.0000
14.250 1.1356 0.07121 0.06598 0.0001 0.0100 1.0000
14.500 1.1214 0.07678 0.07175 -0.0021 0.0100 1.0000
14.750 1.1063 0.08293 0.07811 -0.0051 0.0100 1.0000
15.000 1.0893 0.08993 0.08531 -0.0088 0.0100 1.0000
15.250 1.0710 0.09775 0.09333 -0.0134 0.0100 1.0000
15.500 1.0529 0.10596 0.10169 -0.0182 0.0101 1.0000
15.750 1.0317 0.11541 0.11129 -0.0239 0.0102 1.0000
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Polar data table (+)
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