GOE 621 AIRFOIL (goe621-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 621 AIRFOIL (goe621-il) Reynolds number: 50,000 Max Cl/Cd: 27.65 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe621-il-50000-n5.txt Download as CSV file: xf-goe621-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 621 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.500 -0.2583 0.14747 0.13992 -0.0438 1.0000 0.1349
-12.250 -0.2679 0.14639 0.13893 -0.0438 1.0000 0.1357
-12.000 -0.2754 0.14488 0.13749 -0.0435 1.0000 0.1360
-11.750 -0.2824 0.14314 0.13584 -0.0430 1.0000 0.1362
-11.500 -0.2883 0.14119 0.13396 -0.0423 1.0000 0.1363
-11.250 -0.2926 0.13906 0.13190 -0.0413 1.0000 0.1363
-10.250 -0.3172 0.12357 0.11657 -0.0392 1.0000 0.0941
-10.000 -0.3190 0.12157 0.11463 -0.0377 0.9997 0.0934
-9.750 -0.3019 0.11733 0.11037 -0.0413 0.9941 0.0920
-9.500 -0.2894 0.11303 0.10605 -0.0451 0.9879 0.0903
-9.250 -0.2791 0.10843 0.10144 -0.0496 0.9820 0.0887
-9.000 -0.2736 0.10375 0.09675 -0.0538 0.9746 0.0875
-8.750 -0.2664 0.09901 0.09200 -0.0587 0.9683 0.0871
-8.500 -0.2616 0.09498 0.08797 -0.0622 0.9597 0.0875
-8.250 -0.2580 0.09077 0.08377 -0.0661 0.9517 0.0881
-8.000 -0.2594 0.08602 0.07903 -0.0705 0.9428 0.0879
-7.750 -0.2714 0.08134 0.07436 -0.0737 0.9311 0.0875
-7.500 -0.2878 0.07371 0.06666 -0.0802 0.9192 0.0866
-7.250 -0.3054 0.06409 0.05672 -0.0882 0.9092 0.0859
-7.000 -0.3193 0.05924 0.05159 -0.0888 0.8974 0.0859
-6.750 -0.3085 0.05459 0.04655 -0.0918 0.8916 0.0866
-6.500 -0.3087 0.05206 0.04379 -0.0903 0.8815 0.0875
-6.250 -0.2900 0.04898 0.04032 -0.0919 0.8763 0.0899
-6.000 -0.2842 0.04664 0.03758 -0.0905 0.8681 0.0918
-5.750 -0.2677 0.04392 0.03423 -0.0906 0.8620 0.0942
-5.500 -0.2393 0.04225 0.03242 -0.0917 0.8581 0.0964
-5.250 -0.2295 0.04136 0.03140 -0.0894 0.8501 0.0982
-5.000 -0.2065 0.04015 0.02996 -0.0893 0.8448 0.1015
-4.750 -0.1767 0.03869 0.02804 -0.0901 0.8412 0.1065
-4.500 -0.1642 0.03814 0.02749 -0.0881 0.8338 0.1099
-4.250 -0.1414 0.03741 0.02668 -0.0876 0.8283 0.1147
-4.000 -0.1110 0.03642 0.02538 -0.0882 0.8246 0.1220
-3.750 -0.0929 0.03604 0.02506 -0.0869 0.8188 0.1283
-3.500 -0.0737 0.03556 0.02440 -0.0857 0.8124 0.1372
-3.250 -0.0446 0.03501 0.02385 -0.0861 0.8084 0.1498
-3.000 -0.0104 0.03441 0.02321 -0.0872 0.8054 0.1674
-2.750 -0.0036 0.03457 0.02345 -0.0843 0.7968 0.1791
-2.500 0.0249 0.03429 0.02314 -0.0846 0.7922 0.2020
-2.000 0.0684 0.03437 0.02335 -0.0831 0.7808 0.2519
-1.750 0.0939 0.03440 0.02347 -0.0828 0.7756 0.2838
-1.500 0.1278 0.03422 0.02333 -0.0836 0.7720 0.3199
-1.250 0.1402 0.03464 0.02375 -0.0815 0.7641 0.3449
-1.000 0.1661 0.03468 0.02381 -0.0811 0.7583 0.3749
-0.750 0.2013 0.03446 0.02359 -0.0820 0.7545 0.4100
-0.500 0.2138 0.03489 0.02404 -0.0798 0.7451 0.4346
-0.250 0.2481 0.03450 0.02368 -0.0803 0.7388 0.4707
0.000 0.2727 0.03434 0.02360 -0.0794 0.7300 0.5063
0.250 0.3028 0.03385 0.02324 -0.0790 0.7219 0.5528
0.500 0.3462 0.03262 0.02235 -0.0801 0.7178 0.6224
1.000 0.4380 0.03179 0.02207 -0.0857 0.7017 1.0000
1.250 0.4495 0.03243 0.02253 -0.0835 0.6909 1.0000
1.500 0.4845 0.03224 0.02211 -0.0840 0.6846 1.0000
2.000 0.5313 0.03261 0.02217 -0.0822 0.6671 1.0000
2.500 0.5786 0.03290 0.02222 -0.0802 0.6491 1.0000
3.000 0.6276 0.03303 0.02215 -0.0783 0.6304 1.0000
3.250 0.6410 0.03355 0.02261 -0.0761 0.6185 1.0000
3.500 0.6775 0.03305 0.02201 -0.0765 0.6114 1.0000
3.750 0.6874 0.03368 0.02259 -0.0738 0.5977 1.0000
4.000 0.7305 0.03284 0.02166 -0.0748 0.5919 1.0000
4.250 0.7373 0.03359 0.02239 -0.0719 0.5772 1.0000
4.500 0.7501 0.03413 0.02289 -0.0697 0.5637 1.0000
4.750 0.7914 0.03326 0.02194 -0.0703 0.5569 1.0000
5.000 0.8002 0.03398 0.02264 -0.0677 0.5420 1.0000
5.250 0.8130 0.03458 0.02323 -0.0656 0.5282 1.0000
5.500 0.8336 0.03480 0.02339 -0.0643 0.5162 1.0000
5.750 0.8652 0.03438 0.02290 -0.0639 0.5066 1.0000
6.000 0.8775 0.03512 0.02363 -0.0620 0.4930 1.0000
6.250 0.8969 0.03549 0.02395 -0.0606 0.4812 1.0000
6.500 0.9307 0.03501 0.02337 -0.0605 0.4721 1.0000
6.750 0.9421 0.03592 0.02427 -0.0586 0.4594 1.0000
7.000 0.9628 0.03631 0.02461 -0.0575 0.4488 1.0000
7.250 0.9915 0.03623 0.02443 -0.0570 0.4392 1.0000
7.500 1.0026 0.03727 0.02548 -0.0552 0.4280 1.0000
7.750 1.0309 0.03729 0.02539 -0.0547 0.4190 1.0000
8.000 1.0436 0.03826 0.02637 -0.0531 0.4085 1.0000
8.250 1.0630 0.03887 0.02694 -0.0521 0.3992 1.0000
8.500 1.0826 0.03948 0.02753 -0.0510 0.3899 1.0000
8.750 1.0959 0.04050 0.02856 -0.0496 0.3807 1.0000
9.000 1.1179 0.04100 0.02901 -0.0488 0.3720 1.0000
9.250 1.1275 0.04230 0.03036 -0.0472 0.3633 1.0000
9.500 1.1482 0.04291 0.03094 -0.0463 0.3549 1.0000
9.750 1.1576 0.04428 0.03236 -0.0448 0.3467 1.0000
10.000 1.1719 0.04532 0.03342 -0.0436 0.3385 1.0000
10.250 1.1877 0.04634 0.03446 -0.0426 0.3310 1.0000
10.500 1.1902 0.04824 0.03648 -0.0409 0.3234 1.0000
10.750 1.2244 0.04808 0.03621 -0.0409 0.3164 1.0000
11.000 1.2069 0.05141 0.03978 -0.0383 0.3090 1.0000
11.250 1.2250 0.05228 0.04066 -0.0375 0.3023 1.0000
11.500 1.2309 0.05414 0.04262 -0.0362 0.2959 1.0000
11.750 1.2179 0.05756 0.04622 -0.0346 0.2890 1.0000
12.000 1.2478 0.05744 0.04606 -0.0341 0.2832 1.0000
12.250 1.2197 0.06245 0.05130 -0.0325 0.2769 1.0000
12.500 1.2018 0.06690 0.05592 -0.0316 0.2704 1.0000
12.750 1.2363 0.06613 0.05512 -0.0309 0.2659 1.0000
13.000 1.1192 0.08308 0.07248 -0.0327 0.2558 1.0000
13.250 1.1296 0.08484 0.07429 -0.0323 0.2511 1.0000
13.500 1.1759 0.08176 0.07118 -0.0305 0.2487 1.0000
13.750 1.0415 0.10585 0.09552 -0.0381 0.2337 1.0000
14.000 1.0709 0.10439 0.09410 -0.0365 0.2320 1.0000
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