GOE 457 AIRFOIL (goe457-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 457 AIRFOIL (goe457-il) Reynolds number: 100,000 Max Cl/Cd: 56.89 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe457-il-100000-n5.txt Download as CSV file: xf-goe457-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 457 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3493 0.09337 0.08872 -0.0228 1.0000 0.0486
-7.250 -0.3622 0.09193 0.08739 -0.0223 1.0000 0.0499
-7.000 -0.3692 0.08977 0.08532 -0.0253 1.0000 0.0507
-6.750 -0.3694 0.08682 0.08241 -0.0283 1.0000 0.0510
-6.500 -0.3656 0.08345 0.07904 -0.0306 1.0000 0.0511
-6.250 -0.3595 0.07976 0.07534 -0.0325 1.0000 0.0512
-5.250 -0.3267 0.06210 0.05758 -0.0355 1.0000 0.0388
-5.000 -0.2968 0.05689 0.05225 -0.0410 0.9959 0.0379
-4.750 -0.2546 0.05001 0.04506 -0.0498 0.9901 0.0382
-4.250 -0.1741 0.03687 0.03086 -0.0627 0.9799 0.0404
-4.000 -0.1419 0.03530 0.02923 -0.0652 0.9749 0.0432
-3.750 -0.1072 0.03211 0.02560 -0.0679 0.9692 0.0459
-3.500 -0.0683 0.02859 0.02146 -0.0708 0.9649 0.0469
-3.250 -0.0350 0.02621 0.01843 -0.0720 0.9580 0.0502
-3.000 0.0038 0.02417 0.01580 -0.0741 0.9532 0.0512
-2.750 0.0351 0.02253 0.01391 -0.0748 0.9455 0.0523
-2.500 0.0732 0.02141 0.01264 -0.0769 0.9396 0.0546
-2.250 0.1053 0.02061 0.01169 -0.0775 0.9307 0.0581
-2.000 0.1450 0.01952 0.01038 -0.0795 0.9244 0.0599
-1.750 0.1771 0.01865 0.00939 -0.0799 0.9144 0.0614
-1.500 0.2151 0.01780 0.00846 -0.0814 0.9067 0.0634
-1.250 0.2492 0.01704 0.00774 -0.0823 0.8959 0.0668
-1.000 0.2813 0.01647 0.00715 -0.0827 0.8829 0.0716
-0.750 0.3132 0.01594 0.00661 -0.0831 0.8682 0.0815
-0.500 0.3455 0.01538 0.00607 -0.0835 0.8512 0.0991
-0.250 0.3735 0.01478 0.00565 -0.0831 0.8286 0.1416
0.000 0.4063 0.01404 0.00539 -0.0838 0.8031 0.3139
0.250 0.4580 0.01222 0.00503 -0.0879 0.7729 1.0000
0.500 0.4975 0.01213 0.00458 -0.0894 0.7353 1.0000
0.750 0.5320 0.01221 0.00430 -0.0901 0.6960 1.0000
1.000 0.5620 0.01242 0.00418 -0.0901 0.6599 1.0000
1.250 0.5896 0.01268 0.00416 -0.0897 0.6285 1.0000
1.500 0.6159 0.01296 0.00420 -0.0891 0.6020 1.0000
1.750 0.6419 0.01325 0.00428 -0.0885 0.5802 1.0000
2.000 0.6678 0.01355 0.00440 -0.0879 0.5613 1.0000
2.250 0.6935 0.01385 0.00455 -0.0874 0.5449 1.0000
2.500 0.7192 0.01416 0.00473 -0.0868 0.5299 1.0000
2.750 0.7447 0.01447 0.00495 -0.0862 0.5162 1.0000
3.000 0.7702 0.01479 0.00518 -0.0857 0.5035 1.0000
3.250 0.7956 0.01512 0.00543 -0.0851 0.4919 1.0000
3.500 0.8209 0.01543 0.00572 -0.0845 0.4799 1.0000
3.750 0.8460 0.01574 0.00602 -0.0839 0.4682 1.0000
4.000 0.8710 0.01607 0.00633 -0.0833 0.4572 1.0000
4.250 0.8957 0.01640 0.00665 -0.0826 0.4459 1.0000
4.500 0.9197 0.01671 0.00696 -0.0819 0.4328 1.0000
4.750 0.9432 0.01701 0.00726 -0.0810 0.4180 1.0000
5.000 0.9662 0.01731 0.00757 -0.0800 0.4027 1.0000
5.250 0.9889 0.01761 0.00787 -0.0790 0.3865 1.0000
5.500 1.0114 0.01790 0.00819 -0.0779 0.3690 1.0000
5.750 1.0337 0.01823 0.00854 -0.0769 0.3520 1.0000
6.000 1.0560 0.01857 0.00894 -0.0759 0.3360 1.0000
6.250 1.0781 0.01895 0.00935 -0.0749 0.3199 1.0000
6.500 1.0998 0.01935 0.00978 -0.0738 0.3029 1.0000
6.750 1.1207 0.01979 0.01024 -0.0726 0.2847 1.0000
7.000 1.1416 0.02025 0.01079 -0.0714 0.2645 1.0000
7.250 1.1617 0.02079 0.01135 -0.0702 0.2432 1.0000
7.500 1.1809 0.02141 0.01197 -0.0688 0.2195 1.0000
7.750 1.1987 0.02215 0.01268 -0.0673 0.1940 1.0000
8.000 1.2151 0.02303 0.01347 -0.0657 0.1670 1.0000
8.250 1.2299 0.02405 0.01439 -0.0639 0.1438 1.0000
8.500 1.2436 0.02515 0.01537 -0.0621 0.1264 1.0000
8.750 1.2578 0.02617 0.01638 -0.0603 0.1143 1.0000
9.000 1.2736 0.02703 0.01736 -0.0587 0.1060 1.0000
9.250 1.2865 0.02806 0.01841 -0.0568 0.0989 1.0000
9.500 1.3026 0.02883 0.01937 -0.0552 0.0908 1.0000
9.750 1.3148 0.02980 0.02046 -0.0532 0.0850 1.0000
10.000 1.3279 0.03066 0.02144 -0.0513 0.0749 1.0000
10.250 1.3395 0.03160 0.02249 -0.0492 0.0642 1.0000
10.500 1.3489 0.03269 0.02367 -0.0469 0.0530 1.0000
11.000 1.3541 0.03593 0.02686 -0.0415 0.0259 1.0000
11.250 1.3519 0.03799 0.02894 -0.0388 0.0237 1.0000
11.500 1.3478 0.04033 0.03136 -0.0363 0.0219 1.0000
11.750 1.3434 0.04283 0.03400 -0.0343 0.0209 1.0000
12.000 1.3369 0.04571 0.03703 -0.0327 0.0198 1.0000
12.250 1.3306 0.04880 0.04033 -0.0316 0.0191 1.0000
12.500 1.3237 0.05223 0.04398 -0.0313 0.0187 1.0000
12.750 1.3147 0.05624 0.04825 -0.0316 0.0183 1.0000
13.000 1.3053 0.06065 0.05286 -0.0325 0.0180 1.0000
13.250 1.2945 0.06561 0.05803 -0.0340 0.0178 1.0000
13.500 1.2828 0.07099 0.06361 -0.0360 0.0177 1.0000
13.750 1.2706 0.07667 0.06948 -0.0383 0.0175 1.0000
14.000 1.2574 0.08266 0.07565 -0.0409 0.0174 1.0000
14.250 1.2442 0.08881 0.08198 -0.0435 0.0174 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 457 AIRFOIL (goe457-il)