GOE 617 AIRFOIL (goe617-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: GOE 617 AIRFOIL (goe617-il) Reynolds number: 50,000 Max Cl/Cd: 29.3 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe617-il-50000.txt Download as CSV file: xf-goe617-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 617 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3498 0.10590 0.09839 -0.0194 1.0000 0.3588
-9.250 -0.3566 0.10371 0.09628 -0.0184 1.0000 0.3716
-9.000 -0.3695 0.10229 0.09498 -0.0168 1.0000 0.3863
-8.750 -0.3629 0.09961 0.09232 -0.0153 1.0000 0.4021
-8.500 -0.3578 0.09691 0.08967 -0.0137 1.0000 0.4190
-8.250 -0.3517 0.09415 0.08695 -0.0120 1.0000 0.4358
-8.000 -0.3314 0.08997 0.08278 -0.0116 1.0000 0.4436
-7.750 -0.3455 0.08808 0.08100 -0.0093 1.0000 0.4557
-7.500 -0.6469 0.06502 0.05790 -0.0222 1.0000 0.1849
-7.250 -0.6501 0.06162 0.05445 -0.0194 1.0000 0.1824
-7.000 -0.6588 0.05812 0.05078 -0.0162 1.0000 0.1792
-6.750 -0.6731 0.05421 0.04647 -0.0124 1.0000 0.1743
-6.500 -0.6869 0.05085 0.04236 -0.0078 1.0000 0.1697
-6.250 -0.6844 0.04808 0.03924 -0.0047 1.0000 0.1698
-6.000 -0.6753 0.04553 0.03663 -0.0025 1.0000 0.1723
-5.750 -0.6652 0.04353 0.03452 -0.0002 1.0000 0.1755
-5.500 -0.6559 0.04147 0.03215 0.0023 1.0000 0.1779
-5.250 -0.6451 0.03950 0.02981 0.0047 1.0000 0.1803
-5.000 -0.6337 0.03788 0.02772 0.0072 1.0000 0.1848
-4.750 -0.6186 0.03621 0.02594 0.0089 1.0000 0.1904
-4.500 -0.6025 0.03489 0.02445 0.0106 1.0000 0.1968
-4.250 -0.5869 0.03362 0.02285 0.0124 1.0000 0.2049
-4.000 -0.5693 0.03249 0.02171 0.0138 1.0000 0.2144
-3.750 -0.5513 0.03140 0.02052 0.0151 1.0000 0.2251
-3.500 -0.5326 0.03048 0.01949 0.0163 1.0000 0.2386
-3.250 -0.5124 0.02961 0.01864 0.0172 1.0000 0.2550
-3.000 -0.4898 0.02877 0.01798 0.0176 1.0000 0.2757
-2.750 -0.4657 0.02804 0.01746 0.0177 1.0000 0.3066
-2.500 -0.4442 0.02715 0.01705 0.0183 1.0000 0.3617
-2.250 -0.1422 0.02974 0.02148 -0.0214 1.0000 1.0000
-2.000 -0.1525 0.02960 0.02120 -0.0177 1.0000 1.0000
-1.750 -0.1586 0.02960 0.02107 -0.0144 1.0000 1.0000
-1.500 -0.1614 0.02973 0.02104 -0.0114 1.0000 1.0000
-1.250 -0.1347 0.03027 0.02137 -0.0136 0.9933 1.0000
-1.000 -0.0850 0.03100 0.02185 -0.0197 0.9780 1.0000
-0.750 -0.0415 0.03166 0.02232 -0.0244 0.9630 1.0000
-0.500 -0.0010 0.03229 0.02276 -0.0284 0.9480 1.0000
-0.250 0.0371 0.03289 0.02321 -0.0317 0.9332 1.0000
0.000 0.0741 0.03348 0.02367 -0.0346 0.9186 1.0000
0.250 0.1114 0.03406 0.02413 -0.0374 0.9043 1.0000
0.500 0.1539 0.03464 0.02460 -0.0409 0.8909 1.0000
0.750 0.1799 0.03515 0.02504 -0.0416 0.8761 1.0000
1.000 0.2017 0.03573 0.02555 -0.0415 0.8617 1.0000
1.250 0.2249 0.03634 0.02610 -0.0415 0.8478 1.0000
1.500 0.2515 0.03696 0.02667 -0.0420 0.8346 1.0000
1.750 0.2904 0.03748 0.02715 -0.0443 0.8227 1.0000
2.000 0.3122 0.03811 0.02775 -0.0439 0.8092 1.0000
2.250 0.3230 0.03894 0.02854 -0.0419 0.7956 1.0000
2.500 0.3400 0.03974 0.02932 -0.0409 0.7830 1.0000
2.750 0.3725 0.04034 0.02991 -0.0419 0.7718 1.0000
3.000 0.3924 0.04112 0.03069 -0.0412 0.7598 1.0000
3.250 0.3954 0.04228 0.03183 -0.0382 0.7472 1.0000
3.500 0.4127 0.04324 0.03278 -0.0372 0.7359 1.0000
3.750 0.4498 0.04374 0.03333 -0.0386 0.7257 1.0000
4.000 0.4424 0.04531 0.03488 -0.0346 0.7133 1.0000
4.250 0.4529 0.04656 0.03613 -0.0329 0.7025 1.0000
4.500 0.5038 0.04659 0.03625 -0.0354 0.6920 1.0000
4.750 0.5036 0.04793 0.03757 -0.0321 0.6778 1.0000
5.000 0.5276 0.04830 0.03800 -0.0310 0.6619 1.0000
5.250 0.5779 0.04727 0.03707 -0.0318 0.6443 1.0000
5.500 0.6081 0.04708 0.03698 -0.0307 0.6283 1.0000
5.750 0.6122 0.04838 0.03830 -0.0278 0.6138 1.0000
6.000 0.6279 0.04909 0.03907 -0.0258 0.5989 1.0000
6.250 0.6520 0.04927 0.03934 -0.0243 0.5838 1.0000
6.500 0.6916 0.04831 0.03853 -0.0235 0.5679 1.0000
6.750 0.7704 0.04502 0.03551 -0.0258 0.5496 1.0000
7.000 0.8845 0.03909 0.02980 -0.0310 0.5243 1.0000
7.250 0.9362 0.03675 0.02747 -0.0309 0.4956 1.0000
7.500 0.9543 0.03643 0.02716 -0.0278 0.4684 1.0000
7.750 0.9843 0.03560 0.02625 -0.0259 0.4383 1.0000
8.000 1.0106 0.03519 0.02568 -0.0239 0.4089 1.0000
8.250 1.0324 0.03523 0.02555 -0.0217 0.3817 1.0000
8.500 1.0410 0.03610 0.02639 -0.0182 0.3592 1.0000
8.750 1.0537 0.03699 0.02722 -0.0154 0.3385 1.0000
9.000 1.0713 0.03785 0.02795 -0.0133 0.3193 1.0000
9.250 1.0844 0.03904 0.02911 -0.0109 0.3033 1.0000
9.500 1.0909 0.04066 0.03082 -0.0079 0.2904 1.0000
9.750 1.1016 0.04226 0.03247 -0.0055 0.2788 1.0000
10.000 1.1217 0.04359 0.03373 -0.0043 0.2672 1.0000
10.250 1.1158 0.04573 0.03612 0.0000 0.2596 1.0000
10.500 1.1333 0.04755 0.03794 0.0012 0.2515 1.0000
10.750 1.1156 0.05035 0.04104 0.0062 0.2474 1.0000
11.000 1.1520 0.05168 0.04226 0.0053 0.2381 1.0000
11.250 1.1209 0.05496 0.04586 0.0114 0.2364 1.0000
11.500 1.0820 0.05852 0.04963 0.0176 0.2354 1.0000
11.750 1.0337 0.06349 0.05475 0.0223 0.2356 1.0000
12.000 0.9703 0.07133 0.06270 0.0243 0.2370 1.0000
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Polar data table (+)
Polar graphs
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