GOE 566 AIRFOIL (goe566-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 566 AIRFOIL (goe566-il) Reynolds number: 50,000 Max Cl/Cd: 34.44 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe566-il-50000.txt Download as CSV file: xf-goe566-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 566 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4292 0.10751 0.10036 -0.0162 1.0000 0.2248
-8.750 -0.4197 0.10347 0.09636 -0.0158 1.0000 0.2325
-8.500 -0.4379 0.10273 0.09576 -0.0173 1.0000 0.2415
-8.250 -0.4185 0.09795 0.09098 -0.0158 1.0000 0.2539
-8.000 -0.4126 0.09456 0.08762 -0.0151 1.0000 0.2647
-7.750 -0.4115 0.09172 0.08486 -0.0144 1.0000 0.2790
-7.500 -0.4095 0.08900 0.08220 -0.0131 1.0000 0.2958
-7.250 -0.4456 0.08896 0.08242 -0.0129 1.0000 0.3029
-7.000 -0.4187 0.08460 0.07802 -0.0097 1.0000 0.3313
-6.750 -0.4324 0.08296 0.07654 -0.0075 1.0000 0.3480
-6.500 -0.4121 0.07925 0.07282 -0.0044 1.0000 0.3757
-6.250 -0.4032 0.07642 0.07004 -0.0014 1.0000 0.4037
-6.000 -0.4018 0.07429 0.06799 0.0025 1.0000 0.4362
-5.750 -0.3753 0.07124 0.06490 0.0068 1.0000 0.4964
-5.500 -0.3610 0.06915 0.06285 0.0121 1.0000 0.5614
-5.250 -0.2887 0.06364 0.05712 0.0124 1.0000 0.6569
-5.000 -0.2077 0.05772 0.05098 0.0094 1.0000 0.7823
-4.500 -0.1994 0.05326 0.04667 0.0113 1.0000 0.7968
-4.000 -0.3758 0.04219 0.03416 -0.0254 1.0000 0.1928
-3.750 -0.3534 0.03877 0.03018 -0.0255 1.0000 0.1741
-3.500 -0.3304 0.03610 0.02684 -0.0251 1.0000 0.1633
-3.250 -0.3096 0.03399 0.02438 -0.0242 1.0000 0.1621
-3.000 -0.2878 0.03221 0.02219 -0.0234 1.0000 0.1630
-2.750 -0.2652 0.03053 0.02015 -0.0225 1.0000 0.1624
-2.500 -0.2417 0.02906 0.01835 -0.0217 1.0000 0.1626
-2.250 -0.2180 0.02783 0.01677 -0.0208 1.0000 0.1644
-2.000 -0.1950 0.02669 0.01550 -0.0200 1.0000 0.1696
-1.750 -0.1718 0.02588 0.01454 -0.0192 1.0000 0.1794
-1.500 -0.1484 0.02496 0.01362 -0.0183 1.0000 0.1888
-1.250 -0.1232 0.02420 0.01278 -0.0177 1.0000 0.2023
-1.000 -0.0993 0.02342 0.01207 -0.0170 1.0000 0.2302
-0.750 -0.0488 0.01933 0.01067 -0.0192 1.0000 1.0000
-0.500 -0.0305 0.01965 0.01054 -0.0180 1.0000 1.0000
-0.250 -0.0125 0.02000 0.01057 -0.0170 1.0000 1.0000
0.000 0.0055 0.02039 0.01071 -0.0161 1.0000 1.0000
0.250 0.0234 0.02081 0.01092 -0.0153 1.0000 1.0000
0.500 0.0414 0.02126 0.01120 -0.0146 1.0000 1.0000
0.750 0.0592 0.02176 0.01153 -0.0139 1.0000 1.0000
1.000 0.0769 0.02230 0.01195 -0.0133 1.0000 1.0000
1.250 0.0945 0.02289 0.01243 -0.0128 1.0000 1.0000
1.500 0.1118 0.02353 0.01298 -0.0124 1.0000 1.0000
1.750 0.1292 0.02423 0.01361 -0.0121 0.9998 1.0000
2.000 0.2106 0.02610 0.01543 -0.0236 0.9705 1.0000
2.250 0.2789 0.02723 0.01656 -0.0319 0.9419 1.0000
2.500 0.3323 0.02795 0.01732 -0.0371 0.9160 1.0000
2.750 0.3866 0.02851 0.01797 -0.0419 0.8911 1.0000
3.000 0.4471 0.02875 0.01835 -0.0472 0.8664 1.0000
3.250 0.4954 0.02878 0.01855 -0.0501 0.8392 1.0000
3.500 0.5516 0.02838 0.01835 -0.0535 0.8112 1.0000
3.750 0.6048 0.02764 0.01781 -0.0556 0.7821 1.0000
4.000 0.6545 0.02660 0.01699 -0.0564 0.7510 1.0000
4.250 0.7013 0.02542 0.01595 -0.0562 0.7180 1.0000
4.500 0.7346 0.02481 0.01543 -0.0545 0.6797 1.0000
4.750 0.7697 0.02428 0.01486 -0.0530 0.6436 1.0000
5.000 0.7960 0.02445 0.01501 -0.0512 0.6080 1.0000
5.250 0.8214 0.02477 0.01527 -0.0495 0.5756 1.0000
5.500 0.8457 0.02514 0.01562 -0.0478 0.5456 1.0000
5.750 0.8688 0.02549 0.01591 -0.0459 0.5159 1.0000
6.000 0.8916 0.02596 0.01631 -0.0441 0.4879 1.0000
6.250 0.9165 0.02661 0.01691 -0.0428 0.4636 1.0000
6.500 0.9377 0.02754 0.01792 -0.0413 0.4389 1.0000
6.750 0.9592 0.02856 0.01897 -0.0398 0.4135 1.0000
7.000 0.9814 0.02974 0.02014 -0.0383 0.3869 1.0000
7.250 0.9989 0.03085 0.02128 -0.0361 0.3535 1.0000
7.500 1.0148 0.03149 0.02166 -0.0335 0.3138 1.0000
7.750 1.0271 0.03245 0.02254 -0.0307 0.2753 1.0000
8.000 1.0404 0.03394 0.02390 -0.0282 0.2380 1.0000
8.250 1.0519 0.03561 0.02544 -0.0254 0.2006 1.0000
8.500 1.0644 0.03746 0.02727 -0.0228 0.1724 1.0000
8.750 1.0799 0.03940 0.02921 -0.0209 0.1535 1.0000
9.000 1.0955 0.04164 0.03157 -0.0191 0.1399 1.0000
9.250 1.1135 0.04410 0.03409 -0.0177 0.1296 1.0000
9.500 1.1198 0.04739 0.03792 -0.0152 0.1227 1.0000
9.750 1.1404 0.05029 0.04070 -0.0145 0.1160 1.0000
10.000 1.1377 0.05441 0.04545 -0.0117 0.1142 1.0000
10.250 1.1299 0.05866 0.05023 -0.0089 0.1128 1.0000
10.500 1.1172 0.06302 0.05499 -0.0063 0.1120 1.0000
10.750 1.0989 0.06746 0.05975 -0.0038 0.1119 1.0000
11.000 1.0748 0.07183 0.06434 -0.0014 0.1124 1.0000
11.250 1.0492 0.07673 0.06940 -0.0003 0.1133 1.0000
11.500 1.0250 0.08232 0.07510 -0.0008 0.1145 1.0000
11.750 1.0017 0.08862 0.08147 -0.0025 0.1154 1.0000
12.000 0.8530 0.12200 0.11459 -0.0294 0.1479 1.0000
12.250 0.8441 0.12949 0.12203 -0.0328 0.1523 1.0000
12.500 0.8452 0.13545 0.12799 -0.0345 0.1543 1.0000
12.750 0.6521 0.15002 0.14300 -0.0413 0.2708 1.0000
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Polar data table (+)
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