GOE 440 AIRFOIL (goe440-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 440 AIRFOIL (goe440-il) Reynolds number: 500,000 Max Cl/Cd: 92.2 at α=1.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe440-il-500000.txt Download as CSV file: xf-goe440-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 440 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 0.2806 0.09388 0.09080 -0.1155 0.7190 0.0040
-8.000 0.2906 0.09141 0.08827 -0.1167 0.7100 0.0041
-7.750 0.3008 0.08893 0.08575 -0.1180 0.7012 0.0042
-7.500 0.3111 0.08650 0.08328 -0.1193 0.6930 0.0044
-7.250 0.3217 0.08406 0.08081 -0.1207 0.6845 0.0045
-7.000 0.3324 0.08166 0.07838 -0.1221 0.6768 0.0047
-6.750 0.3435 0.07929 0.07599 -0.1236 0.6689 0.0048
-6.500 0.3549 0.07699 0.07366 -0.1253 0.6615 0.0050
-6.250 0.3663 0.07478 0.07143 -0.1270 0.6534 0.0051
-6.000 0.3769 0.07267 0.06930 -0.1289 0.6455 0.0052
-5.750 0.3867 0.07073 0.06734 -0.1307 0.6375 0.0053
-5.500 0.4021 0.06831 0.06491 -0.1339 0.6290 0.0053
-5.250 0.4188 0.06578 0.06234 -0.1371 0.6207 0.0053
-5.000 0.4375 0.06308 0.05961 -0.1408 0.6116 0.0054
-4.750 0.4578 0.06028 0.05678 -0.1449 0.6029 0.0054
-4.500 0.4734 0.05656 0.05302 -0.1484 0.5940 0.0056
-4.250 0.4882 0.05392 0.05035 -0.1500 0.5838 0.0060
-4.000 0.5112 0.05110 0.04747 -0.1544 0.5732 0.0064
-3.750 0.5390 0.04805 0.04433 -0.1602 0.5618 0.0069
-3.500 0.5725 0.04467 0.04086 -0.1673 0.5495 0.0074
-3.250 0.6137 0.04084 0.03691 -0.1760 0.5371 0.0081
-2.250 0.8194 0.02443 0.01928 -0.2121 0.4889 0.0122
-2.000 0.8616 0.02225 0.01677 -0.2164 0.4786 0.0150
-1.750 0.8981 0.02195 0.01603 -0.2172 0.4703 0.0196
-1.500 0.9437 0.01865 0.01234 -0.2228 0.4623 0.0237
-1.250 0.9812 0.01754 0.01078 -0.2248 0.4555 0.0315
-1.000 1.0180 0.01599 0.00895 -0.2259 0.4489 0.0169
-0.750 1.0529 0.01492 0.00760 -0.2268 0.4430 0.0089
-0.500 1.0883 0.01416 0.00668 -0.2284 0.4373 0.0088
-0.250 1.1208 0.01385 0.00616 -0.2297 0.4321 0.0119
0.000 1.1534 0.01361 0.00592 -0.2309 0.4273 0.1014
0.250 1.1857 0.01329 0.00618 -0.2328 0.4223 0.4054
0.500 1.2113 0.01347 0.00625 -0.2328 0.4176 0.4343
0.750 1.2363 0.01362 0.00636 -0.2327 0.4054 0.4632
1.000 1.2610 0.01381 0.00657 -0.2327 0.3969 0.5149
1.250 1.2862 0.01395 0.00678 -0.2327 0.3888 0.5824
1.500 1.3056 0.01434 0.00700 -0.2317 0.3572 0.6310
1.750 1.3260 0.01469 0.00733 -0.2310 0.3345 0.6960
2.000 1.3320 0.01550 0.00804 -0.2276 0.2806 0.8444
2.250 1.3223 0.01693 0.00909 -0.2213 0.2180 1.0000
2.500 1.2953 0.02022 0.01173 -0.2131 0.1046 1.0000
2.750 1.3024 0.02151 0.01291 -0.2104 0.0263 1.0000
3.000 1.3208 0.02206 0.01347 -0.2095 0.0266 1.0000
3.250 1.3388 0.02265 0.01407 -0.2085 0.0286 1.0000
4.000 1.3860 0.02495 0.01658 -0.2047 0.0222 1.0000
4.250 1.3977 0.02605 0.01779 -0.2030 0.0180 1.0000
4.500 1.4110 0.02704 0.01884 -0.2016 0.0160 1.0000
4.750 1.4244 0.02804 0.01989 -0.2002 0.0134 1.0000
5.000 1.4336 0.02941 0.02137 -0.1984 0.0117 1.0000
5.250 1.4374 0.03124 0.02333 -0.1961 0.0105 1.0000
5.500 1.4328 0.03385 0.02610 -0.1932 0.0097 1.0000
5.750 1.4176 0.03756 0.03001 -0.1897 0.0091 1.0000
6.000 1.4173 0.04008 0.03263 -0.1878 0.0090 1.0000
6.250 1.4509 0.03939 0.03185 -0.1883 0.0079 1.0000
6.500 1.4586 0.04121 0.03374 -0.1870 0.0071 1.0000
6.750 1.4592 0.04382 0.03645 -0.1854 0.0065 1.0000
7.000 1.4561 0.04692 0.03966 -0.1838 0.0061 1.0000
7.250 1.4493 0.05061 0.04348 -0.1822 0.0058 1.0000
7.500 1.4384 0.05498 0.04799 -0.1808 0.0056 1.0000
7.750 1.4242 0.05993 0.05309 -0.1795 0.0055 1.0000
8.000 1.4068 0.06548 0.05879 -0.1785 0.0054 1.0000
8.250 1.3882 0.07143 0.06488 -0.1777 0.0053 1.0000
8.500 1.3693 0.07756 0.07115 -0.1771 0.0053 1.0000
8.750 1.3498 0.08392 0.07764 -0.1767 0.0052 1.0000
9.000 1.3306 0.09035 0.08418 -0.1765 0.0052 1.0000
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Polar data table (+)
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