GOE 5K AIRFOIL (goe05k-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 5K AIRFOIL (goe05k-il) Reynolds number: 100,000 Max Cl/Cd: 25.64 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe05k-il-100000.txt Download as CSV file: xf-goe05k-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 5K 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5298 0.10174 0.09680 -0.0115 1.0000 0.0432
-8.000 -0.5317 0.09954 0.09468 -0.0137 1.0000 0.0438
-7.750 -0.5333 0.09724 0.09245 -0.0168 1.0000 0.0442
-7.500 -0.5289 0.09458 0.08982 -0.0219 1.0000 0.0445
-7.250 -0.5259 0.08852 0.08380 -0.0165 1.0000 0.0470
-7.000 -0.5210 0.08499 0.08029 -0.0169 1.0000 0.0494
-6.750 -0.5151 0.08149 0.07681 -0.0192 1.0000 0.0519
-6.500 -0.5058 0.07806 0.07336 -0.0232 1.0000 0.0542
-6.250 -0.4887 0.07550 0.07063 -0.0306 1.0000 0.0558
-6.000 -0.4847 0.07002 0.06524 -0.0292 1.0000 0.0577
-5.750 -0.4771 0.06630 0.06155 -0.0279 1.0000 0.0618
-5.500 -0.4520 0.06431 0.05917 -0.0338 1.0000 0.0678
-5.250 -0.4462 0.05858 0.05356 -0.0327 1.0000 0.0707
-5.000 -0.4338 0.05512 0.05005 -0.0323 1.0000 0.0777
-4.750 -0.4153 0.05141 0.04611 -0.0341 1.0000 0.0851
-4.500 -0.3954 0.04861 0.04298 -0.0352 1.0000 0.0977
-4.250 -0.3792 0.04530 0.03952 -0.0351 1.0000 0.1121
-4.000 -0.3650 0.04152 0.03576 -0.0343 1.0000 0.1281
-3.750 -0.3484 0.03878 0.03287 -0.0338 1.0000 0.1566
-2.750 -0.3049 0.02861 0.02322 -0.0222 1.0000 0.3948
-2.500 -0.2183 0.02529 0.01738 -0.0308 1.0000 0.1388
-2.250 -0.1831 0.02212 0.01350 -0.0288 1.0000 0.0730
-2.000 -0.1539 0.02059 0.01137 -0.0269 1.0000 0.0550
-1.750 -0.1269 0.01841 0.00893 -0.0259 1.0000 0.0491
-1.500 -0.1009 0.01693 0.00730 -0.0247 1.0000 0.0514
-1.250 -0.0760 0.01564 0.00592 -0.0235 1.0000 0.0542
-1.000 -0.0521 0.01446 0.00470 -0.0222 1.0000 0.0681
-0.750 -0.0008 0.01071 0.00326 -0.0261 1.0000 1.0000
-0.500 0.0219 0.01078 0.00291 -0.0252 1.0000 1.0000
-0.250 0.0445 0.01086 0.00274 -0.0243 1.0000 1.0000
0.000 0.0669 0.01096 0.00265 -0.0234 1.0000 1.0000
0.250 0.0893 0.01107 0.00263 -0.0226 1.0000 1.0000
0.500 0.1116 0.01120 0.00265 -0.0218 1.0000 1.0000
0.750 0.1339 0.01134 0.00274 -0.0210 1.0000 1.0000
1.000 0.1560 0.01150 0.00289 -0.0203 1.0000 1.0000
1.250 0.1781 0.01167 0.00310 -0.0195 1.0000 1.0000
1.500 0.1998 0.01187 0.00336 -0.0187 1.0000 1.0000
1.750 0.2216 0.01209 0.00368 -0.0180 1.0000 1.0000
2.000 0.2431 0.01232 0.00406 -0.0172 1.0000 1.0000
2.250 0.2646 0.01259 0.00459 -0.0164 1.0000 1.0000
2.500 0.2859 0.01288 0.00512 -0.0157 1.0000 1.0000
2.750 0.3072 0.01320 0.00578 -0.0149 1.0000 1.0000
3.000 0.4898 0.01910 0.00886 -0.0385 0.0426 1.0000
3.250 0.5176 0.02109 0.01097 -0.0375 0.0416 1.0000
3.500 0.5471 0.02416 0.01424 -0.0366 0.0456 1.0000
3.750 0.5771 0.02683 0.01739 -0.0350 0.0569 1.0000
4.000 0.6156 0.02926 0.02083 -0.0321 0.1047 1.0000
4.250 0.6959 0.03443 0.02885 -0.0363 0.3627 1.0000
4.500 0.7103 0.03740 0.03178 -0.0331 0.3226 1.0000
5.000 0.7290 0.04345 0.03767 -0.0266 0.2210 1.0000
5.250 0.6965 0.02922 0.02397 -0.0208 0.1885 1.0000
5.500 0.7479 0.04994 0.04416 -0.0225 0.1459 1.0000
5.750 0.7565 0.05352 0.04790 -0.0211 0.1217 1.0000
6.000 0.7665 0.05709 0.05163 -0.0201 0.1056 1.0000
6.250 0.7750 0.06139 0.05605 -0.0192 0.0955 1.0000
6.500 0.7867 0.06631 0.06086 -0.0182 0.0858 1.0000
6.750 0.7889 0.06936 0.06438 -0.0186 0.0802 1.0000
7.000 0.7958 0.07322 0.06833 -0.0182 0.0740 1.0000
7.250 0.8031 0.08092 0.07583 -0.0174 0.0702 1.0000
7.500 0.7982 0.08252 0.07782 -0.0192 0.0682 1.0000
7.750 0.7963 0.08631 0.08171 -0.0212 0.0659 1.0000
8.000 0.7934 0.09045 0.08589 -0.0228 0.0634 1.0000
8.250 0.8097 0.09800 0.09327 -0.0186 0.0579 1.0000
8.500 0.8022 0.10125 0.09661 -0.0205 0.0578 1.0000
8.750 0.7927 0.10473 0.10011 -0.0227 0.0576 1.0000
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