GOE 548 AIRFOIL (goe548-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 548 AIRFOIL (goe548-il) Reynolds number: 100,000 Max Cl/Cd: 51.06 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe548-il-100000-n5.txt Download as CSV file: xf-goe548-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 548 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.4805 0.09248 0.08759 -0.0507 1.0000 0.0236
-10.500 -0.4928 0.08664 0.08182 -0.0526 1.0000 0.0234
-10.250 -0.5107 0.07936 0.07462 -0.0555 1.0000 0.0232
-10.000 -0.5331 0.07236 0.06766 -0.0591 1.0000 0.0228
-9.750 -0.5582 0.06716 0.06246 -0.0612 1.0000 0.0225
-9.500 -0.5874 0.06358 0.05886 -0.0606 1.0000 0.0223
-9.250 -0.6160 0.06142 0.05667 -0.0569 1.0000 0.0221
-9.000 -0.6435 0.05892 0.05409 -0.0526 1.0000 0.0220
-8.750 -0.6654 0.05578 0.05079 -0.0489 1.0000 0.0220
-8.500 -0.6814 0.05258 0.04738 -0.0454 1.0000 0.0219
-8.250 -0.6945 0.04901 0.04350 -0.0420 1.0000 0.0221
-8.000 -0.6946 0.04467 0.03868 -0.0406 0.9976 0.0224
-7.750 -0.6820 0.04014 0.03349 -0.0409 0.9928 0.0230
-7.500 -0.6644 0.03634 0.02898 -0.0408 0.9882 0.0240
-7.250 -0.6415 0.03452 0.02697 -0.0415 0.9841 0.0261
-7.000 -0.6174 0.03275 0.02491 -0.0419 0.9797 0.0282
-6.750 -0.5902 0.03055 0.02228 -0.0424 0.9764 0.0305
-6.500 -0.5665 0.02879 0.02013 -0.0419 0.9718 0.0335
-6.250 -0.5395 0.02764 0.01894 -0.0426 0.9677 0.0379
-6.000 -0.5092 0.02636 0.01732 -0.0433 0.9643 0.0430
-5.750 -0.4869 0.02499 0.01598 -0.0429 0.9587 0.0490
-5.500 -0.4568 0.02409 0.01477 -0.0436 0.9547 0.0564
-5.250 -0.4312 0.02300 0.01372 -0.0439 0.9503 0.0648
-5.000 -0.4068 0.02225 0.01289 -0.0436 0.9445 0.0732
-4.750 -0.3764 0.02155 0.01209 -0.0445 0.9404 0.0806
-4.500 -0.3506 0.02093 0.01140 -0.0444 0.9349 0.0879
-4.250 -0.3241 0.02035 0.01078 -0.0444 0.9293 0.1018
-4.000 -0.2933 0.01961 0.01017 -0.0454 0.9256 0.1322
-3.750 -0.2708 0.01889 0.00979 -0.0449 0.9192 0.2084
-3.500 -0.2440 0.01830 0.00947 -0.0451 0.9138 0.2873
-3.250 -0.2122 0.01781 0.00921 -0.0462 0.9105 0.3673
-3.000 -0.1918 0.01747 0.00911 -0.0449 0.9028 0.4385
-2.750 -0.1624 0.01711 0.00898 -0.0452 0.8982 0.5116
-2.500 -0.1299 0.01679 0.00881 -0.0459 0.8944 0.5791
-2.250 -0.1085 0.01657 0.00879 -0.0443 0.8860 0.6436
-2.000 -0.0732 0.01628 0.00869 -0.0452 0.8826 0.7127
-1.750 -0.0471 0.01620 0.00866 -0.0445 0.8751 0.7584
-1.500 -0.0106 0.01608 0.00855 -0.0458 0.8708 0.7971
-1.250 0.0328 0.01599 0.00842 -0.0486 0.8682 0.8350
-1.000 0.0676 0.01605 0.00846 -0.0498 0.8611 0.8692
-0.750 0.1193 0.01608 0.00844 -0.0544 0.8580 0.9026
-0.500 0.1763 0.01609 0.00838 -0.0602 0.8559 0.9280
-0.250 0.2268 0.01607 0.00829 -0.0649 0.8523 0.9459
0.000 0.2680 0.01612 0.00829 -0.0678 0.8452 0.9629
0.250 0.3176 0.01601 0.00814 -0.0724 0.8414 0.9751
0.500 0.3613 0.01600 0.00810 -0.0760 0.8341 0.9884
0.750 0.4101 0.01588 0.00797 -0.0807 0.8289 0.9990
1.000 0.4292 0.01584 0.00791 -0.0793 0.8188 1.0000
1.250 0.4565 0.01569 0.00774 -0.0793 0.8116 1.0000
1.500 0.4744 0.01567 0.00771 -0.0775 0.8012 1.0000
1.750 0.4956 0.01564 0.00768 -0.0763 0.7918 1.0000
2.000 0.5270 0.01547 0.00752 -0.0769 0.7849 1.0000
2.250 0.5482 0.01549 0.00755 -0.0756 0.7748 1.0000
2.500 0.5763 0.01541 0.00751 -0.0756 0.7661 1.0000
2.750 0.6102 0.01522 0.00734 -0.0766 0.7562 1.0000
3.000 0.6396 0.01510 0.00722 -0.0766 0.7420 1.0000
3.250 0.6664 0.01503 0.00715 -0.0760 0.7239 1.0000
3.500 0.6916 0.01503 0.00714 -0.0752 0.7048 1.0000
3.750 0.7155 0.01510 0.00722 -0.0742 0.6871 1.0000
4.000 0.7383 0.01519 0.00734 -0.0729 0.6687 1.0000
4.250 0.7542 0.01535 0.00756 -0.0704 0.6480 1.0000
4.500 0.7722 0.01548 0.00770 -0.0682 0.6256 1.0000
4.750 0.7878 0.01564 0.00792 -0.0657 0.6017 1.0000
5.000 0.8052 0.01579 0.00806 -0.0634 0.5736 1.0000
5.250 0.8190 0.01604 0.00807 -0.0602 0.5237 1.0000
5.500 0.8294 0.01650 0.00826 -0.0566 0.4680 1.0000
5.750 0.8369 0.01711 0.00857 -0.0526 0.4047 1.0000
6.000 0.8377 0.01804 0.00899 -0.0477 0.3103 1.0000
6.250 0.8375 0.01924 0.00963 -0.0430 0.2290 1.0000
6.500 0.8454 0.02014 0.01031 -0.0397 0.1884 1.0000
6.750 0.8531 0.02105 0.01101 -0.0364 0.1478 1.0000
7.000 0.8604 0.02203 0.01175 -0.0331 0.0974 1.0000
7.250 0.8682 0.02300 0.01255 -0.0299 0.0537 1.0000
7.500 0.8721 0.02430 0.01355 -0.0263 0.0337 1.0000
7.750 0.8816 0.02531 0.01459 -0.0235 0.0301 1.0000
8.000 0.8911 0.02636 0.01574 -0.0207 0.0272 1.0000
8.250 0.9016 0.02738 0.01691 -0.0182 0.0250 1.0000
8.500 0.9112 0.02847 0.01815 -0.0156 0.0238 1.0000
8.750 0.9194 0.02969 0.01952 -0.0130 0.0228 1.0000
9.000 0.9267 0.03101 0.02097 -0.0105 0.0220 1.0000
9.250 0.9335 0.03240 0.02250 -0.0080 0.0215 1.0000
9.500 0.9388 0.03402 0.02422 -0.0055 0.0206 1.0000
9.750 0.9438 0.03590 0.02626 -0.0031 0.0197 1.0000
10.000 0.9527 0.03798 0.02841 -0.0012 0.0188 1.0000
10.250 0.9659 0.03949 0.03008 0.0003 0.0181 1.0000
10.500 0.9806 0.04123 0.03197 0.0017 0.0179 1.0000
10.750 0.9957 0.04315 0.03408 0.0029 0.0175 1.0000
11.000 1.0096 0.04528 0.03642 0.0042 0.0173 1.0000
11.250 1.0208 0.04757 0.03895 0.0056 0.0171 1.0000
11.500 1.0279 0.05010 0.04176 0.0072 0.0169 1.0000
11.750 1.0303 0.05272 0.04466 0.0090 0.0166 1.0000
12.000 1.0290 0.05557 0.04778 0.0108 0.0165 1.0000
12.250 1.0246 0.05848 0.05096 0.0124 0.0163 1.0000
12.500 1.0172 0.06158 0.05433 0.0137 0.0160 1.0000
12.750 1.0072 0.06511 0.05811 0.0146 0.0158 1.0000
13.000 0.9940 0.06910 0.06236 0.0151 0.0155 1.0000
13.250 0.9814 0.07312 0.06660 0.0149 0.0152 1.0000
13.500 0.9656 0.07791 0.07163 0.0141 0.0153 1.0000
13.750 0.9483 0.08329 0.07723 0.0126 0.0157 1.0000
14.000 0.9296 0.08920 0.08335 0.0102 0.0158 1.0000
14.250 0.9098 0.09587 0.09020 0.0069 0.0156 1.0000
14.500 0.8890 0.10351 0.09802 0.0026 0.0159 1.0000
14.750 0.8664 0.11246 0.10712 -0.0027 0.0163 1.0000
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