GOE 548 AIRFOIL (goe548-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: GOE 548 AIRFOIL (goe548-il) Reynolds number: 200,000 Max Cl/Cd: 74.23 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe548-il-200000.txt Download as CSV file: xf-goe548-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4874 0.08737 0.08412 -0.0448 1.0000 0.0591
-9.250 -0.5048 0.08321 0.08002 -0.0451 1.0000 0.0590
-9.000 -0.5256 0.07876 0.07564 -0.0457 1.0000 0.0581
-8.750 -0.5613 0.07337 0.07030 -0.0475 1.0000 0.0569
-8.500 -0.5978 0.07033 0.06724 -0.0445 1.0000 0.0558
-8.250 -0.6225 0.06671 0.06357 -0.0423 1.0000 0.0560
-8.000 -0.6415 0.06312 0.05987 -0.0399 1.0000 0.0564
-7.750 -0.6573 0.05939 0.05600 -0.0374 1.0000 0.0576
-7.500 -0.6991 0.04540 0.04091 -0.0327 0.9993 0.0357
-7.250 -0.6800 0.04042 0.03538 -0.0342 0.9947 0.0356
-7.000 -0.6594 0.03643 0.03083 -0.0346 0.9894 0.0360
-6.750 -0.6345 0.03256 0.02640 -0.0353 0.9852 0.0358
-6.500 -0.6083 0.02987 0.02321 -0.0356 0.9803 0.0366
-6.250 -0.5775 0.02849 0.02137 -0.0363 0.9752 0.0379
-6.000 -0.5487 0.02544 0.01814 -0.0376 0.9726 0.0417
-5.750 -0.5201 0.02422 0.01676 -0.0380 0.9675 0.0451
-5.500 -0.4875 0.02326 0.01553 -0.0389 0.9626 0.0489
-5.250 -0.4543 0.02141 0.01365 -0.0404 0.9598 0.0552
-5.000 -0.4251 0.02056 0.01270 -0.0408 0.9543 0.0610
-4.750 -0.3962 0.01945 0.01159 -0.0414 0.9492 0.0686
-4.500 -0.3602 0.01876 0.01084 -0.0432 0.9460 0.0767
-4.250 -0.3351 0.01816 0.01027 -0.0430 0.9391 0.0855
-4.000 -0.3032 0.01756 0.00968 -0.0441 0.9344 0.0970
-3.750 -0.2668 0.01682 0.00904 -0.0460 0.9316 0.1220
-3.500 -0.2507 0.01565 0.00867 -0.0443 0.9228 0.2751
-3.250 -0.2182 0.01489 0.00842 -0.0456 0.9189 0.4074
-3.000 -0.1908 0.01452 0.00826 -0.0455 0.9128 0.4780
-2.750 -0.1627 0.01418 0.00813 -0.0455 0.9068 0.5456
-2.500 -0.1273 0.01380 0.00794 -0.0468 0.9040 0.6116
-2.250 -0.1049 0.01362 0.00791 -0.0455 0.8969 0.6632
-2.000 -0.0739 0.01334 0.00779 -0.0457 0.8924 0.7185
-1.750 -0.0368 0.01303 0.00765 -0.0470 0.8901 0.7777
-1.500 0.0056 0.01280 0.00754 -0.0493 0.8888 0.8331
-1.250 0.0354 0.01285 0.00765 -0.0492 0.8820 0.8756
-1.000 0.0901 0.01282 0.00761 -0.0544 0.8804 0.9069
-0.750 0.1454 0.01279 0.00750 -0.0598 0.8788 0.9264
-0.500 0.2015 0.01274 0.00739 -0.0654 0.8773 0.9405
-0.250 0.2647 0.01271 0.00731 -0.0725 0.8764 0.9498
0.000 0.3258 0.01266 0.00721 -0.0791 0.8751 0.9597
0.250 0.3748 0.01262 0.00715 -0.0835 0.8706 0.9702
0.500 0.4262 0.01252 0.00704 -0.0885 0.8651 0.9777
0.750 0.4790 0.01222 0.00672 -0.0936 0.8605 0.9843
1.000 0.5269 0.01196 0.00646 -0.0979 0.8509 0.9921
1.250 0.5777 0.01154 0.00603 -0.1026 0.8414 0.9985
1.500 0.6076 0.01124 0.00572 -0.1031 0.8306 1.0000
1.750 0.6242 0.01108 0.00555 -0.1009 0.8172 1.0000
2.000 0.6422 0.01096 0.00542 -0.0990 0.8045 1.0000
2.250 0.6630 0.01081 0.00525 -0.0975 0.7910 1.0000
2.500 0.6852 0.01071 0.00514 -0.0963 0.7776 1.0000
2.750 0.7090 0.01062 0.00501 -0.0953 0.7626 1.0000
3.000 0.7335 0.01060 0.00491 -0.0945 0.7458 1.0000
3.250 0.7531 0.01067 0.00496 -0.0927 0.7275 1.0000
3.500 0.7731 0.01079 0.00505 -0.0910 0.7106 1.0000
3.750 0.7918 0.01091 0.00514 -0.0890 0.6919 1.0000
4.000 0.8093 0.01104 0.00523 -0.0867 0.6714 1.0000
4.250 0.8223 0.01114 0.00529 -0.0834 0.6442 1.0000
4.500 0.8343 0.01124 0.00537 -0.0800 0.6137 1.0000
4.750 0.8456 0.01141 0.00543 -0.0764 0.5755 1.0000
5.000 0.8573 0.01169 0.00555 -0.0729 0.5382 1.0000
5.250 0.8660 0.01211 0.00575 -0.0690 0.4936 1.0000
5.500 0.8720 0.01264 0.00602 -0.0646 0.4407 1.0000
5.750 0.8743 0.01330 0.00636 -0.0596 0.3703 1.0000
6.000 0.8713 0.01424 0.00681 -0.0538 0.2794 1.0000
6.250 0.8699 0.01531 0.00743 -0.0485 0.2086 1.0000
6.500 0.8740 0.01614 0.00799 -0.0443 0.1639 1.0000
6.750 0.8775 0.01703 0.00854 -0.0399 0.1057 1.0000
7.000 0.8821 0.01796 0.00921 -0.0359 0.0679 1.0000
7.250 0.8887 0.01889 0.00992 -0.0322 0.0403 1.0000
7.500 0.8990 0.01969 0.01073 -0.0292 0.0357 1.0000
7.750 0.9088 0.02056 0.01170 -0.0261 0.0331 1.0000
8.000 0.9193 0.02138 0.01264 -0.0233 0.0319 1.0000
8.250 0.9290 0.02228 0.01365 -0.0204 0.0302 1.0000
8.500 0.9374 0.02330 0.01475 -0.0175 0.0287 1.0000
8.750 0.9450 0.02444 0.01596 -0.0147 0.0279 1.0000
9.000 0.9534 0.02565 0.01724 -0.0120 0.0273 1.0000
9.250 0.9620 0.02706 0.01870 -0.0094 0.0265 1.0000
9.500 0.9751 0.02844 0.02015 -0.0075 0.0263 1.0000
9.750 0.9926 0.02991 0.02167 -0.0061 0.0261 1.0000
10.000 1.0162 0.03153 0.02337 -0.0055 0.0260 1.0000
10.250 1.0493 0.03393 0.02585 -0.0063 0.0254 1.0000
10.500 1.0903 0.03800 0.03018 -0.0089 0.0247 1.0000
10.750 1.1052 0.03962 0.03200 -0.0071 0.0250 1.0000
11.000 1.1143 0.04111 0.03377 -0.0044 0.0254 1.0000
11.250 1.1169 0.04296 0.03608 -0.0007 0.0270 1.0000
11.500 1.1124 0.04735 0.04108 0.0032 0.0298 1.0000
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Polar data table (+)
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