GOE 457 AIRFOIL (goe457-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 457 AIRFOIL (goe457-il) Reynolds number: 1,000,000 Max Cl/Cd: 108.17 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe457-il-1000000-n5.txt Download as CSV file: xf-goe457-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 457 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.3871 0.11237 0.11069 -0.0210 1.0000 0.0046
-9.500 -0.5540 0.02021 0.01707 -0.0995 0.9775 0.0060
-9.250 -0.5307 0.01767 0.01421 -0.1003 0.9743 0.0064
-9.000 -0.5061 0.01621 0.01255 -0.1005 0.9707 0.0067
-8.750 -0.4779 0.01507 0.01124 -0.1012 0.9684 0.0071
-8.500 -0.4482 0.01411 0.01013 -0.1021 0.9667 0.0075
-8.250 -0.4178 0.01329 0.00915 -0.1030 0.9654 0.0080
-8.000 -0.3929 0.01264 0.00837 -0.1025 0.9605 0.0084
-7.750 -0.3640 0.01197 0.00761 -0.1029 0.9571 0.0094
-7.500 -0.3329 0.01141 0.00697 -0.1037 0.9542 0.0102
-7.250 -0.3032 0.01093 0.00640 -0.1042 0.9500 0.0111
-7.000 -0.2738 0.01051 0.00590 -0.1045 0.9442 0.0121
-6.750 -0.2418 0.01033 0.00571 -0.1053 0.9397 0.0134
-6.500 -0.2125 0.01014 0.00549 -0.1055 0.9341 0.0147
-6.250 -0.1831 0.00990 0.00518 -0.1057 0.9276 0.0157
-6.000 -0.1529 0.00993 0.00523 -0.1060 0.9206 0.0168
-5.750 -0.1241 0.00994 0.00523 -0.1061 0.9130 0.0176
-5.500 -0.0958 0.00984 0.00507 -0.1060 0.9052 0.0186
-5.250 -0.0683 0.00972 0.00488 -0.1057 0.8971 0.0196
-5.000 -0.0407 0.00966 0.00477 -0.1055 0.8890 0.0204
-4.750 -0.0134 0.00958 0.00463 -0.1052 0.8807 0.0209
-4.500 0.0138 0.00956 0.00456 -0.1049 0.8716 0.0212
-4.250 0.0388 0.00888 0.00376 -0.1043 0.8611 0.0224
-4.000 0.0650 0.00869 0.00351 -0.1038 0.8469 0.0232
-3.750 0.0907 0.00858 0.00332 -0.1031 0.8238 0.0240
-3.500 0.1134 0.00853 0.00304 -0.1018 0.7716 0.0245
-3.250 0.1345 0.00864 0.00287 -0.1001 0.7125 0.0252
-3.000 0.1572 0.00869 0.00270 -0.0989 0.6645 0.0259
-2.750 0.1802 0.00872 0.00251 -0.0978 0.6132 0.0263
-2.500 0.2038 0.00876 0.00232 -0.0968 0.5682 0.0266
-2.250 0.2288 0.00873 0.00215 -0.0961 0.5408 0.0269
-2.000 0.2544 0.00868 0.00199 -0.0955 0.5210 0.0272
-1.750 0.2804 0.00862 0.00184 -0.0950 0.5057 0.0275
-1.250 0.3332 0.00853 0.00162 -0.0941 0.4819 0.0283
-1.000 0.3597 0.00849 0.00152 -0.0937 0.4716 0.0286
-0.750 0.3862 0.00846 0.00144 -0.0933 0.4617 0.0289
-0.500 0.4131 0.00843 0.00138 -0.0930 0.4532 0.0291
-0.250 0.4397 0.00840 0.00129 -0.0926 0.4451 0.0303
0.000 0.4664 0.00838 0.00123 -0.0922 0.4360 0.0323
0.250 0.4931 0.00839 0.00121 -0.0918 0.4272 0.0341
0.750 0.5464 0.00842 0.00120 -0.0911 0.4091 0.0390
1.000 0.5728 0.00841 0.00123 -0.0907 0.4008 0.0633
1.250 0.5993 0.00842 0.00127 -0.0903 0.3917 0.0822
1.500 0.6255 0.00847 0.00131 -0.0899 0.3793 0.0957
1.750 0.6506 0.00834 0.00139 -0.0894 0.3654 0.2111
2.000 0.6763 0.00832 0.00148 -0.0890 0.3538 0.2749
2.250 0.7011 0.00824 0.00158 -0.0885 0.3393 0.3790
2.500 0.7191 0.00753 0.00176 -0.0867 0.3252 0.7639
3.000 0.7944 0.00751 0.00202 -0.0911 0.2887 1.0000
3.250 0.8198 0.00767 0.00213 -0.0905 0.2784 1.0000
3.500 0.8450 0.00784 0.00224 -0.0899 0.2670 1.0000
3.750 0.8698 0.00805 0.00237 -0.0893 0.2511 1.0000
4.000 0.8945 0.00828 0.00252 -0.0886 0.2346 1.0000
4.250 0.9194 0.00850 0.00266 -0.0880 0.2210 1.0000
4.500 0.9441 0.00873 0.00283 -0.0874 0.2070 1.0000
4.750 0.9685 0.00898 0.00301 -0.0867 0.1908 1.0000
5.000 0.9920 0.00934 0.00324 -0.0859 0.1661 1.0000
5.250 1.0127 0.00997 0.00362 -0.0847 0.1203 1.0000
5.500 1.0354 0.01041 0.00394 -0.0838 0.1001 1.0000
5.750 1.0596 0.01069 0.00419 -0.0831 0.0902 1.0000
6.000 1.0834 0.01100 0.00444 -0.0824 0.0800 1.0000
6.250 1.1068 0.01134 0.00472 -0.0816 0.0692 1.0000
6.500 1.1302 0.01169 0.00501 -0.0809 0.0598 1.0000
6.750 1.1538 0.01200 0.00529 -0.0801 0.0535 1.0000
7.000 1.1771 0.01232 0.00559 -0.0794 0.0470 1.0000
7.250 1.1970 0.01298 0.00609 -0.0781 0.0246 1.0000
7.500 1.2178 0.01355 0.00662 -0.0769 0.0131 1.0000
7.750 1.2399 0.01396 0.00706 -0.0759 0.0107 1.0000
8.000 1.2622 0.01433 0.00748 -0.0749 0.0093 1.0000
8.250 1.2845 0.01469 0.00788 -0.0740 0.0085 1.0000
8.500 1.3061 0.01510 0.00832 -0.0730 0.0077 1.0000
8.750 1.3263 0.01560 0.00887 -0.0718 0.0067 1.0000
9.000 1.3475 0.01601 0.00931 -0.0707 0.0063 1.0000
9.250 1.3680 0.01645 0.00980 -0.0695 0.0059 1.0000
9.500 1.3878 0.01692 0.01032 -0.0683 0.0055 1.0000
9.750 1.4069 0.01742 0.01086 -0.0669 0.0052 1.0000
10.000 1.4246 0.01801 0.01150 -0.0653 0.0048 1.0000
10.250 1.4406 0.01870 0.01225 -0.0635 0.0045 1.0000
10.500 1.4579 0.01923 0.01284 -0.0619 0.0043 1.0000
10.750 1.4730 0.01977 0.01345 -0.0599 0.0041 1.0000
11.000 1.4866 0.02034 0.01408 -0.0576 0.0039 1.0000
11.250 1.4994 0.02095 0.01474 -0.0553 0.0037 1.0000
11.500 1.5108 0.02164 0.01550 -0.0528 0.0036 1.0000
11.750 1.5216 0.02237 0.01630 -0.0504 0.0035 1.0000
12.000 1.5319 0.02315 0.01713 -0.0480 0.0033 1.0000
12.250 1.5394 0.02411 0.01817 -0.0453 0.0031 1.0000
12.500 1.5451 0.02523 0.01938 -0.0426 0.0030 1.0000
12.750 1.5485 0.02653 0.02078 -0.0398 0.0029 1.0000
13.000 1.5546 0.02769 0.02205 -0.0376 0.0029 1.0000
13.250 1.5591 0.02903 0.02348 -0.0353 0.0028 1.0000
13.500 1.5622 0.03055 0.02510 -0.0332 0.0028 1.0000
13.750 1.5629 0.03235 0.02701 -0.0312 0.0028 1.0000
14.000 1.5636 0.03428 0.02905 -0.0296 0.0027 1.0000
14.250 1.5638 0.03640 0.03127 -0.0282 0.0026 1.0000
14.500 1.5613 0.03898 0.03397 -0.0273 0.0025 1.0000
14.750 1.5553 0.04218 0.03730 -0.0268 0.0025 1.0000
15.000 1.5502 0.04560 0.04084 -0.0269 0.0025 1.0000
15.250 1.5421 0.04971 0.04508 -0.0276 0.0025 1.0000
15.500 1.5333 0.05425 0.04975 -0.0289 0.0024 1.0000
15.750 1.5255 0.05888 0.05451 -0.0305 0.0024 1.0000
16.000 1.5091 0.06499 0.06077 -0.0328 0.0024 1.0000
16.250 1.5041 0.06935 0.06522 -0.0344 0.0023 1.0000
16.500 1.4835 0.07631 0.07233 -0.0372 0.0024 1.0000
16.750 1.4666 0.08280 0.07895 -0.0399 0.0023 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 457 AIRFOIL (goe457-il)