GOE 567 AIRFOIL (goe567-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 567 AIRFOIL (goe567-il) Reynolds number: 500,000 Max Cl/Cd: 115.17 at α=6.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe567-il-500000.txt Download as CSV file: xf-goe567-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 567 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
------ -------- --------- --------- -------- -------- --------
-15.750 -0.7041 0.08328 0.08032 -0.0770 1.0000 0.0296
-15.500 -0.7434 0.07385 0.07075 -0.0825 1.0000 0.0299
-15.250 -0.7733 0.06647 0.06324 -0.0867 1.0000 0.0301
-15.000 -0.7998 0.06018 0.05684 -0.0900 1.0000 0.0303
-14.750 -0.8246 0.05475 0.05133 -0.0925 1.0000 0.0305
-14.500 -0.8490 0.04981 0.04628 -0.0943 1.0000 0.0308
-14.250 -0.8735 0.04557 0.04195 -0.0950 1.0000 0.0310
-14.000 -0.9022 0.04188 0.03821 -0.0943 1.0000 0.0310
-13.750 -0.9406 0.03883 0.03514 -0.0914 1.0000 0.0309
-13.500 -0.9265 0.03308 0.02930 -0.1028 0.9939 0.0322
-13.250 -0.8954 0.02937 0.02547 -0.1120 0.9886 0.0339
-13.000 -0.8602 0.02692 0.02276 -0.1189 0.9846 0.0359
-12.750 -0.8248 0.02521 0.02105 -0.1240 0.9821 0.0379
-12.500 -0.7954 0.02431 0.02007 -0.1262 0.9763 0.0398
-12.250 -0.7609 0.02323 0.01880 -0.1294 0.9727 0.0418
-12.000 -0.7237 0.02252 0.01813 -0.1326 0.9703 0.0438
-11.750 -0.6881 0.02208 0.01763 -0.1350 0.9669 0.0458
-11.500 -0.6563 0.02147 0.01685 -0.1367 0.9611 0.0475
-11.250 -0.6207 0.02050 0.01584 -0.1395 0.9575 0.0494
-11.000 -0.5817 0.02023 0.01557 -0.1422 0.9545 0.0510
-10.750 -0.5532 0.01992 0.01520 -0.1428 0.9445 0.0527
-10.500 -0.5170 0.01945 0.01456 -0.1450 0.9383 0.0545
-10.250 -0.4896 0.01858 0.01358 -0.1457 0.9259 0.0562
-10.000 -0.4602 0.01820 0.01318 -0.1464 0.9132 0.0577
-9.750 -0.4306 0.01797 0.01289 -0.1470 0.8995 0.0592
-9.500 -0.4024 0.01770 0.01249 -0.1473 0.8849 0.0609
-9.250 -0.3752 0.01749 0.01211 -0.1474 0.8702 0.0626
-9.000 -0.3495 0.01707 0.01151 -0.1473 0.8563 0.0641
-8.750 -0.3255 0.01635 0.01074 -0.1470 0.8431 0.0659
-8.500 -0.2996 0.01611 0.01043 -0.1468 0.8316 0.0675
-8.250 -0.2736 0.01584 0.01005 -0.1466 0.8215 0.0691
-8.000 -0.2478 0.01550 0.00961 -0.1464 0.8117 0.0706
-7.750 -0.2213 0.01526 0.00924 -0.1462 0.8034 0.0721
-7.500 -0.1945 0.01512 0.00898 -0.1460 0.7950 0.0731
-7.250 -0.1703 0.01416 0.00792 -0.1457 0.7879 0.0752
-7.000 -0.1444 0.01371 0.00746 -0.1456 0.7806 0.0768
-6.750 -0.1176 0.01344 0.00712 -0.1454 0.7740 0.0786
-6.500 -0.0907 0.01319 0.00682 -0.1453 0.7676 0.0805
-6.250 -0.0637 0.01295 0.00650 -0.1451 0.7610 0.0823
-6.000 -0.0362 0.01277 0.00620 -0.1451 0.7553 0.0837
-5.750 -0.0091 0.01249 0.00588 -0.1449 0.7494 0.0849
-5.500 0.0168 0.01192 0.00529 -0.1448 0.7434 0.0879
-5.250 0.0443 0.01170 0.00500 -0.1447 0.7379 0.0902
-5.000 0.0716 0.01147 0.00476 -0.1446 0.7324 0.0924
-4.750 0.0992 0.01128 0.00452 -0.1445 0.7267 0.0946
-4.500 0.1272 0.01118 0.00432 -0.1445 0.7213 0.0965
-4.250 0.1544 0.01088 0.00401 -0.1444 0.7158 0.0997
-4.000 0.1820 0.01064 0.00378 -0.1444 0.7102 0.1032
-3.750 0.2101 0.01052 0.00359 -0.1444 0.7049 0.1067
-3.500 0.2380 0.01040 0.00344 -0.1443 0.6994 0.1103
-3.250 0.2656 0.01018 0.00325 -0.1442 0.6935 0.1175
-3.000 0.2937 0.01005 0.00309 -0.1442 0.6882 0.1271
-2.750 0.3214 0.00980 0.00296 -0.1443 0.6828 0.1552
-2.500 0.3488 0.00947 0.00284 -0.1443 0.6768 0.2130
-2.250 0.3765 0.00932 0.00277 -0.1443 0.6712 0.2572
-2.000 0.4043 0.00922 0.00276 -0.1443 0.6653 0.2902
-1.750 0.4320 0.00914 0.00274 -0.1442 0.6588 0.3159
-1.500 0.4599 0.00915 0.00272 -0.1440 0.6525 0.3372
-1.250 0.4874 0.00910 0.00272 -0.1439 0.6451 0.3557
-1.000 0.5149 0.00911 0.00270 -0.1437 0.6381 0.3728
-0.750 0.5426 0.00911 0.00273 -0.1435 0.6316 0.3888
-0.500 0.5703 0.00912 0.00276 -0.1434 0.6250 0.4041
-0.250 0.5978 0.00917 0.00278 -0.1432 0.6190 0.4174
0.000 0.6253 0.00917 0.00281 -0.1430 0.6115 0.4305
0.250 0.6525 0.00923 0.00283 -0.1428 0.6043 0.4432
0.500 0.6799 0.00925 0.00288 -0.1426 0.5973 0.4549
0.750 0.7072 0.00930 0.00293 -0.1424 0.5907 0.4666
1.000 0.7344 0.00939 0.00298 -0.1421 0.5845 0.4790
1.250 0.7616 0.00941 0.00306 -0.1419 0.5776 0.4905
1.500 0.7885 0.00950 0.00312 -0.1416 0.5714 0.5020
2.000 0.8426 0.00963 0.00330 -0.1412 0.5589 0.5256
2.250 0.8692 0.00974 0.00338 -0.1408 0.5528 0.5360
2.500 0.8962 0.00979 0.00347 -0.1406 0.5462 0.5475
2.750 0.9221 0.00988 0.00356 -0.1402 0.5388 0.5585
3.000 0.9482 0.00994 0.00365 -0.1398 0.5304 0.5702
3.500 1.0001 0.01013 0.00388 -0.1389 0.5166 0.5957
3.750 1.0257 0.01020 0.00399 -0.1385 0.5089 0.6097
4.000 1.0508 0.01030 0.00411 -0.1379 0.5012 0.6255
4.250 1.0761 0.01036 0.00424 -0.1374 0.4934 0.6441
4.500 1.1006 0.01046 0.00439 -0.1368 0.4862 0.6680
4.750 1.1254 0.01045 0.00454 -0.1362 0.4789 0.7066
5.000 1.1477 0.01011 0.00468 -0.1349 0.4717 1.0000
5.250 1.1733 0.01025 0.00482 -0.1344 0.4643 1.0000
5.500 1.1971 0.01046 0.00498 -0.1337 0.4560 1.0000
5.750 1.2218 0.01063 0.00515 -0.1331 0.4474 1.0000
6.000 1.2450 0.01085 0.00533 -0.1323 0.4395 1.0000
6.250 1.2692 0.01102 0.00552 -0.1316 0.4314 1.0000
6.500 1.2911 0.01128 0.00574 -0.1305 0.4225 1.0000
6.750 1.3138 0.01148 0.00595 -0.1296 0.4117 1.0000
7.000 1.3350 0.01175 0.00619 -0.1284 0.4012 1.0000
7.250 1.3544 0.01205 0.00646 -0.1269 0.3890 1.0000
7.500 1.3724 0.01236 0.00673 -0.1252 0.3755 1.0000
7.750 1.3884 0.01271 0.00704 -0.1230 0.3605 1.0000
8.000 1.4027 0.01313 0.00741 -0.1207 0.3442 1.0000
8.250 1.4146 0.01366 0.00786 -0.1180 0.3234 1.0000
8.500 1.4217 0.01440 0.00846 -0.1146 0.2960 1.0000
8.750 1.4256 0.01533 0.00922 -0.1109 0.2652 1.0000
9.000 1.4264 0.01646 0.01016 -0.1069 0.2327 1.0000
9.250 1.4252 0.01777 0.01129 -0.1029 0.1994 1.0000
9.500 1.4182 0.01949 0.01277 -0.0985 0.1580 1.0000
9.750 1.4169 0.02104 0.01418 -0.0952 0.1354 1.0000
10.000 1.4228 0.02226 0.01537 -0.0928 0.1263 1.0000
10.250 1.4282 0.02359 0.01668 -0.0906 0.1188 1.0000
10.500 1.4384 0.02466 0.01780 -0.0889 0.1143 1.0000
10.750 1.4454 0.02601 0.01916 -0.0871 0.1095 1.0000
11.000 1.4522 0.02743 0.02062 -0.0854 0.1044 1.0000
11.250 1.4630 0.02861 0.02184 -0.0841 0.0993 1.0000
11.500 1.4686 0.03023 0.02347 -0.0826 0.0934 1.0000
11.750 1.4786 0.03155 0.02483 -0.0814 0.0865 1.0000
12.000 1.4829 0.03337 0.02662 -0.0800 0.0748 1.0000
12.250 1.4859 0.03537 0.02857 -0.0787 0.0684 1.0000
12.500 1.4884 0.03748 0.03068 -0.0774 0.0646 1.0000
12.750 1.4916 0.03960 0.03282 -0.0763 0.0619 1.0000
13.000 1.4956 0.04170 0.03499 -0.0753 0.0602 1.0000
13.250 1.4985 0.04395 0.03731 -0.0744 0.0587 1.0000
13.500 1.5000 0.04642 0.03985 -0.0736 0.0571 1.0000
13.750 1.5003 0.04911 0.04260 -0.0729 0.0558 1.0000
14.000 1.4992 0.05202 0.04556 -0.0723 0.0543 1.0000
14.500 1.5032 0.05738 0.05108 -0.0715 0.0517 1.0000
14.750 1.5071 0.05993 0.05372 -0.0713 0.0504 1.0000
15.000 1.5093 0.06274 0.05661 -0.0711 0.0491 1.0000
15.250 1.5092 0.06587 0.05981 -0.0711 0.0477 1.0000
15.500 1.5057 0.06952 0.06351 -0.0712 0.0460 1.0000
15.750 1.5059 0.07275 0.06683 -0.0714 0.0445 1.0000
16.000 1.5148 0.07489 0.06907 -0.0715 0.0427 1.0000
16.250 1.5200 0.07752 0.07176 -0.0717 0.0403 1.0000
16.500 1.5180 0.08120 0.07547 -0.0722 0.0379 1.0000
16.750 1.5304 0.08289 0.07722 -0.0724 0.0336 1.0000
17.000 1.5342 0.08580 0.08016 -0.0728 0.0295 1.0000
17.250 1.5358 0.08906 0.08342 -0.0734 0.0254 1.0000
17.500 1.5341 0.09283 0.08720 -0.0741 0.0228 1.0000
17.750 1.5317 0.09677 0.09119 -0.0750 0.0212 1.0000
18.000 1.5282 0.10091 0.09537 -0.0760 0.0200 1.0000
18.250 1.5237 0.10529 0.09982 -0.0772 0.0192 1.0000
18.500 1.5210 0.10942 0.10405 -0.0784 0.0185 1.0000
18.750 1.5181 0.11358 0.10830 -0.0797 0.0180 1.0000
19.000 1.5138 0.11803 0.11283 -0.0812 0.0175 1.0000
19.250 1.5090 0.12261 0.11749 -0.0828 0.0170 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 567 AIRFOIL (goe567-il)