GOE 600 AIRFOIL (goe600-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 600 AIRFOIL (goe600-il) Reynolds number: 200,000 Max Cl/Cd: 69.58 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe600-il-200000.txt Download as CSV file: xf-goe600-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 600 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.5973 0.10126 0.09769 -0.0305 1.0000 0.0488
-11.750 -0.5825 0.10081 0.09722 -0.0285 1.0000 0.0469
-11.500 -0.7964 0.05432 0.04993 -0.0613 1.0000 0.0338
-11.250 -0.8039 0.05161 0.04712 -0.0607 1.0000 0.0342
-11.000 -0.8126 0.04904 0.04441 -0.0590 1.0000 0.0346
-10.750 -0.8181 0.04617 0.04134 -0.0575 1.0000 0.0351
-10.500 -0.8213 0.04322 0.03812 -0.0557 1.0000 0.0356
-10.250 -0.8220 0.04026 0.03486 -0.0536 1.0000 0.0361
-10.000 -0.8204 0.03747 0.03173 -0.0512 1.0000 0.0366
-9.750 -0.8166 0.03505 0.02899 -0.0485 1.0000 0.0372
-9.500 -0.8111 0.03301 0.02664 -0.0457 1.0000 0.0380
-9.250 -0.8044 0.03139 0.02470 -0.0427 1.0000 0.0388
-9.000 -0.7976 0.02981 0.02282 -0.0397 1.0000 0.0400
-8.750 -0.7868 0.02794 0.02096 -0.0376 1.0000 0.0417
-8.500 -0.7751 0.02697 0.01994 -0.0354 1.0000 0.0434
-8.250 -0.7628 0.02593 0.01876 -0.0331 1.0000 0.0452
-8.000 -0.7497 0.02496 0.01760 -0.0309 1.0000 0.0471
-7.750 -0.7271 0.02325 0.01576 -0.0307 0.9985 0.0499
-7.500 -0.6903 0.02203 0.01454 -0.0333 0.9948 0.0547
-7.250 -0.6548 0.02068 0.01304 -0.0353 0.9905 0.0596
-7.000 -0.6187 0.01955 0.01195 -0.0377 0.9860 0.0664
-6.750 -0.5803 0.01851 0.01091 -0.0404 0.9827 0.0743
-6.500 -0.5453 0.01781 0.01020 -0.0423 0.9761 0.0835
-6.250 -0.5053 0.01725 0.00966 -0.0452 0.9720 0.0957
-6.000 -0.4662 0.01669 0.00914 -0.0479 0.9678 0.1122
-5.750 -0.4305 0.01611 0.00856 -0.0498 0.9616 0.1369
-5.500 -0.3925 0.01538 0.00791 -0.0524 0.9575 0.1644
-5.250 -0.3580 0.01478 0.00737 -0.0542 0.9517 0.1852
-5.000 -0.3251 0.01425 0.00694 -0.0555 0.9447 0.2063
-4.750 -0.2900 0.01370 0.00653 -0.0573 0.9401 0.2361
-4.500 -0.2639 0.01322 0.00627 -0.0572 0.9306 0.2887
-4.250 -0.2345 0.01265 0.00604 -0.0578 0.9249 0.3701
-4.000 -0.2097 0.01239 0.00597 -0.0572 0.9156 0.4296
-3.750 -0.1816 0.01219 0.00587 -0.0571 0.9096 0.4777
-3.500 -0.1564 0.01211 0.00587 -0.0565 0.9006 0.5148
-3.250 -0.1291 0.01203 0.00581 -0.0561 0.8945 0.5466
-3.000 -0.1031 0.01201 0.00581 -0.0556 0.8868 0.5735
-2.750 -0.0763 0.01194 0.00575 -0.0552 0.8804 0.5971
-2.500 -0.0496 0.01191 0.00571 -0.0548 0.8739 0.6194
-2.250 -0.0234 0.01184 0.00570 -0.0543 0.8669 0.6423
-2.000 0.0036 0.01178 0.00564 -0.0538 0.8622 0.6657
-1.750 0.0295 0.01174 0.00568 -0.0534 0.8546 0.6872
-1.500 0.0557 0.01163 0.00560 -0.0526 0.8482 0.7094
-1.250 0.0816 0.01156 0.00557 -0.0520 0.8403 0.7313
-1.000 0.1070 0.01142 0.00548 -0.0509 0.8328 0.7545
-0.750 0.1323 0.01134 0.00545 -0.0500 0.8246 0.7800
-0.500 0.1575 0.01123 0.00541 -0.0490 0.8177 0.8063
-0.250 0.1823 0.01120 0.00545 -0.0478 0.8115 0.8361
0.000 0.2068 0.01118 0.00552 -0.0465 0.8045 0.8697
0.250 0.2340 0.01116 0.00550 -0.0456 0.7995 0.9033
0.500 0.2669 0.01124 0.00563 -0.0462 0.7918 0.9326
0.750 0.3058 0.01125 0.00560 -0.0481 0.7859 0.9557
1.000 0.3501 0.01131 0.00564 -0.0515 0.7787 0.9721
1.250 0.3980 0.01130 0.00559 -0.0556 0.7716 0.9848
1.500 0.4476 0.01130 0.00557 -0.0602 0.7638 0.9961
1.750 0.4797 0.01125 0.00547 -0.0614 0.7555 1.0000
2.000 0.5015 0.01125 0.00544 -0.0604 0.7463 1.0000
2.250 0.5248 0.01122 0.00533 -0.0595 0.7377 1.0000
2.500 0.5475 0.01124 0.00536 -0.0586 0.7264 1.0000
2.750 0.5711 0.01127 0.00535 -0.0577 0.7156 1.0000
3.000 0.5957 0.01129 0.00530 -0.0568 0.7054 1.0000
3.250 0.6196 0.01133 0.00534 -0.0560 0.6932 1.0000
3.500 0.6437 0.01136 0.00538 -0.0551 0.6800 1.0000
3.750 0.6679 0.01138 0.00537 -0.0542 0.6653 1.0000
4.000 0.6923 0.01139 0.00536 -0.0533 0.6489 1.0000
4.250 0.7169 0.01141 0.00538 -0.0525 0.6317 1.0000
4.500 0.7418 0.01146 0.00542 -0.0518 0.6142 1.0000
4.750 0.7663 0.01153 0.00548 -0.0510 0.5923 1.0000
5.000 0.7904 0.01162 0.00552 -0.0501 0.5662 1.0000
5.250 0.8140 0.01178 0.00562 -0.0492 0.5350 1.0000
5.500 0.8370 0.01203 0.00577 -0.0482 0.4998 1.0000
5.750 0.8591 0.01241 0.00600 -0.0472 0.4660 1.0000
6.000 0.8804 0.01290 0.00631 -0.0460 0.4337 1.0000
6.250 0.9014 0.01344 0.00670 -0.0449 0.4028 1.0000
6.500 0.9225 0.01397 0.00712 -0.0439 0.3739 1.0000
6.750 0.9435 0.01449 0.00756 -0.0429 0.3463 1.0000
7.000 0.9641 0.01502 0.00800 -0.0418 0.3191 1.0000
7.250 0.9843 0.01556 0.00846 -0.0407 0.2919 1.0000
7.500 1.0041 0.01613 0.00896 -0.0396 0.2640 1.0000
7.750 1.0226 0.01677 0.00951 -0.0383 0.2350 1.0000
8.000 1.0405 0.01747 0.01010 -0.0370 0.2078 1.0000
8.250 1.0574 0.01823 0.01076 -0.0355 0.1827 1.0000
8.500 1.0732 0.01905 0.01147 -0.0339 0.1588 1.0000
8.750 1.0891 0.01986 0.01222 -0.0324 0.1350 1.0000
9.000 1.1023 0.02083 0.01308 -0.0305 0.1103 1.0000
9.250 1.1094 0.02215 0.01422 -0.0277 0.0872 1.0000
9.500 1.1137 0.02352 0.01553 -0.0245 0.0723 1.0000
9.750 1.1195 0.02489 0.01688 -0.0217 0.0636 1.0000
10.000 1.1293 0.02605 0.01804 -0.0197 0.0572 1.0000
10.250 1.1380 0.02736 0.01941 -0.0176 0.0527 1.0000
10.500 1.1481 0.02860 0.02070 -0.0159 0.0490 1.0000
10.750 1.1518 0.03043 0.02249 -0.0137 0.0463 1.0000
11.000 1.1627 0.03177 0.02396 -0.0122 0.0440 1.0000
11.250 1.1722 0.03323 0.02548 -0.0107 0.0416 1.0000
11.500 1.1799 0.03488 0.02714 -0.0093 0.0397 1.0000
11.750 1.1878 0.03695 0.02922 -0.0076 0.0380 1.0000
12.000 1.1990 0.03856 0.03098 -0.0063 0.0368 1.0000
12.250 1.2097 0.04029 0.03284 -0.0050 0.0356 1.0000
12.500 1.2194 0.04207 0.03472 -0.0038 0.0342 1.0000
12.750 1.2287 0.04389 0.03659 -0.0028 0.0330 1.0000
13.000 1.2436 0.04620 0.03888 -0.0017 0.0315 1.0000
13.250 1.2494 0.04869 0.04158 -0.0004 0.0308 1.0000
13.500 1.2521 0.05116 0.04431 0.0007 0.0304 1.0000
13.750 1.2526 0.05394 0.04732 0.0017 0.0299 1.0000
14.000 1.2506 0.05702 0.05065 0.0026 0.0295 1.0000
14.250 1.2458 0.06042 0.05429 0.0033 0.0292 1.0000
14.500 1.2382 0.06416 0.05828 0.0036 0.0289 1.0000
14.750 1.2279 0.06826 0.06261 0.0036 0.0286 1.0000
15.000 1.2152 0.07275 0.06733 0.0031 0.0284 1.0000
15.250 1.1999 0.07774 0.07255 0.0020 0.0283 1.0000
15.500 1.1814 0.08338 0.07843 0.0003 0.0282 1.0000
15.750 1.1601 0.08990 0.08519 -0.0023 0.0283 1.0000
16.000 1.1350 0.09758 0.09313 -0.0060 0.0284 1.0000
16.250 1.1054 0.10681 0.10261 -0.0113 0.0288 1.0000
16.500 1.0728 0.11761 0.11365 -0.0180 0.0293 1.0000
16.750 1.0376 0.12999 0.12623 -0.0260 0.0300 1.0000
17.000 1.0019 0.14356 0.13992 -0.0346 0.0307 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 600 AIRFOIL (goe600-il)