GOE 501 AIRFOIL (goe501-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 501 AIRFOIL (goe501-il) Reynolds number: 500,000 Max Cl/Cd: 106.17 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe501-il-500000-n5.txt Download as CSV file: xf-goe501-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 501 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.000 -0.5399 0.04619 0.04298 -0.1351 0.9624 0.0281
-12.750 -0.5661 0.03666 0.03322 -0.1474 0.9486 0.0283
-12.500 -0.5709 0.02939 0.02566 -0.1584 0.9364 0.0287
-12.250 -0.5539 0.02507 0.02101 -0.1659 0.9289 0.0292
-12.000 -0.5362 0.02362 0.01949 -0.1672 0.9217 0.0299
-11.750 -0.5140 0.02256 0.01834 -0.1683 0.9157 0.0304
-11.500 -0.4908 0.02169 0.01736 -0.1690 0.9102 0.0311
-11.250 -0.4681 0.02090 0.01647 -0.1693 0.9031 0.0319
-11.000 -0.4437 0.02004 0.01547 -0.1699 0.8974 0.0327
-10.750 -0.4203 0.01926 0.01453 -0.1701 0.8904 0.0334
-10.500 -0.3956 0.01849 0.01360 -0.1705 0.8839 0.0341
-10.250 -0.3704 0.01784 0.01289 -0.1707 0.8776 0.0348
-10.000 -0.3444 0.01740 0.01238 -0.1709 0.8705 0.0355
-9.750 -0.3178 0.01694 0.01183 -0.1711 0.8644 0.0362
-9.500 -0.2917 0.01647 0.01126 -0.1712 0.8572 0.0371
-9.250 -0.2651 0.01597 0.01062 -0.1713 0.8506 0.0378
-9.000 -0.2386 0.01550 0.01002 -0.1714 0.8440 0.0386
-8.750 -0.2119 0.01507 0.00945 -0.1714 0.8366 0.0393
-8.500 -0.1856 0.01453 0.00884 -0.1715 0.8289 0.0401
-8.250 -0.1588 0.01417 0.00839 -0.1715 0.8202 0.0408
-8.000 -0.1317 0.01385 0.00799 -0.1715 0.8125 0.0415
-7.750 -0.1045 0.01357 0.00762 -0.1715 0.8054 0.0424
-7.500 -0.0771 0.01329 0.00723 -0.1715 0.7997 0.0435
-7.250 -0.0496 0.01299 0.00685 -0.1715 0.7939 0.0444
-7.000 -0.0221 0.01271 0.00647 -0.1715 0.7881 0.0451
-6.750 0.0053 0.01236 0.00604 -0.1715 0.7828 0.0461
-6.500 0.0330 0.01207 0.00573 -0.1715 0.7778 0.0471
-6.250 0.0609 0.01186 0.00547 -0.1715 0.7728 0.0482
-6.000 0.0887 0.01169 0.00522 -0.1715 0.7676 0.0494
-5.750 0.1167 0.01150 0.00498 -0.1715 0.7627 0.0508
-5.500 0.1447 0.01132 0.00474 -0.1715 0.7581 0.0519
-5.250 0.1726 0.01111 0.00446 -0.1715 0.7532 0.0530
-5.000 0.2005 0.01089 0.00420 -0.1715 0.7483 0.0545
-4.750 0.2285 0.01072 0.00401 -0.1715 0.7433 0.0560
-4.500 0.2566 0.01059 0.00384 -0.1715 0.7386 0.0576
-4.250 0.2847 0.01048 0.00368 -0.1714 0.7342 0.0594
-4.000 0.3129 0.01039 0.00353 -0.1714 0.7301 0.0608
-3.750 0.3411 0.01017 0.00332 -0.1715 0.7258 0.0629
-3.500 0.3692 0.01004 0.00317 -0.1715 0.7211 0.0649
-3.250 0.3972 0.00995 0.00303 -0.1714 0.7163 0.0673
-3.000 0.4254 0.00987 0.00292 -0.1714 0.7118 0.0695
-2.750 0.4534 0.00976 0.00279 -0.1713 0.7065 0.0718
-2.500 0.4811 0.00966 0.00266 -0.1712 0.6998 0.0750
-2.250 0.5087 0.00958 0.00256 -0.1710 0.6924 0.0787
-2.000 0.5360 0.00952 0.00246 -0.1708 0.6837 0.0831
-1.750 0.5635 0.00942 0.00237 -0.1707 0.6757 0.0910
-1.500 0.5905 0.00927 0.00234 -0.1705 0.6656 0.1284
-1.250 0.6175 0.00925 0.00232 -0.1702 0.6540 0.1472
-1.000 0.6441 0.00923 0.00229 -0.1699 0.6410 0.1607
-0.750 0.6701 0.00925 0.00227 -0.1694 0.6241 0.1700
-0.500 0.6950 0.00932 0.00227 -0.1687 0.6019 0.1793
-0.250 0.7187 0.00944 0.00230 -0.1679 0.5773 0.1920
0.000 0.7422 0.00959 0.00238 -0.1670 0.5551 0.2101
0.250 0.7661 0.00973 0.00248 -0.1662 0.5378 0.2330
0.500 0.7908 0.00985 0.00259 -0.1656 0.5248 0.2564
0.750 0.8155 0.00998 0.00272 -0.1649 0.5138 0.2813
1.000 0.8406 0.01010 0.00284 -0.1644 0.5031 0.3052
1.250 0.8658 0.01023 0.00296 -0.1638 0.4944 0.3217
1.500 0.8911 0.01035 0.00308 -0.1633 0.4860 0.3371
1.750 0.9161 0.01048 0.00321 -0.1627 0.4783 0.3524
2.000 0.9413 0.01060 0.00334 -0.1621 0.4700 0.3673
2.250 0.9659 0.01074 0.00348 -0.1615 0.4630 0.3817
2.500 0.9914 0.01083 0.00360 -0.1610 0.4562 0.3943
3.000 1.0407 0.01108 0.00388 -0.1597 0.4442 0.4224
3.250 1.0655 0.01117 0.00403 -0.1592 0.4378 0.4427
3.500 1.0891 0.01130 0.00422 -0.1584 0.4311 0.4747
3.750 1.1134 0.01136 0.00442 -0.1577 0.4253 0.5288
4.000 1.1371 0.01145 0.00462 -0.1569 0.4190 0.5812
4.250 1.1594 0.01158 0.00483 -0.1559 0.4123 0.6238
4.500 1.1823 0.01161 0.00503 -0.1550 0.4054 0.6879
5.000 1.2252 0.01154 0.00540 -0.1524 0.3929 1.0000
5.250 1.2458 0.01178 0.00561 -0.1510 0.3838 1.0000
5.500 1.2651 0.01207 0.00585 -0.1495 0.3719 1.0000
5.750 1.2834 0.01240 0.00612 -0.1477 0.3594 1.0000
6.000 1.3019 0.01273 0.00641 -0.1461 0.3478 1.0000
6.250 1.3204 0.01308 0.00671 -0.1445 0.3343 1.0000
6.500 1.3350 0.01360 0.00711 -0.1423 0.3131 1.0000
6.750 1.3482 0.01421 0.00759 -0.1399 0.2868 1.0000
7.000 1.3553 0.01516 0.00830 -0.1366 0.2445 1.0000
7.250 1.3574 0.01646 0.00929 -0.1328 0.1953 1.0000
7.500 1.3708 0.01716 0.00992 -0.1307 0.1832 1.0000
7.750 1.3863 0.01776 0.01051 -0.1290 0.1765 1.0000
8.000 1.4024 0.01833 0.01108 -0.1273 0.1720 1.0000
8.250 1.4180 0.01895 0.01169 -0.1257 0.1679 1.0000
8.500 1.4331 0.01959 0.01234 -0.1240 0.1641 1.0000
8.750 1.4502 0.02013 0.01293 -0.1226 0.1614 1.0000
9.000 1.4663 0.02073 0.01357 -0.1211 0.1586 1.0000
9.250 1.4815 0.02141 0.01427 -0.1195 0.1557 1.0000
9.500 1.4956 0.02216 0.01505 -0.1179 0.1527 1.0000
9.750 1.5091 0.02297 0.01589 -0.1162 0.1498 1.0000
10.000 1.5240 0.02370 0.01667 -0.1148 0.1470 1.0000
10.250 1.5401 0.02437 0.01740 -0.1135 0.1432 1.0000
10.500 1.5533 0.02525 0.01829 -0.1120 0.1375 1.0000
10.750 1.5654 0.02623 0.01929 -0.1105 0.1316 1.0000
11.000 1.5771 0.02726 0.02030 -0.1089 0.1205 1.0000
11.250 1.5822 0.02881 0.02171 -0.1069 0.0960 1.0000
11.500 1.5839 0.03069 0.02349 -0.1047 0.0817 1.0000
11.750 1.5882 0.03242 0.02521 -0.1028 0.0747 1.0000
12.000 1.5961 0.03392 0.02675 -0.1013 0.0707 1.0000
12.250 1.6025 0.03559 0.02847 -0.0998 0.0672 1.0000
12.500 1.6075 0.03744 0.03034 -0.0983 0.0636 1.0000
12.750 1.6155 0.03906 0.03202 -0.0971 0.0612 1.0000
13.000 1.6215 0.04091 0.03392 -0.0959 0.0585 1.0000
13.250 1.6263 0.04293 0.03599 -0.0948 0.0562 1.0000
13.500 1.6311 0.04500 0.03812 -0.0937 0.0540 1.0000
13.750 1.6381 0.04690 0.04009 -0.0928 0.0522 1.0000
14.000 1.6435 0.04901 0.04226 -0.0919 0.0503 1.0000
14.250 1.6473 0.05135 0.04465 -0.0911 0.0481 1.0000
14.500 1.6495 0.05393 0.04729 -0.0904 0.0461 1.0000
14.750 1.6549 0.05621 0.04965 -0.0898 0.0443 1.0000
15.000 1.6585 0.05874 0.05225 -0.0894 0.0420 1.0000
15.250 1.6598 0.06162 0.05517 -0.0890 0.0395 1.0000
15.500 1.6628 0.06437 0.05800 -0.0887 0.0372 1.0000
15.750 1.6636 0.06747 0.06116 -0.0886 0.0341 1.0000
16.000 1.6637 0.07072 0.06446 -0.0886 0.0308 1.0000
16.250 1.6621 0.07429 0.06806 -0.0887 0.0270 1.0000
16.500 1.6586 0.07818 0.07199 -0.0890 0.0232 1.0000
16.750 1.6548 0.08221 0.07608 -0.0895 0.0206 1.0000
17.000 1.6511 0.08629 0.08022 -0.0900 0.0187 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 501 AIRFOIL (goe501-il)