GOE 481A AIRFOIL (goe481a-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 481A AIRFOIL (goe481a-il) Reynolds number: 500,000 Max Cl/Cd: 74.08 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe481a-il-500000-n5.txt Download as CSV file: xf-goe481a-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 481A 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
------ -------- --------- --------- -------- -------- --------
-12.250 -0.6223 0.03953 0.03570 -0.1404 0.9500 0.0432
-12.000 -0.6449 0.03465 0.03048 -0.1406 0.9461 0.0434
-11.750 -0.6600 0.03276 0.02842 -0.1357 0.9387 0.0436
-11.500 -0.6539 0.03074 0.02620 -0.1344 0.9356 0.0438
-11.250 -0.6399 0.02899 0.02425 -0.1339 0.9338 0.0441
-11.000 -0.6203 0.02744 0.02252 -0.1339 0.9325 0.0443
-10.750 -0.6185 0.02653 0.02148 -0.1298 0.9259 0.0445
-10.500 -0.5992 0.02537 0.02017 -0.1290 0.9223 0.0447
-10.250 -0.5737 0.02419 0.01882 -0.1293 0.9200 0.0449
-10.000 -0.5476 0.02302 0.01756 -0.1296 0.9184 0.0452
-9.750 -0.5192 0.02213 0.01663 -0.1301 0.9173 0.0456
-9.500 -0.5143 0.02166 0.01611 -0.1258 0.9089 0.0458
-9.250 -0.4898 0.02096 0.01536 -0.1253 0.9062 0.0461
-9.000 -0.4640 0.02028 0.01463 -0.1251 0.9045 0.0464
-8.750 -0.4370 0.01963 0.01391 -0.1250 0.9032 0.0467
-8.500 -0.4085 0.01896 0.01317 -0.1253 0.9020 0.0470
-8.250 -0.3981 0.01863 0.01279 -0.1218 0.8957 0.0473
-8.000 -0.3767 0.01813 0.01223 -0.1206 0.8918 0.0476
-7.750 -0.3506 0.01757 0.01160 -0.1202 0.8894 0.0480
-7.500 -0.3220 0.01699 0.01094 -0.1203 0.8873 0.0483
-7.250 -0.2903 0.01640 0.01027 -0.1210 0.8855 0.0488
-7.000 -0.2748 0.01610 0.00991 -0.1184 0.8792 0.0491
-6.750 -0.2533 0.01574 0.00948 -0.1170 0.8740 0.0495
-6.500 -0.2235 0.01527 0.00894 -0.1172 0.8707 0.0498
-6.250 -0.1883 0.01477 0.00837 -0.1186 0.8680 0.0501
-6.000 -0.1740 0.01445 0.00803 -0.1157 0.8597 0.0505
-5.750 -0.1457 0.01403 0.00759 -0.1156 0.8538 0.0511
-5.500 -0.1084 0.01357 0.00711 -0.1174 0.8496 0.0518
-5.250 -0.0902 0.01337 0.00687 -0.1151 0.8393 0.0521
-5.000 -0.0524 0.01294 0.00640 -0.1170 0.8317 0.0528
-4.750 -0.0260 0.01267 0.00607 -0.1164 0.8184 0.0534
-4.500 0.0069 0.01235 0.00568 -0.1172 0.8040 0.0540
-4.250 0.0405 0.01207 0.00528 -0.1181 0.7831 0.0546
-4.000 0.0707 0.01190 0.00494 -0.1183 0.7546 0.0552
-3.750 0.0917 0.01189 0.00472 -0.1166 0.7207 0.0556
-3.500 0.1064 0.01188 0.00456 -0.1136 0.6883 0.0563
-3.250 0.1210 0.01191 0.00445 -0.1106 0.6630 0.0569
-3.000 0.1375 0.01193 0.00437 -0.1080 0.6432 0.0576
-2.750 0.1556 0.01193 0.00429 -0.1058 0.6270 0.0585
-2.500 0.1744 0.01193 0.00422 -0.1037 0.6134 0.0593
-2.000 0.2124 0.01195 0.00410 -0.0996 0.5878 0.0611
-1.750 0.2314 0.01195 0.00403 -0.0975 0.5772 0.0620
-1.500 0.2524 0.01190 0.00397 -0.0959 0.5678 0.0632
-1.000 0.2941 0.01189 0.00389 -0.0926 0.5507 0.0660
-0.750 0.3144 0.01191 0.00386 -0.0908 0.5413 0.0675
-0.500 0.3360 0.01190 0.00383 -0.0893 0.5340 0.0696
-0.250 0.3576 0.01189 0.00382 -0.0879 0.5264 0.0726
0.000 0.3780 0.01191 0.00382 -0.0862 0.5190 0.0766
0.500 0.4214 0.01189 0.00386 -0.0833 0.5055 0.0925
0.750 0.4416 0.01197 0.00394 -0.0816 0.4984 0.1033
1.000 0.4641 0.01202 0.00402 -0.0803 0.4914 0.1127
1.250 0.4834 0.01215 0.00409 -0.0784 0.4802 0.1188
1.500 0.5054 0.01224 0.00418 -0.0771 0.4728 0.1246
1.750 0.5269 0.01234 0.00424 -0.0756 0.4644 0.1286
2.000 0.5474 0.01245 0.00433 -0.0740 0.4576 0.1322
2.250 0.5702 0.01252 0.00441 -0.0728 0.4511 0.1359
2.500 0.5915 0.01263 0.00449 -0.0714 0.4441 0.1388
2.750 0.6119 0.01276 0.00458 -0.0698 0.4376 0.1411
3.000 0.6336 0.01282 0.00466 -0.0684 0.4308 0.1442
3.250 0.6528 0.01297 0.00479 -0.0666 0.4218 0.1481
3.500 0.6737 0.01309 0.00489 -0.0651 0.4135 0.1514
3.750 0.6928 0.01328 0.00503 -0.0634 0.4029 0.1541
4.000 0.7135 0.01338 0.00513 -0.0619 0.3960 0.1577
4.250 0.7341 0.01350 0.00526 -0.0604 0.3885 0.1612
4.500 0.7531 0.01368 0.00541 -0.0587 0.3816 0.1639
4.750 0.7745 0.01380 0.00554 -0.0574 0.3757 0.1665
5.000 0.7947 0.01397 0.00569 -0.0559 0.3691 0.1687
5.250 0.8138 0.01417 0.00587 -0.0542 0.3619 0.1714
5.500 0.8345 0.01433 0.00604 -0.0529 0.3544 0.1752
5.750 0.8536 0.01455 0.00625 -0.0513 0.3472 0.1794
6.000 0.8736 0.01475 0.00645 -0.0498 0.3417 0.1841
6.250 0.8940 0.01493 0.00667 -0.0485 0.3355 0.1941
6.750 1.1090 0.01497 0.00840 -0.0845 0.3109 1.0000
7.000 1.1280 0.01528 0.00868 -0.0830 0.3045 1.0000
7.250 1.1472 0.01560 0.00896 -0.0815 0.2966 1.0000
7.500 1.1646 0.01599 0.00930 -0.0797 0.2896 1.0000
7.750 1.1845 0.01628 0.00959 -0.0784 0.2838 1.0000
8.000 1.2019 0.01668 0.00995 -0.0767 0.2755 1.0000
8.250 1.2196 0.01708 0.01032 -0.0750 0.2680 1.0000
8.500 1.2361 0.01754 0.01073 -0.0733 0.2577 1.0000
8.750 1.2530 0.01798 0.01115 -0.0716 0.2493 1.0000
9.000 1.2679 0.01853 0.01164 -0.0696 0.2391 1.0000
9.250 1.2835 0.01906 0.01213 -0.0678 0.2279 1.0000
9.500 1.2975 0.01967 0.01269 -0.0659 0.2170 1.0000
9.750 1.3097 0.02038 0.01334 -0.0637 0.2052 1.0000
10.000 1.3219 0.02112 0.01402 -0.0616 0.1944 1.0000
10.250 1.3347 0.02183 0.01472 -0.0596 0.1872 1.0000
10.500 1.3471 0.02259 0.01544 -0.0576 0.1810 1.0000
10.750 1.3594 0.02336 0.01620 -0.0557 0.1757 1.0000
11.000 1.3716 0.02415 0.01699 -0.0538 0.1710 1.0000
11.250 1.3838 0.02496 0.01781 -0.0520 0.1673 1.0000
11.500 1.3933 0.02596 0.01880 -0.0499 0.1629 1.0000
11.750 1.4069 0.02673 0.01961 -0.0484 0.1606 1.0000
12.000 1.4193 0.02758 0.02050 -0.0468 0.1579 1.0000
12.250 1.4309 0.02851 0.02146 -0.0452 0.1556 1.0000
12.500 1.4411 0.02956 0.02252 -0.0435 0.1531 1.0000
12.750 1.4495 0.03076 0.02375 -0.0417 0.1504 1.0000
13.000 1.4590 0.03192 0.02494 -0.0402 0.1483 1.0000
13.250 1.4711 0.03293 0.02601 -0.0389 0.1464 1.0000
13.500 1.4815 0.03407 0.02720 -0.0376 0.1444 1.0000
13.750 1.4912 0.03530 0.02848 -0.0362 0.1420 1.0000
14.000 1.4982 0.03679 0.03001 -0.0348 0.1392 1.0000
14.250 1.5027 0.03852 0.03176 -0.0334 0.1359 1.0000
14.500 1.5117 0.03993 0.03323 -0.0323 0.1336 1.0000
14.750 1.5201 0.04142 0.03478 -0.0313 0.1304 1.0000
15.000 1.5242 0.04333 0.03670 -0.0302 0.1254 1.0000
15.250 1.5281 0.04531 0.03872 -0.0292 0.1218 1.0000
15.500 1.5320 0.04732 0.04077 -0.0283 0.1155 1.0000
15.750 1.5325 0.04974 0.04320 -0.0275 0.1096 1.0000
16.000 1.5297 0.05255 0.04601 -0.0267 0.1014 1.0000
16.250 1.5261 0.05549 0.04896 -0.0260 0.0949 1.0000
16.500 1.5207 0.05870 0.05219 -0.0254 0.0901 1.0000
16.750 1.5174 0.06175 0.05529 -0.0250 0.0873 1.0000
17.000 1.5114 0.06515 0.05873 -0.0247 0.0842 1.0000
17.250 1.5072 0.06839 0.06203 -0.0245 0.0823 1.0000
17.500 1.5049 0.07143 0.06515 -0.0244 0.0808 1.0000
17.750 1.5004 0.07481 0.06859 -0.0244 0.0792 1.0000
18.000 1.4937 0.07849 0.07233 -0.0246 0.0774 1.0000
18.250 1.4878 0.08211 0.07602 -0.0248 0.0763 1.0000
18.500 1.4815 0.08580 0.07978 -0.0251 0.0749 1.0000
18.750 1.4784 0.08915 0.08320 -0.0255 0.0738 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 481A AIRFOIL (goe481a-il)