GOE 496 AIRFOIL (goe496-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 496 AIRFOIL (goe496-il) Reynolds number: 500,000 Max Cl/Cd: 115.08 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe496-il-500000-n5.txt Download as CSV file: xf-goe496-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 496 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
------ -------- --------- --------- -------- -------- --------
-11.500 -0.6994 0.03219 0.02916 -0.1127 0.9842 0.0189
-11.250 -0.6736 0.02740 0.02396 -0.1210 0.9794 0.0196
-11.000 -0.6421 0.02625 0.02275 -0.1239 0.9774 0.0202
-10.750 -0.6093 0.02555 0.02200 -0.1262 0.9759 0.0208
-10.500 -0.5817 0.02462 0.02097 -0.1277 0.9723 0.0215
-10.250 -0.5523 0.02348 0.01967 -0.1298 0.9691 0.0223
-10.000 -0.5209 0.02214 0.01811 -0.1322 0.9669 0.0232
-9.750 -0.4881 0.02102 0.01677 -0.1346 0.9653 0.0240
-9.500 -0.4613 0.02004 0.01560 -0.1355 0.9606 0.0245
-9.250 -0.4314 0.01900 0.01448 -0.1370 0.9572 0.0253
-9.000 -0.3989 0.01850 0.01392 -0.1385 0.9548 0.0260
-8.750 -0.3704 0.01798 0.01333 -0.1392 0.9499 0.0267
-8.500 -0.3405 0.01736 0.01260 -0.1401 0.9451 0.0275
-8.250 -0.3075 0.01667 0.01178 -0.1416 0.9417 0.0283
-8.000 -0.2804 0.01605 0.01103 -0.1418 0.9340 0.0291
-7.750 -0.2480 0.01537 0.01019 -0.1431 0.9287 0.0297
-7.500 -0.2200 0.01486 0.00955 -0.1434 0.9198 0.0301
-7.250 -0.1901 0.01411 0.00868 -0.1441 0.9122 0.0308
-7.000 -0.1616 0.01332 0.00778 -0.1446 0.9032 0.0315
-6.750 -0.1328 0.01279 0.00716 -0.1450 0.8942 0.0321
-6.500 -0.1036 0.01235 0.00663 -0.1453 0.8850 0.0327
-6.250 -0.0756 0.01197 0.00616 -0.1454 0.8746 0.0334
-6.000 -0.0473 0.01164 0.00573 -0.1455 0.8649 0.0342
-5.750 -0.0192 0.01135 0.00534 -0.1456 0.8547 0.0351
-5.500 0.0085 0.01105 0.00495 -0.1455 0.8447 0.0358
-5.250 0.0363 0.01077 0.00457 -0.1454 0.8352 0.0364
-5.000 0.0638 0.01052 0.00423 -0.1453 0.8253 0.0371
-4.750 0.0915 0.01031 0.00393 -0.1452 0.8163 0.0376
-4.500 0.1191 0.01010 0.00363 -0.1451 0.8072 0.0382
-4.250 0.1466 0.00979 0.00324 -0.1450 0.7985 0.0397
-4.000 0.1742 0.00959 0.00298 -0.1449 0.7897 0.0412
-3.750 0.2019 0.00943 0.00277 -0.1448 0.7812 0.0429
-3.500 0.2295 0.00931 0.00258 -0.1446 0.7730 0.0450
-3.250 0.2572 0.00920 0.00241 -0.1444 0.7641 0.0469
-3.000 0.2848 0.00909 0.00223 -0.1442 0.7560 0.0500
-2.750 0.3124 0.00894 0.00207 -0.1441 0.7474 0.0553
-2.500 0.3399 0.00880 0.00194 -0.1440 0.7392 0.0677
-2.250 0.3675 0.00871 0.00191 -0.1438 0.7302 0.0960
-2.000 0.3950 0.00867 0.00186 -0.1436 0.7220 0.1060
-1.750 0.4225 0.00867 0.00181 -0.1434 0.7130 0.1115
-1.500 0.4500 0.00862 0.00176 -0.1432 0.7044 0.1199
-1.250 0.4773 0.00863 0.00172 -0.1430 0.6955 0.1256
-1.000 0.5049 0.00859 0.00169 -0.1428 0.6865 0.1321
-0.750 0.5321 0.00859 0.00167 -0.1426 0.6775 0.1402
-0.500 0.5593 0.00855 0.00165 -0.1424 0.6670 0.1533
-0.250 0.5863 0.00850 0.00165 -0.1421 0.6556 0.1765
0.000 0.6130 0.00843 0.00167 -0.1419 0.6426 0.2268
0.250 0.6395 0.00842 0.00172 -0.1416 0.6300 0.2684
0.500 0.6663 0.00845 0.00176 -0.1412 0.6177 0.2918
0.750 0.6930 0.00849 0.00181 -0.1409 0.6052 0.3127
1.000 0.7196 0.00854 0.00186 -0.1406 0.5923 0.3328
1.500 0.7722 0.00868 0.00202 -0.1398 0.5669 0.3782
1.750 0.7982 0.00876 0.00212 -0.1394 0.5528 0.4115
2.000 0.8246 0.00882 0.00223 -0.1390 0.5404 0.4438
2.250 0.8511 0.00888 0.00233 -0.1387 0.5305 0.4720
2.500 0.8773 0.00894 0.00245 -0.1383 0.5213 0.5056
2.750 0.9038 0.00896 0.00258 -0.1380 0.5118 0.5479
3.000 0.9292 0.00893 0.00276 -0.1375 0.5013 0.6426
3.250 0.9533 0.00884 0.00292 -0.1367 0.4899 0.7429
3.750 1.0043 0.00876 0.00315 -0.1354 0.4679 1.0000
4.000 1.0300 0.00895 0.00330 -0.1349 0.4555 1.0000
4.250 1.0552 0.00917 0.00347 -0.1343 0.4403 1.0000
4.500 1.0800 0.00941 0.00365 -0.1337 0.4224 1.0000
4.750 1.1045 0.00967 0.00385 -0.1331 0.4031 1.0000
5.000 1.1282 0.00999 0.00408 -0.1323 0.3813 1.0000
5.250 1.1509 0.01037 0.00435 -0.1313 0.3536 1.0000
5.500 1.1710 0.01095 0.00472 -0.1300 0.3121 1.0000
6.000 1.2112 0.01212 0.00552 -0.1274 0.2471 1.0000
6.250 1.2319 0.01264 0.00593 -0.1262 0.2256 1.0000
6.500 1.2533 0.01310 0.00631 -0.1251 0.2085 1.0000
6.750 1.2746 0.01354 0.00668 -0.1240 0.1926 1.0000
7.000 1.2960 0.01396 0.00705 -0.1229 0.1784 1.0000
7.250 1.3169 0.01439 0.00744 -0.1218 0.1631 1.0000
7.500 1.3368 0.01488 0.00785 -0.1205 0.1447 1.0000
7.750 1.3540 0.01553 0.00836 -0.1188 0.1195 1.0000
8.000 1.3669 0.01643 0.00905 -0.1164 0.0863 1.0000
8.250 1.3792 0.01723 0.00974 -0.1139 0.0658 1.0000
8.500 1.3921 0.01795 0.01040 -0.1114 0.0527 1.0000
8.750 1.4059 0.01862 0.01104 -0.1092 0.0436 1.0000
9.000 1.4191 0.01932 0.01172 -0.1069 0.0357 1.0000
9.250 1.4316 0.02007 0.01245 -0.1045 0.0286 1.0000
9.500 1.4440 0.02084 0.01321 -0.1023 0.0233 1.0000
9.750 1.4561 0.02164 0.01403 -0.1000 0.0198 1.0000
10.000 1.4688 0.02242 0.01485 -0.0980 0.0177 1.0000
10.250 1.4804 0.02328 0.01575 -0.0958 0.0162 1.0000
10.500 1.4905 0.02428 0.01678 -0.0936 0.0148 1.0000
10.750 1.5023 0.02517 0.01776 -0.0916 0.0140 1.0000
11.000 1.5131 0.02617 0.01882 -0.0897 0.0132 1.0000
11.250 1.5226 0.02729 0.02000 -0.0877 0.0124 1.0000
11.500 1.5306 0.02857 0.02135 -0.0857 0.0118 1.0000
11.750 1.5363 0.03008 0.02294 -0.0837 0.0112 1.0000
12.000 1.5448 0.03141 0.02436 -0.0820 0.0109 1.0000
12.250 1.5522 0.03289 0.02593 -0.0804 0.0105 1.0000
12.500 1.5586 0.03450 0.02763 -0.0788 0.0102 1.0000
12.750 1.5641 0.03627 0.02949 -0.0774 0.0099 1.0000
13.000 1.5690 0.03814 0.03146 -0.0760 0.0095 1.0000
13.250 1.5733 0.04015 0.03356 -0.0748 0.0092 1.0000
13.500 1.5763 0.04236 0.03585 -0.0737 0.0089 1.0000
13.750 1.5769 0.04490 0.03849 -0.0727 0.0087 1.0000
14.000 1.5750 0.04783 0.04152 -0.0719 0.0084 1.0000
14.250 1.5747 0.05071 0.04450 -0.0713 0.0082 1.0000
14.500 1.5752 0.05359 0.04750 -0.0709 0.0081 1.0000
14.750 1.5760 0.05652 0.05055 -0.0707 0.0079 1.0000
15.000 1.5747 0.05985 0.05399 -0.0707 0.0077 1.0000
15.250 1.5726 0.06341 0.05767 -0.0709 0.0076 1.0000
15.500 1.5704 0.06708 0.06147 -0.0713 0.0074 1.0000
15.750 1.5666 0.07112 0.06562 -0.0719 0.0073 1.0000
16.000 1.5622 0.07537 0.07001 -0.0728 0.0072 1.0000
16.250 1.5572 0.07985 0.07461 -0.0739 0.0071 1.0000
16.500 1.5516 0.08454 0.07942 -0.0752 0.0069 1.0000
16.750 1.5453 0.08946 0.08446 -0.0767 0.0068 1.0000
17.000 1.5386 0.09458 0.08971 -0.0784 0.0068 1.0000
17.250 1.5311 0.09995 0.09519 -0.0804 0.0067 1.0000
17.500 1.5231 0.10551 0.10088 -0.0827 0.0066 1.0000
17.750 1.5148 0.11123 0.10672 -0.0851 0.0065 1.0000
18.000 1.5060 0.11718 0.11279 -0.0878 0.0065 1.0000
18.250 1.4970 0.12323 0.11896 -0.0907 0.0064 1.0000
18.500 1.4882 0.12933 0.12518 -0.0937 0.0063 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 496 AIRFOIL (goe496-il)