AH-7-47-6 AIRFOIL (ah7476-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: AH-7-47-6 AIRFOIL (ah7476-il) Reynolds number: 50,000 Max Cl/Cd: 48.92 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ah7476-il-50000-n5.txt Download as CSV file: xf-ah7476-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AH-7-47-6 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3379 0.12202 0.11525 -0.0260 1.0000 0.0666
-8.000 -0.3492 0.12195 0.11530 -0.0245 1.0000 0.0670
-7.750 -0.3585 0.12174 0.11521 -0.0242 1.0000 0.0673
-7.500 -0.3462 0.11523 0.10872 -0.0216 1.0000 0.0692
-7.250 -0.3444 0.11235 0.10588 -0.0199 1.0000 0.0716
-7.000 -0.3460 0.11041 0.10401 -0.0188 1.0000 0.0744
-6.750 -0.3486 0.10901 0.10266 -0.0189 1.0000 0.0775
-6.500 -0.3501 0.10875 0.10253 -0.0224 1.0000 0.0794
-6.250 -0.3469 0.10577 0.09960 -0.0233 1.0000 0.0806
-6.000 -0.3439 0.10154 0.09539 -0.0187 1.0000 0.0842
-5.750 -0.3394 0.09910 0.09299 -0.0191 1.0000 0.0882
-5.500 -0.3255 0.09780 0.09172 -0.0269 1.0000 0.0928
-5.250 -0.3167 0.09423 0.08820 -0.0283 1.0000 0.0945
-5.000 -0.3137 0.09075 0.08475 -0.0249 1.0000 0.0976
-4.750 -0.3013 0.08795 0.08195 -0.0270 1.0000 0.1030
-4.500 -0.2737 0.08479 0.07877 -0.0362 1.0000 0.1089
-4.250 -0.2690 0.08168 0.07569 -0.0331 1.0000 0.1144
-4.000 -0.2351 0.07832 0.07227 -0.0427 1.0000 0.1237
-3.750 -0.2242 0.07524 0.06918 -0.0416 1.0000 0.1289
-3.500 -0.1812 0.07155 0.06536 -0.0523 0.9996 0.1385
-3.000 -0.0941 0.06419 0.05776 -0.0684 0.9931 0.1679
-2.750 -0.0584 0.06085 0.05434 -0.0732 0.9897 0.1846
-2.250 0.0871 0.05112 0.04379 -0.1015 0.9858 0.1259
-1.750 0.2317 0.04232 0.03364 -0.1249 0.9828 0.0782
-1.500 0.2837 0.03998 0.03092 -0.1312 0.9801 0.0750
-1.250 0.3393 0.03783 0.02823 -0.1378 0.9770 0.0723
-1.000 0.3904 0.03631 0.02622 -0.1431 0.9711 0.0738
-0.750 0.4409 0.03501 0.02447 -0.1480 0.9640 0.0754
-0.500 0.4918 0.03383 0.02287 -0.1526 0.9574 0.0750
-0.250 0.5360 0.03296 0.02167 -0.1557 0.9497 0.0752
0.000 0.5731 0.03240 0.02083 -0.1575 0.9420 0.0759
0.250 0.6147 0.03190 0.02009 -0.1600 0.9368 0.0773
0.500 0.6469 0.03163 0.01967 -0.1609 0.9278 0.0791
0.750 0.6833 0.03131 0.01936 -0.1628 0.9200 0.0847
1.000 0.7221 0.03104 0.01906 -0.1649 0.9124 0.0939
1.250 0.7546 0.03086 0.01897 -0.1659 0.9021 0.1051
1.750 0.8234 0.02896 0.01883 -0.1685 0.8839 1.0000
2.000 0.8528 0.02907 0.01881 -0.1685 0.8716 1.0000
2.250 0.8848 0.02909 0.01872 -0.1690 0.8601 1.0000
2.500 0.9243 0.02881 0.01838 -0.1706 0.8514 1.0000
2.750 0.9535 0.02880 0.01838 -0.1705 0.8378 1.0000
3.000 0.9836 0.02870 0.01830 -0.1704 0.8242 1.0000
3.500 1.0462 0.02836 0.01811 -0.1704 0.7959 1.0000
3.750 1.0782 0.02815 0.01799 -0.1705 0.7802 1.0000
4.000 1.1101 0.02790 0.01783 -0.1704 0.7630 1.0000
4.250 1.1388 0.02761 0.01764 -0.1695 0.7409 1.0000
4.500 1.1689 0.02716 0.01730 -0.1686 0.7153 1.0000
4.750 1.1971 0.02687 0.01708 -0.1675 0.6861 1.0000
5.000 1.2282 0.02661 0.01686 -0.1669 0.6572 1.0000
5.250 1.2598 0.02651 0.01682 -0.1666 0.6288 1.0000
5.500 1.2899 0.02661 0.01693 -0.1661 0.6000 1.0000
5.750 1.3173 0.02694 0.01726 -0.1653 0.5711 1.0000
6.000 1.3424 0.02744 0.01782 -0.1642 0.5424 1.0000
6.250 1.3653 0.02806 0.01846 -0.1628 0.5133 1.0000
6.500 1.3850 0.02881 0.01921 -0.1609 0.4817 1.0000
6.750 1.4018 0.02968 0.02006 -0.1586 0.4480 1.0000
7.000 1.4154 0.03066 0.02101 -0.1558 0.4124 1.0000
7.250 1.4267 0.03170 0.02210 -0.1527 0.3768 1.0000
7.500 1.4368 0.03277 0.02327 -0.1496 0.3413 1.0000
7.750 1.4459 0.03391 0.02448 -0.1464 0.3061 1.0000
8.000 1.4514 0.03525 0.02572 -0.1428 0.2712 1.0000
8.250 1.4549 0.03682 0.02717 -0.1390 0.2363 1.0000
8.500 1.4565 0.03873 0.02889 -0.1354 0.1999 1.0000
8.750 1.4573 0.04087 0.03085 -0.1320 0.1627 1.0000
9.000 1.4590 0.04311 0.03304 -0.1289 0.1274 1.0000
9.250 1.4597 0.04559 0.03542 -0.1259 0.0989 1.0000
9.500 1.4571 0.04847 0.03816 -0.1226 0.0808 1.0000
9.750 1.4532 0.05156 0.04123 -0.1194 0.0688 1.0000
10.000 1.4519 0.05451 0.04433 -0.1165 0.0594 1.0000
10.250 1.4476 0.05778 0.04763 -0.1137 0.0539 1.0000
10.500 1.4495 0.06068 0.05082 -0.1111 0.0493 1.0000
10.750 1.4505 0.06369 0.05401 -0.1088 0.0460 1.0000
11.000 1.4498 0.06687 0.05728 -0.1067 0.0434 1.0000
11.250 1.4544 0.06989 0.06059 -0.1047 0.0406 1.0000
11.500 1.4568 0.07305 0.06411 -0.1030 0.0380 1.0000
11.750 1.4567 0.07640 0.06769 -0.1015 0.0360 1.0000
12.000 1.4566 0.07987 0.07135 -0.1001 0.0348 1.0000
12.250 1.4570 0.08352 0.07515 -0.0989 0.0338 1.0000
12.500 1.4545 0.08779 0.07971 -0.0979 0.0332 1.0000
12.750 1.4472 0.09262 0.08492 -0.0974 0.0328 1.0000
13.000 1.4363 0.09795 0.09061 -0.0977 0.0325 1.0000
13.250 1.4228 0.10379 0.09678 -0.0988 0.0323 1.0000
13.500 1.4075 0.11015 0.10343 -0.1008 0.0322 1.0000
13.750 1.3911 0.11712 0.11068 -0.1038 0.0322 1.0000
14.000 1.3742 0.12476 0.11857 -0.1079 0.0323 1.0000
14.250 1.3572 0.13311 0.12712 -0.1129 0.0325 1.0000
14.500 1.3408 0.14203 0.13622 -0.1188 0.0327 1.0000
14.750 1.3252 0.15154 0.14588 -0.1252 0.0330 1.0000
15.000 1.3113 0.16134 0.15579 -0.1319 0.0332 1.0000
15.250 1.3001 0.17109 0.16560 -0.1384 0.0335 1.0000
|
Polar data table (+)
Polar graphs
<< Back to AH-7-47-6 AIRFOIL (ah7476-il)