HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/9 AIRFOIL (hq259-il) Reynolds number: 100,000 Max Cl/Cd: 58.58 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq259-il-100000.txt Download as CSV file: xf-hq259-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/9 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3221 0.09298 0.08845 -0.0321 1.0000 0.0851
-8.750 -0.3312 0.08985 0.08540 -0.0345 1.0000 0.0879
-8.500 -0.3491 0.08711 0.08277 -0.0380 1.0000 0.0890
-8.250 -0.3658 0.08431 0.08008 -0.0403 1.0000 0.0893
-8.000 -0.3296 0.07856 0.07428 -0.0343 1.0000 0.0950
-7.750 -0.3345 0.07541 0.07121 -0.0344 1.0000 0.0983
-7.500 -0.3491 0.07268 0.06858 -0.0348 1.0000 0.1007
-7.250 -0.3748 0.07064 0.06670 -0.0344 1.0000 0.1017
-7.000 -0.3982 0.06789 0.06404 -0.0363 1.0000 0.1022
-6.750 -0.4231 0.06531 0.06142 -0.0381 1.0000 0.1027
-6.500 -0.4115 0.06114 0.05740 -0.0327 1.0000 0.1056
-6.250 -0.4132 0.05867 0.05497 -0.0301 1.0000 0.1091
-6.000 -0.4484 0.06392 0.05976 -0.0373 1.0000 0.1090
-5.750 -0.4520 0.06025 0.05579 -0.0412 1.0000 0.1177
-5.500 -0.4446 0.05743 0.05313 -0.0379 1.0000 0.1216
-5.250 -0.4385 0.05405 0.04956 -0.0394 1.0000 0.1328
-5.000 -0.4290 0.05114 0.04649 -0.0397 1.0000 0.1458
-4.750 -0.4179 0.04827 0.04356 -0.0393 1.0000 0.1599
-4.500 -0.4057 0.04560 0.04088 -0.0385 1.0000 0.1750
-4.250 -0.3560 0.03693 0.03069 -0.0433 1.0000 0.0875
-4.000 -0.3240 0.03112 0.02406 -0.0433 1.0000 0.0581
-3.750 -0.2983 0.02839 0.02085 -0.0431 1.0000 0.0569
-3.500 -0.2727 0.02650 0.01842 -0.0426 1.0000 0.0597
-3.250 -0.2463 0.02448 0.01596 -0.0420 1.0000 0.0604
-3.000 -0.2208 0.02300 0.01408 -0.0411 1.0000 0.0623
-2.750 -0.1966 0.02113 0.01224 -0.0406 1.0000 0.0696
-2.500 -0.1732 0.02008 0.01116 -0.0400 1.0000 0.0856
-2.250 -0.1498 0.01890 0.01012 -0.0394 1.0000 0.1140
-2.000 -0.1267 0.01798 0.00940 -0.0390 1.0000 0.1570
-1.750 -0.1032 0.01556 0.00948 -0.0377 0.9972 0.7018
-1.500 -0.0809 0.01568 0.00978 -0.0351 0.9897 0.8278
-1.250 -0.0338 0.01550 0.00965 -0.0379 0.9830 1.0000
-1.000 0.0076 0.01582 0.00964 -0.0417 0.9750 1.0000
-0.750 0.0554 0.01630 0.00979 -0.0465 0.9686 1.0000
-0.500 0.0923 0.01656 0.00983 -0.0492 0.9592 1.0000
-0.250 0.1331 0.01692 0.00999 -0.0525 0.9511 1.0000
0.000 0.1757 0.01724 0.01014 -0.0560 0.9431 1.0000
0.250 0.2113 0.01752 0.01028 -0.0581 0.9335 1.0000
0.500 0.2535 0.01780 0.01046 -0.0613 0.9257 1.0000
0.750 0.2924 0.01803 0.01061 -0.0638 0.9165 1.0000
1.000 0.3276 0.01826 0.01077 -0.0655 0.9063 1.0000
1.250 0.3706 0.01833 0.01081 -0.0684 0.8963 1.0000
1.500 0.4268 0.01806 0.01055 -0.0732 0.8874 1.0000
1.750 0.4670 0.01791 0.01043 -0.0751 0.8752 1.0000
2.000 0.5034 0.01782 0.01035 -0.0763 0.8634 1.0000
2.250 0.5397 0.01767 0.01024 -0.0774 0.8519 1.0000
2.500 0.5760 0.01744 0.01007 -0.0782 0.8402 1.0000
2.750 0.6110 0.01716 0.00989 -0.0786 0.8277 1.0000
3.000 0.6436 0.01686 0.00966 -0.0785 0.8138 1.0000
3.250 0.6744 0.01655 0.00942 -0.0779 0.7986 1.0000
3.500 0.7047 0.01622 0.00916 -0.0772 0.7825 1.0000
3.750 0.7350 0.01584 0.00889 -0.0764 0.7656 1.0000
4.000 0.7608 0.01562 0.00875 -0.0749 0.7439 1.0000
4.250 0.7882 0.01532 0.00850 -0.0736 0.7208 1.0000
4.500 0.8139 0.01511 0.00835 -0.0721 0.6933 1.0000
4.750 0.8385 0.01499 0.00826 -0.0705 0.6601 1.0000
5.000 0.8622 0.01497 0.00827 -0.0688 0.6195 1.0000
5.250 0.8845 0.01510 0.00831 -0.0669 0.5691 1.0000
5.500 0.9049 0.01550 0.00848 -0.0649 0.5092 1.0000
5.750 0.9238 0.01621 0.00888 -0.0629 0.4484 1.0000
6.000 0.9422 0.01712 0.00949 -0.0611 0.3943 1.0000
6.250 0.9610 0.01808 0.01025 -0.0596 0.3478 1.0000
6.500 0.9802 0.01911 0.01118 -0.0582 0.3088 1.0000
6.750 0.9996 0.02021 0.01212 -0.0568 0.2746 1.0000
7.000 1.0175 0.02117 0.01304 -0.0554 0.2397 1.0000
7.250 1.0325 0.02178 0.01363 -0.0537 0.2006 1.0000
7.500 1.0473 0.02238 0.01413 -0.0523 0.1480 1.0000
7.750 1.0509 0.02527 0.01622 -0.0491 0.0570 1.0000
8.000 1.0606 0.02748 0.01842 -0.0463 0.0435 1.0000
8.250 1.0695 0.02992 0.02080 -0.0438 0.0383 1.0000
8.500 1.0877 0.03199 0.02308 -0.0421 0.0355 1.0000
8.750 1.1086 0.03447 0.02574 -0.0408 0.0334 1.0000
9.000 1.1306 0.03734 0.02885 -0.0398 0.0321 1.0000
9.250 1.1506 0.04062 0.03254 -0.0386 0.0314 1.0000
9.500 1.1646 0.04379 0.03603 -0.0371 0.0304 1.0000
9.750 1.1741 0.04723 0.03968 -0.0356 0.0289 1.0000
10.000 1.1775 0.05168 0.04449 -0.0336 0.0282 1.0000
10.250 1.1734 0.05631 0.04953 -0.0312 0.0279 1.0000
10.500 1.1644 0.06046 0.05404 -0.0286 0.0277 1.0000
10.750 1.1545 0.06343 0.05729 -0.0256 0.0279 1.0000
11.000 1.1399 0.06684 0.06096 -0.0231 0.0279 1.0000
11.250 1.1274 0.06977 0.06413 -0.0214 0.0281 1.0000
11.500 1.1101 0.07363 0.06824 -0.0207 0.0282 1.0000
11.750 1.0926 0.07751 0.07240 -0.0210 0.0288 1.0000
12.000 1.0699 0.08305 0.07818 -0.0229 0.0290 1.0000
12.250 1.0038 0.09773 0.09336 -0.0332 0.0337 1.0000
12.500 0.9785 0.10794 0.10365 -0.0402 0.0354 1.0000
12.750 0.9605 0.11759 0.11331 -0.0462 0.0367 1.0000
13.000 0.9503 0.12557 0.12121 -0.0503 0.0377 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/9 AIRFOIL (hq259-il)