HQ 3.0/9 AIRFOIL (hq309-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.0/9 AIRFOIL (hq309-il) Reynolds number: 500,000 Max Cl/Cd: 96.32 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq309-il-500000-n5.txt Download as CSV file: xf-hq309-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.0/9 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3811 0.08630 0.08411 -0.0364 1.0000 0.0042
-8.500 -0.3847 0.08281 0.08067 -0.0370 1.0000 0.0042
-8.250 -0.3894 0.08022 0.07813 -0.0367 1.0000 0.0043
-8.000 -0.3799 0.07438 0.07231 -0.0432 0.9915 0.0043
-7.750 -0.3672 0.06888 0.06682 -0.0509 0.9825 0.0044
-7.500 -0.3452 0.06232 0.06022 -0.0629 0.9751 0.0046
-7.250 -0.3232 0.05361 0.05139 -0.0767 0.9651 0.0046
-7.000 -0.3031 0.04594 0.04350 -0.0854 0.9550 0.0047
-6.750 -0.2821 0.04040 0.03773 -0.0902 0.9458 0.0049
-6.500 -0.2631 0.03581 0.03288 -0.0924 0.9352 0.0053
-6.250 -0.2439 0.03072 0.02744 -0.0938 0.9256 0.0057
-6.000 -0.2246 0.02475 0.02096 -0.0942 0.9169 0.0059
-5.750 -0.2044 0.01933 0.01486 -0.0939 0.9084 0.0058
-5.500 -0.1804 0.01651 0.01154 -0.0936 0.9018 0.0059
-5.250 -0.1551 0.01473 0.00940 -0.0932 0.8950 0.0061
-5.000 -0.1290 0.01346 0.00787 -0.0929 0.8891 0.0065
-4.750 -0.1028 0.01259 0.00681 -0.0925 0.8824 0.0068
-4.500 -0.0760 0.01196 0.00604 -0.0923 0.8767 0.0072
-4.250 -0.0505 0.01083 0.00475 -0.0920 0.8700 0.0080
-4.000 -0.0236 0.01044 0.00426 -0.0919 0.8642 0.0093
-3.750 0.0034 0.01000 0.00374 -0.0917 0.8577 0.0099
-3.500 0.0307 0.00960 0.00322 -0.0915 0.8516 0.0105
-3.250 0.0581 0.00928 0.00281 -0.0914 0.8452 0.0113
-3.000 0.0857 0.00901 0.00244 -0.0913 0.8388 0.0133
-2.750 0.1132 0.00874 0.00220 -0.0912 0.8324 0.0244
-2.500 0.1408 0.00858 0.00203 -0.0912 0.8254 0.0379
-2.250 0.1683 0.00842 0.00188 -0.0911 0.8177 0.0545
-2.000 0.1956 0.00827 0.00173 -0.0910 0.8089 0.0725
-1.750 0.2227 0.00794 0.00158 -0.0910 0.7992 0.1391
-1.500 0.2496 0.00748 0.00144 -0.0911 0.7908 0.2600
-1.250 0.2759 0.00691 0.00134 -0.0912 0.7827 0.4333
-1.000 0.3029 0.00669 0.00133 -0.0911 0.7745 0.5176
-0.750 0.3299 0.00658 0.00132 -0.0909 0.7660 0.5774
-0.500 0.3563 0.00649 0.00135 -0.0906 0.7544 0.6391
-0.250 0.3822 0.00644 0.00139 -0.0900 0.7405 0.6937
0.000 0.4089 0.00645 0.00140 -0.0896 0.7269 0.7160
0.250 0.4363 0.00648 0.00139 -0.0895 0.7152 0.7256
0.500 0.4636 0.00651 0.00140 -0.0893 0.7036 0.7350
0.750 0.4906 0.00656 0.00141 -0.0891 0.6897 0.7445
1.000 0.5172 0.00664 0.00142 -0.0888 0.6720 0.7547
1.250 0.5438 0.00672 0.00146 -0.0885 0.6538 0.7654
1.500 0.5704 0.00680 0.00150 -0.0882 0.6368 0.7765
1.750 0.5966 0.00689 0.00156 -0.0878 0.6182 0.7883
2.000 0.6226 0.00698 0.00162 -0.0874 0.5978 0.8011
2.250 0.6481 0.00711 0.00172 -0.0869 0.5736 0.8154
2.500 0.6729 0.00726 0.00181 -0.0863 0.5452 0.8325
2.750 0.6968 0.00742 0.00192 -0.0855 0.5132 0.8539
3.000 0.7189 0.00757 0.00205 -0.0843 0.4788 0.8937
3.250 0.7503 0.00779 0.00221 -0.0852 0.4404 1.0000
3.500 0.7752 0.00813 0.00240 -0.0848 0.4083 1.0000
3.750 0.8001 0.00847 0.00260 -0.0844 0.3755 1.0000
4.000 0.8249 0.00881 0.00282 -0.0840 0.3439 1.0000
4.250 0.8494 0.00919 0.00306 -0.0835 0.3120 1.0000
4.500 0.8740 0.00956 0.00334 -0.0831 0.2842 1.0000
5.000 0.9218 0.01043 0.00392 -0.0821 0.2199 1.0000
5.250 0.9436 0.01107 0.00429 -0.0814 0.1707 1.0000
5.500 0.9661 0.01165 0.00466 -0.0807 0.1338 1.0000
5.750 0.9885 0.01223 0.00508 -0.0800 0.0999 1.0000
6.000 1.0102 0.01288 0.00554 -0.0792 0.0673 1.0000
6.250 1.0319 0.01353 0.00604 -0.0784 0.0409 1.0000
6.500 1.0532 0.01421 0.00659 -0.0775 0.0206 1.0000
6.750 1.0737 0.01500 0.00727 -0.0765 0.0053 1.0000
7.000 1.0963 0.01553 0.00783 -0.0757 0.0042 1.0000
7.250 1.1188 0.01604 0.00843 -0.0749 0.0035 1.0000
7.500 1.1410 0.01658 0.00908 -0.0741 0.0033 1.0000
7.750 1.1626 0.01717 0.00978 -0.0731 0.0031 1.0000
8.000 1.1836 0.01779 0.01050 -0.0722 0.0030 1.0000
8.250 1.2038 0.01848 0.01130 -0.0710 0.0029 1.0000
8.500 1.2227 0.01928 0.01221 -0.0697 0.0028 1.0000
8.750 1.2407 0.02012 0.01316 -0.0683 0.0028 1.0000
9.000 1.2572 0.02104 0.01421 -0.0667 0.0027 1.0000
9.250 1.2727 0.02200 0.01528 -0.0650 0.0026 1.0000
9.500 1.2864 0.02303 0.01642 -0.0631 0.0025 1.0000
9.750 1.2974 0.02406 0.01760 -0.0607 0.0023 1.0000
10.000 1.3068 0.02514 0.01878 -0.0582 0.0021 1.0000
10.250 1.3134 0.02643 0.02018 -0.0555 0.0020 1.0000
10.500 1.3182 0.02788 0.02175 -0.0527 0.0019 1.0000
10.750 1.3200 0.02963 0.02364 -0.0499 0.0018 1.0000
11.000 1.3190 0.03173 0.02590 -0.0470 0.0018 1.0000
11.250 1.3176 0.03400 0.02833 -0.0444 0.0017 1.0000
11.500 1.3168 0.03632 0.03082 -0.0423 0.0017 1.0000
11.750 1.3133 0.03905 0.03372 -0.0402 0.0017 1.0000
12.000 1.3137 0.04142 0.03627 -0.0388 0.0017 1.0000
12.250 1.3150 0.04371 0.03872 -0.0378 0.0016 1.0000
12.500 1.3120 0.04667 0.04187 -0.0369 0.0016 1.0000
12.750 1.3096 0.04966 0.04502 -0.0364 0.0016 1.0000
13.000 1.3045 0.05319 0.04874 -0.0363 0.0016 1.0000
13.250 1.2993 0.05692 0.05265 -0.0366 0.0015 1.0000
13.500 1.2887 0.06165 0.05757 -0.0376 0.0015 1.0000
13.750 1.2779 0.06676 0.06286 -0.0391 0.0015 1.0000
14.000 1.2667 0.07231 0.06860 -0.0413 0.0015 1.0000
14.250 1.2538 0.07861 0.07507 -0.0442 0.0015 1.0000
14.500 1.2404 0.08549 0.08213 -0.0478 0.0015 1.0000
14.750 1.2254 0.09316 0.08997 -0.0521 0.0015 1.0000
15.000 1.2098 0.10136 0.09833 -0.0567 0.0015 1.0000
15.250 1.1942 0.10999 0.10711 -0.0618 0.0015 1.0000
15.500 1.1790 0.11881 0.11608 -0.0669 0.0015 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.0/9 AIRFOIL (hq309-il)