HQ 2.0/9 AIRFOIL (hq209-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.0/9 AIRFOIL (hq209-il) Reynolds number: 50,000 Max Cl/Cd: 36.9 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq209-il-50000.txt Download as CSV file: xf-hq209-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.0/9 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4447 0.10784 0.10099 -0.0129 1.0000 0.2364
-8.750 -0.4624 0.10705 0.10034 -0.0148 1.0000 0.2463
-8.500 -0.4529 0.10324 0.09656 -0.0141 1.0000 0.2599
-8.250 -0.4501 0.10006 0.09346 -0.0136 1.0000 0.2738
-8.000 -0.4513 0.09734 0.09083 -0.0132 1.0000 0.2881
-7.750 -0.4273 0.09235 0.08582 -0.0113 1.0000 0.3056
-7.500 -0.4193 0.08919 0.08271 -0.0102 1.0000 0.3236
-7.250 -0.4336 0.08788 0.08155 -0.0091 1.0000 0.3433
-7.000 -0.4225 0.08469 0.07841 -0.0067 1.0000 0.3706
-6.750 -0.4001 0.08076 0.07448 -0.0040 1.0000 0.4038
-6.500 -0.3908 0.07787 0.07164 -0.0017 1.0000 0.4325
-5.750 -0.3850 0.07028 0.06432 0.0055 1.0000 0.4985
-5.500 -0.4520 0.04941 0.04217 -0.0425 1.0000 0.1338
-5.250 -0.4357 0.04472 0.03696 -0.0427 1.0000 0.1194
-5.000 -0.4192 0.04110 0.03301 -0.0422 1.0000 0.1164
-4.750 -0.4006 0.03790 0.02934 -0.0416 1.0000 0.1160
-4.500 -0.3797 0.03489 0.02582 -0.0409 1.0000 0.1161
-4.250 -0.3566 0.03205 0.02246 -0.0401 1.0000 0.1157
-4.000 -0.3325 0.02961 0.01953 -0.0390 1.0000 0.1183
-3.750 -0.3099 0.02764 0.01734 -0.0380 1.0000 0.1309
-3.500 -0.2859 0.02570 0.01522 -0.0367 1.0000 0.1446
-3.250 -0.2633 0.02403 0.01358 -0.0354 1.0000 0.1725
-3.000 -0.2404 0.02232 0.01197 -0.0339 1.0000 0.2111
-2.750 -0.2162 0.01973 0.01030 -0.0333 1.0000 0.3294
-2.500 -0.2303 0.01827 0.01115 -0.0206 1.0000 0.7763
-2.250 -0.1254 0.01847 0.01061 -0.0271 1.0000 1.0000
-2.000 -0.1360 0.01806 0.01010 -0.0229 1.0000 1.0000
-1.750 -0.1333 0.01781 0.00962 -0.0206 1.0000 1.0000
-1.500 -0.1186 0.01776 0.00921 -0.0201 1.0000 1.0000
-1.250 -0.0993 0.01783 0.00897 -0.0201 1.0000 1.0000
-1.000 -0.0783 0.01798 0.00882 -0.0202 1.0000 1.0000
-0.750 -0.0568 0.01819 0.00877 -0.0203 1.0000 1.0000
-0.500 -0.0354 0.01845 0.00879 -0.0204 1.0000 1.0000
-0.250 -0.0139 0.01876 0.00886 -0.0205 1.0000 1.0000
0.000 0.0073 0.01911 0.00902 -0.0205 1.0000 1.0000
0.250 0.0283 0.01950 0.00925 -0.0205 1.0000 1.0000
0.500 0.0490 0.01994 0.00955 -0.0205 1.0000 1.0000
0.750 0.0694 0.02041 0.00988 -0.0205 1.0000 1.0000
1.000 0.0896 0.02093 0.01030 -0.0205 1.0000 1.0000
1.250 0.1094 0.02150 0.01078 -0.0206 1.0000 1.0000
1.500 0.1288 0.02211 0.01133 -0.0206 1.0000 1.0000
1.750 0.1479 0.02278 0.01195 -0.0207 1.0000 1.0000
2.000 0.1666 0.02350 0.01264 -0.0208 1.0000 1.0000
2.250 0.1891 0.02436 0.01350 -0.0217 0.9978 1.0000
2.500 0.2511 0.02599 0.01518 -0.0299 0.9748 1.0000
2.750 0.3126 0.02732 0.01662 -0.0374 0.9489 1.0000
3.000 0.3701 0.02835 0.01777 -0.0435 0.9243 1.0000
3.250 0.4250 0.02914 0.01873 -0.0488 0.9007 1.0000
3.500 0.4706 0.02972 0.01954 -0.0522 0.8758 1.0000
3.750 0.5194 0.03010 0.02015 -0.0555 0.8501 1.0000
4.000 0.5712 0.03016 0.02049 -0.0587 0.8232 1.0000
4.250 0.6272 0.02973 0.02046 -0.0616 0.7950 1.0000
4.500 0.6799 0.02882 0.01989 -0.0628 0.7653 1.0000
4.750 0.7325 0.02733 0.01874 -0.0628 0.7341 1.0000
5.000 0.7689 0.02620 0.01793 -0.0606 0.6953 1.0000
5.250 0.8059 0.02483 0.01672 -0.0579 0.6521 1.0000
5.500 0.8330 0.02407 0.01597 -0.0544 0.6004 1.0000
5.750 0.8572 0.02362 0.01538 -0.0509 0.5437 1.0000
6.000 0.8768 0.02376 0.01533 -0.0475 0.4827 1.0000
6.250 0.8937 0.02448 0.01571 -0.0444 0.4188 1.0000
6.500 0.9080 0.02577 0.01657 -0.0414 0.3492 1.0000
6.750 0.9219 0.02775 0.01796 -0.0387 0.2781 1.0000
7.000 0.9428 0.03033 0.02024 -0.0371 0.2256 1.0000
7.250 0.9634 0.03242 0.02224 -0.0357 0.1928 1.0000
7.500 0.9871 0.03468 0.02455 -0.0348 0.1713 1.0000
7.750 1.0119 0.03737 0.02738 -0.0339 0.1568 1.0000
8.000 1.0318 0.03981 0.03000 -0.0328 0.1415 1.0000
8.250 1.0495 0.04311 0.03388 -0.0312 0.1325 1.0000
8.500 1.0649 0.04730 0.03850 -0.0296 0.1265 1.0000
8.750 1.0782 0.05058 0.04193 -0.0281 0.1148 1.0000
9.000 1.0766 0.05513 0.04726 -0.0254 0.1110 1.0000
9.250 1.0870 0.05883 0.05098 -0.0240 0.1016 1.0000
9.500 1.0768 0.06341 0.05615 -0.0216 0.0998 1.0000
9.750 1.0628 0.06821 0.06138 -0.0196 0.0989 1.0000
10.000 1.0470 0.07308 0.06654 -0.0182 0.0994 1.0000
10.250 1.0287 0.07782 0.07147 -0.0170 0.1003 1.0000
10.500 1.0107 0.08268 0.07643 -0.0164 0.1012 1.0000
10.750 0.9939 0.08807 0.08190 -0.0170 0.1020 1.0000
11.000 0.9813 0.09401 0.08791 -0.0183 0.1028 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.0/9 AIRFOIL (hq209-il)