HQ 2.5/8 AIRFOIL (hq258-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/8 AIRFOIL (hq258-il) Reynolds number: 50,000 Max Cl/Cd: 40.49 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq258-il-50000-n5.txt Download as CSV file: xf-hq258-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/8 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4265 0.11570 0.10867 -0.0215 1.0000 0.0896
-9.250 -0.4324 0.11372 0.10679 -0.0247 1.0000 0.0926
-9.000 -0.4410 0.11188 0.10509 -0.0282 1.0000 0.0934
-8.750 -0.4240 0.10618 0.09939 -0.0261 1.0000 0.0954
-8.500 -0.4164 0.10255 0.09580 -0.0260 1.0000 0.0979
-8.250 -0.4138 0.09935 0.09266 -0.0267 1.0000 0.1005
-7.750 -0.3351 0.07886 0.07275 -0.0334 1.0000 0.0626
-7.500 -0.3451 0.07465 0.06862 -0.0355 1.0000 0.0515
-7.250 -0.3454 0.07107 0.06512 -0.0345 1.0000 0.0486
-7.000 -0.4151 0.07940 0.07307 -0.0393 1.0000 0.0514
-6.750 -0.4119 0.07578 0.06951 -0.0388 1.0000 0.0478
-6.500 -0.4105 0.07160 0.06536 -0.0409 1.0000 0.0449
-6.250 -0.4060 0.06639 0.05990 -0.0460 1.0000 0.0395
-6.000 -0.4009 0.06293 0.05646 -0.0454 1.0000 0.0385
-5.750 -0.3946 0.05921 0.05267 -0.0458 1.0000 0.0375
-5.500 -0.3860 0.05552 0.04885 -0.0463 1.0000 0.0367
-5.250 -0.3747 0.05174 0.04486 -0.0469 1.0000 0.0359
-5.000 -0.3605 0.04797 0.04081 -0.0476 1.0000 0.0351
-4.750 -0.3435 0.04430 0.03679 -0.0481 1.0000 0.0346
-4.500 -0.3240 0.04077 0.03283 -0.0485 1.0000 0.0343
-4.250 -0.3023 0.03752 0.02911 -0.0487 1.0000 0.0343
-4.000 -0.2790 0.03455 0.02564 -0.0487 1.0000 0.0347
-3.750 -0.2546 0.03206 0.02261 -0.0484 1.0000 0.0368
-3.500 -0.2293 0.02998 0.01994 -0.0479 1.0000 0.0408
-3.250 -0.2065 0.02807 0.01786 -0.0475 1.0000 0.0455
-3.000 -0.1819 0.02631 0.01577 -0.0465 1.0000 0.0483
-2.750 -0.1579 0.02486 0.01386 -0.0451 1.0000 0.0524
-2.500 -0.1351 0.02355 0.01247 -0.0441 1.0000 0.0620
-2.250 -0.1109 0.02250 0.01130 -0.0436 1.0000 0.0831
-2.000 -0.0846 0.02130 0.01015 -0.0439 1.0000 0.1248
-1.750 -0.0570 0.01851 0.00945 -0.0448 1.0000 0.5260
-1.500 -0.0524 0.01796 0.00955 -0.0385 1.0000 0.7814
-1.250 -0.0164 0.01751 0.00903 -0.0394 0.9915 1.0000
-1.000 0.0227 0.01784 0.00883 -0.0424 0.9846 1.0000
-0.750 0.0600 0.01814 0.00865 -0.0450 0.9768 1.0000
-0.500 0.0969 0.01847 0.00863 -0.0475 0.9689 1.0000
-0.250 0.1355 0.01882 0.00869 -0.0503 0.9614 1.0000
0.000 0.1699 0.01912 0.00876 -0.0522 0.9523 1.0000
0.250 0.2067 0.01944 0.00883 -0.0545 0.9437 1.0000
0.500 0.2445 0.01975 0.00900 -0.0570 0.9351 1.0000
0.750 0.2784 0.02003 0.00917 -0.0586 0.9248 1.0000
1.000 0.3138 0.02031 0.00937 -0.0605 0.9147 1.0000
1.250 0.3520 0.02056 0.00957 -0.0627 0.9052 1.0000
1.500 0.3891 0.02079 0.00978 -0.0647 0.8950 1.0000
1.750 0.4231 0.02101 0.01001 -0.0661 0.8834 1.0000
2.000 0.4571 0.02120 0.01028 -0.0673 0.8716 1.0000
2.250 0.4919 0.02130 0.01044 -0.0684 0.8585 1.0000
2.500 0.5274 0.02129 0.01050 -0.0694 0.8438 1.0000
2.750 0.5628 0.02120 0.01052 -0.0701 0.8283 1.0000
3.000 0.5949 0.02114 0.01057 -0.0703 0.8117 1.0000
3.250 0.6235 0.02114 0.01080 -0.0698 0.7933 1.0000
3.500 0.6550 0.02101 0.01083 -0.0696 0.7752 1.0000
3.750 0.6828 0.02094 0.01092 -0.0687 0.7538 1.0000
4.000 0.7123 0.02077 0.01092 -0.0680 0.7314 1.0000
4.250 0.7396 0.02066 0.01098 -0.0668 0.7050 1.0000
4.500 0.7652 0.02063 0.01125 -0.0654 0.6738 1.0000
4.750 0.7919 0.02056 0.01132 -0.0640 0.6382 1.0000
5.000 0.8161 0.02064 0.01151 -0.0624 0.5940 1.0000
5.250 0.8400 0.02084 0.01170 -0.0606 0.5425 1.0000
5.500 0.8620 0.02129 0.01206 -0.0588 0.4849 1.0000
5.750 0.8821 0.02203 0.01264 -0.0569 0.4274 1.0000
6.000 0.9008 0.02299 0.01345 -0.0552 0.3742 1.0000
6.250 0.9189 0.02408 0.01444 -0.0536 0.3269 1.0000
6.500 0.9370 0.02526 0.01576 -0.0521 0.2849 1.0000
6.750 0.9548 0.02653 0.01705 -0.0506 0.2466 1.0000
7.000 0.9686 0.02789 0.01840 -0.0489 0.1950 1.0000
7.250 0.9750 0.02999 0.01965 -0.0473 0.0897 1.0000
7.500 0.9858 0.03293 0.02215 -0.0454 0.0459 1.0000
7.750 0.9993 0.03524 0.02453 -0.0437 0.0341 1.0000
8.000 1.0107 0.03765 0.02700 -0.0421 0.0281 1.0000
8.250 1.0257 0.03992 0.02964 -0.0403 0.0259 1.0000
8.500 1.0413 0.04253 0.03261 -0.0386 0.0241 1.0000
8.750 1.0570 0.04542 0.03586 -0.0370 0.0228 1.0000
9.000 1.0705 0.04862 0.03946 -0.0355 0.0218 1.0000
9.250 1.0800 0.05205 0.04328 -0.0339 0.0211 1.0000
9.500 1.0845 0.05565 0.04728 -0.0322 0.0207 1.0000
9.750 1.0826 0.05930 0.05129 -0.0302 0.0204 1.0000
10.000 1.0759 0.06302 0.05535 -0.0282 0.0203 1.0000
10.250 1.0652 0.06698 0.05962 -0.0269 0.0203 1.0000
10.500 1.0514 0.07135 0.06427 -0.0264 0.0204 1.0000
10.750 1.0353 0.07626 0.06944 -0.0271 0.0205 1.0000
11.000 1.0174 0.08192 0.07532 -0.0291 0.0207 1.0000
11.250 0.9982 0.08849 0.08208 -0.0325 0.0209 1.0000
11.500 0.9778 0.09639 0.09014 -0.0376 0.0213 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/8 AIRFOIL (hq258-il)