HQ 3.0/11 AIRFOIL (hq3011-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.0/11 AIRFOIL (hq3011-il) Reynolds number: 100,000 Max Cl/Cd: 57.45 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq3011-il-100000-n5.txt Download as CSV file: xf-hq3011-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.0/11 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3709 0.09141 0.08655 -0.0442 1.0000 0.0242
-9.000 -0.3772 0.08746 0.08268 -0.0451 1.0000 0.0239
-8.750 -0.3859 0.08361 0.07893 -0.0458 1.0000 0.0234
-8.500 -0.3995 0.07983 0.07528 -0.0461 1.0000 0.0230
-8.250 -0.4209 0.07638 0.07196 -0.0454 1.0000 0.0226
-8.000 -0.4512 0.07313 0.06883 -0.0440 1.0000 0.0222
-7.500 -0.4576 0.05356 0.04866 -0.0630 0.9835 0.0207
-7.250 -0.4450 0.04812 0.04287 -0.0672 0.9752 0.0210
-7.000 -0.4226 0.04547 0.04006 -0.0702 0.9697 0.0222
-6.750 -0.4012 0.04273 0.03707 -0.0723 0.9628 0.0237
-6.500 -0.3752 0.03828 0.03207 -0.0755 0.9581 0.0248
-6.250 -0.3544 0.03436 0.02758 -0.0764 0.9508 0.0253
-6.000 -0.3257 0.03078 0.02338 -0.0779 0.9463 0.0260
-5.750 -0.2925 0.02778 0.01978 -0.0796 0.9434 0.0274
-5.500 -0.2680 0.02584 0.01741 -0.0792 0.9365 0.0287
-5.250 -0.2369 0.02424 0.01568 -0.0805 0.9326 0.0317
-5.000 -0.2021 0.02312 0.01434 -0.0820 0.9296 0.0372
-4.750 -0.1777 0.02186 0.01298 -0.0815 0.9229 0.0416
-4.500 -0.1457 0.02110 0.01206 -0.0824 0.9186 0.0515
-4.250 -0.1119 0.02029 0.01122 -0.0839 0.9155 0.0645
-4.000 -0.0883 0.01966 0.01059 -0.0834 0.9082 0.0753
-3.750 -0.0572 0.01895 0.00984 -0.0841 0.9039 0.0861
-3.500 -0.0273 0.01836 0.00920 -0.0847 0.8992 0.0998
-3.250 -0.0012 0.01785 0.00869 -0.0845 0.8926 0.1167
-3.000 0.0303 0.01713 0.00815 -0.0856 0.8886 0.1634
-2.750 0.0529 0.01604 0.00781 -0.0854 0.8819 0.3339
-2.250 0.1076 0.01542 0.00791 -0.0846 0.8721 0.6080
-2.000 0.1297 0.01544 0.00793 -0.0832 0.8643 0.6483
-1.750 0.1572 0.01535 0.00790 -0.0824 0.8597 0.6973
-1.500 0.1762 0.01537 0.00804 -0.0799 0.8520 0.7413
-1.250 0.2013 0.01533 0.00801 -0.0785 0.8465 0.7740
-1.000 0.2259 0.01532 0.00795 -0.0775 0.8395 0.7919
-0.750 0.2538 0.01525 0.00781 -0.0772 0.8331 0.8060
-0.500 0.2812 0.01519 0.00768 -0.0769 0.8266 0.8206
-0.250 0.3078 0.01513 0.00756 -0.0764 0.8193 0.8369
0.000 0.3356 0.01504 0.00743 -0.0762 0.8128 0.8551
0.250 0.3619 0.01496 0.00733 -0.0757 0.8051 0.8749
0.500 0.3916 0.01483 0.00719 -0.0757 0.7987 0.8964
0.750 0.4253 0.01474 0.00709 -0.0767 0.7912 0.9209
1.000 0.4670 0.01461 0.00695 -0.0794 0.7847 0.9542
1.250 0.5048 0.01448 0.00676 -0.0814 0.7750 1.0000
1.500 0.5310 0.01446 0.00667 -0.0811 0.7632 1.0000
1.750 0.5581 0.01444 0.00658 -0.0809 0.7518 1.0000
2.000 0.5865 0.01442 0.00652 -0.0809 0.7419 1.0000
2.250 0.6141 0.01445 0.00651 -0.0808 0.7315 1.0000
2.500 0.6400 0.01453 0.00657 -0.0804 0.7197 1.0000
2.750 0.6666 0.01459 0.00661 -0.0800 0.7075 1.0000
3.000 0.6932 0.01464 0.00667 -0.0797 0.6945 1.0000
3.250 0.7197 0.01469 0.00671 -0.0792 0.6800 1.0000
3.500 0.7459 0.01475 0.00675 -0.0787 0.6640 1.0000
3.750 0.7720 0.01482 0.00682 -0.0782 0.6468 1.0000
4.000 0.7975 0.01493 0.00692 -0.0776 0.6283 1.0000
4.250 0.8220 0.01507 0.00707 -0.0768 0.6077 1.0000
4.500 0.8469 0.01522 0.00719 -0.0761 0.5858 1.0000
4.750 0.8711 0.01541 0.00738 -0.0753 0.5607 1.0000
5.000 0.8946 0.01566 0.00757 -0.0743 0.5329 1.0000
5.250 0.9174 0.01597 0.00779 -0.0733 0.5026 1.0000
5.500 0.9392 0.01636 0.00808 -0.0722 0.4708 1.0000
5.750 0.9603 0.01683 0.00846 -0.0710 0.4389 1.0000
6.000 0.9807 0.01736 0.00888 -0.0698 0.4083 1.0000
6.250 1.0004 0.01793 0.00937 -0.0685 0.3790 1.0000
6.500 1.0199 0.01854 0.00990 -0.0672 0.3517 1.0000
6.750 1.0387 0.01919 0.01047 -0.0659 0.3262 1.0000
7.000 1.0571 0.01987 0.01112 -0.0646 0.3033 1.0000
7.250 1.0755 0.02056 0.01178 -0.0632 0.2818 1.0000
7.500 1.0931 0.02130 0.01247 -0.0618 0.2615 1.0000
7.750 1.1104 0.02203 0.01320 -0.0604 0.2419 1.0000
8.000 1.1274 0.02276 0.01395 -0.0590 0.2228 1.0000
8.250 1.1432 0.02355 0.01473 -0.0575 0.2049 1.0000
8.500 1.1580 0.02437 0.01558 -0.0558 0.1881 1.0000
8.750 1.1721 0.02518 0.01643 -0.0540 0.1723 1.0000
9.000 1.1850 0.02605 0.01734 -0.0521 0.1575 1.0000
9.250 1.1971 0.02700 0.01832 -0.0502 0.1431 1.0000
9.500 1.2084 0.02801 0.01937 -0.0483 0.1278 1.0000
9.750 1.2185 0.02911 0.02047 -0.0464 0.1122 1.0000
10.000 1.2280 0.03030 0.02167 -0.0445 0.0985 1.0000
10.250 1.2373 0.03157 0.02300 -0.0428 0.0880 1.0000
10.500 1.2449 0.03298 0.02449 -0.0409 0.0795 1.0000
10.750 1.2519 0.03447 0.02602 -0.0393 0.0700 1.0000
11.000 1.2598 0.03596 0.02758 -0.0378 0.0601 1.0000
11.250 1.2660 0.03762 0.02933 -0.0363 0.0522 1.0000
11.500 1.2693 0.03956 0.03133 -0.0349 0.0467 1.0000
11.750 1.2745 0.04142 0.03335 -0.0336 0.0402 1.0000
12.000 1.2754 0.04374 0.03571 -0.0324 0.0339 1.0000
12.250 1.2762 0.04616 0.03830 -0.0313 0.0288 1.0000
12.500 1.2717 0.04921 0.04143 -0.0304 0.0255 1.0000
12.750 1.2687 0.05226 0.04470 -0.0297 0.0226 1.0000
13.000 1.2625 0.05581 0.04840 -0.0294 0.0203 1.0000
13.250 1.2537 0.05981 0.05253 -0.0296 0.0191 1.0000
13.500 1.2438 0.06418 0.05707 -0.0300 0.0181 1.0000
13.750 1.2346 0.06872 0.06183 -0.0309 0.0172 1.0000
14.000 1.2251 0.07355 0.06686 -0.0321 0.0165 1.0000
14.250 1.2146 0.07881 0.07232 -0.0338 0.0159 1.0000
14.500 1.2041 0.08434 0.07804 -0.0359 0.0156 1.0000
14.750 1.1929 0.09027 0.08420 -0.0385 0.0152 1.0000
15.000 1.1810 0.09661 0.09072 -0.0414 0.0148 1.0000
15.250 1.1689 0.10325 0.09753 -0.0447 0.0145 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.0/11 AIRFOIL (hq3011-il)