HQ 1.5/11 AIRFOIL (hq1511-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/11 AIRFOIL (hq1511-il) Reynolds number: 100,000 Max Cl/Cd: 48.41 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1511-il-100000-n5.txt Download as CSV file: xf-hq1511-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/11 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5639 0.08449 0.07940 -0.0336 1.0000 0.0212
-10.500 -0.5789 0.07560 0.07053 -0.0395 1.0000 0.0209
-10.250 -0.6015 0.06694 0.06184 -0.0465 1.0000 0.0206
-10.000 -0.6245 0.06060 0.05540 -0.0510 1.0000 0.0204
-9.750 -0.6463 0.05600 0.05068 -0.0528 1.0000 0.0203
-9.500 -0.6648 0.05228 0.04678 -0.0521 1.0000 0.0202
-9.250 -0.6780 0.04819 0.04240 -0.0511 1.0000 0.0202
-9.000 -0.6856 0.04419 0.03804 -0.0497 1.0000 0.0204
-8.750 -0.6871 0.04055 0.03400 -0.0480 1.0000 0.0206
-8.500 -0.6847 0.03709 0.03007 -0.0460 1.0000 0.0210
-8.250 -0.6776 0.03416 0.02668 -0.0440 1.0000 0.0214
-8.000 -0.6669 0.03216 0.02449 -0.0422 1.0000 0.0222
-7.750 -0.6541 0.03082 0.02304 -0.0405 1.0000 0.0231
-7.500 -0.6409 0.02962 0.02167 -0.0387 1.0000 0.0248
-7.250 -0.6276 0.02803 0.01978 -0.0367 1.0000 0.0271
-7.000 -0.6145 0.02660 0.01822 -0.0347 1.0000 0.0290
-6.750 -0.6010 0.02559 0.01715 -0.0328 1.0000 0.0313
-6.500 -0.5825 0.02425 0.01558 -0.0315 0.9987 0.0353
-6.250 -0.5504 0.02303 0.01429 -0.0331 0.9925 0.0409
-6.000 -0.5170 0.02239 0.01354 -0.0350 0.9860 0.0507
-5.750 -0.4840 0.02173 0.01277 -0.0368 0.9799 0.0612
-5.500 -0.4514 0.02083 0.01174 -0.0381 0.9737 0.0681
-5.250 -0.4187 0.01992 0.01079 -0.0396 0.9683 0.0754
-5.000 -0.3873 0.01924 0.00998 -0.0406 0.9616 0.0838
-4.750 -0.3531 0.01842 0.00918 -0.0425 0.9569 0.0948
-4.500 -0.3237 0.01773 0.00848 -0.0432 0.9495 0.1103
-4.000 -0.2640 0.01612 0.00741 -0.0453 0.9367 0.2170
-3.750 -0.2364 0.01517 0.00699 -0.0460 0.9301 0.3269
-3.500 -0.2107 0.01453 0.00687 -0.0458 0.9232 0.4605
-3.250 -0.1837 0.01431 0.00693 -0.0454 0.9163 0.5463
-3.000 -0.1541 0.01428 0.00699 -0.0453 0.9106 0.6068
-2.750 -0.1282 0.01435 0.00706 -0.0445 0.9028 0.6519
-2.500 -0.0981 0.01442 0.00711 -0.0443 0.8978 0.6898
-2.250 -0.0733 0.01448 0.00713 -0.0433 0.8897 0.7120
-2.000 -0.0422 0.01444 0.00699 -0.0436 0.8840 0.7263
-1.750 -0.0157 0.01442 0.00691 -0.0431 0.8754 0.7407
-1.500 0.0142 0.01437 0.00680 -0.0429 0.8683 0.7583
-1.250 0.0389 0.01436 0.00675 -0.0418 0.8583 0.7751
-1.000 0.0669 0.01432 0.00667 -0.0413 0.8514 0.7896
-0.750 0.0925 0.01432 0.00664 -0.0406 0.8432 0.8021
-0.500 0.1210 0.01427 0.00656 -0.0404 0.8373 0.8140
-0.250 0.1461 0.01427 0.00654 -0.0397 0.8285 0.8257
0.000 0.1740 0.01422 0.00645 -0.0395 0.8213 0.8363
0.250 0.2003 0.01418 0.00641 -0.0389 0.8122 0.8461
0.500 0.2269 0.01413 0.00635 -0.0384 0.8028 0.8568
0.750 0.2552 0.01402 0.00621 -0.0380 0.7935 0.8682
1.000 0.2813 0.01394 0.00616 -0.0373 0.7812 0.8801
1.250 0.3090 0.01385 0.00608 -0.0368 0.7685 0.8927
1.500 0.3385 0.01375 0.00600 -0.0368 0.7556 0.9066
1.750 0.3704 0.01368 0.00594 -0.0372 0.7434 0.9214
2.000 0.4048 0.01362 0.00591 -0.0383 0.7314 0.9364
2.250 0.4414 0.01357 0.00588 -0.0399 0.7196 0.9514
2.500 0.4787 0.01356 0.00593 -0.0419 0.7066 0.9675
2.750 0.5166 0.01354 0.00597 -0.0440 0.6922 0.9863
3.000 0.5463 0.01357 0.00601 -0.0446 0.6768 1.0000
3.250 0.5679 0.01364 0.00607 -0.0435 0.6608 1.0000
3.500 0.5909 0.01374 0.00619 -0.0426 0.6433 1.0000
3.750 0.6145 0.01385 0.00633 -0.0418 0.6237 1.0000
4.000 0.6386 0.01396 0.00642 -0.0410 0.6015 1.0000
4.250 0.6624 0.01411 0.00655 -0.0401 0.5739 1.0000
4.500 0.6853 0.01429 0.00665 -0.0390 0.5348 1.0000
4.750 0.7068 0.01460 0.00672 -0.0376 0.4856 1.0000
5.000 0.7275 0.01509 0.00694 -0.0363 0.4392 1.0000
5.250 0.7489 0.01559 0.00731 -0.0352 0.3990 1.0000
5.500 0.7703 0.01612 0.00774 -0.0342 0.3602 1.0000
5.750 0.7912 0.01670 0.00817 -0.0332 0.3273 1.0000
6.000 0.8120 0.01732 0.00867 -0.0323 0.3005 1.0000
6.250 0.8328 0.01796 0.00923 -0.0313 0.2770 1.0000
6.500 0.8540 0.01858 0.00984 -0.0304 0.2530 1.0000
6.750 0.8745 0.01923 0.01046 -0.0295 0.2282 1.0000
7.000 0.8945 0.01991 0.01110 -0.0285 0.1998 1.0000
7.250 0.9144 0.02059 0.01172 -0.0276 0.1626 1.0000
7.500 0.9319 0.02152 0.01246 -0.0264 0.1295 1.0000
7.750 0.9481 0.02261 0.01340 -0.0251 0.1065 1.0000
8.000 0.9648 0.02368 0.01445 -0.0238 0.0893 1.0000
8.250 0.9807 0.02481 0.01556 -0.0224 0.0756 1.0000
8.500 0.9952 0.02602 0.01677 -0.0209 0.0660 1.0000
8.750 1.0095 0.02721 0.01800 -0.0194 0.0583 1.0000
9.000 1.0232 0.02844 0.01935 -0.0177 0.0521 1.0000
9.250 1.0356 0.02969 0.02067 -0.0160 0.0479 1.0000
9.500 1.0447 0.03119 0.02220 -0.0140 0.0451 1.0000
9.750 1.0566 0.03254 0.02373 -0.0121 0.0421 1.0000
10.000 1.0670 0.03389 0.02517 -0.0104 0.0392 1.0000
10.250 1.0754 0.03542 0.02673 -0.0087 0.0371 1.0000
10.500 1.0856 0.03726 0.02868 -0.0071 0.0355 1.0000
10.750 1.0975 0.03911 0.03078 -0.0057 0.0341 1.0000
11.000 1.1075 0.04104 0.03293 -0.0043 0.0326 1.0000
11.250 1.1149 0.04297 0.03503 -0.0030 0.0311 1.0000
11.500 1.1204 0.04490 0.03707 -0.0017 0.0298 1.0000
11.750 1.1247 0.04705 0.03929 -0.0006 0.0287 1.0000
12.000 1.1276 0.04981 0.04219 0.0004 0.0278 1.0000
12.250 1.1269 0.05280 0.04550 0.0014 0.0273 1.0000
12.500 1.1227 0.05612 0.04912 0.0020 0.0269 1.0000
12.750 1.1147 0.05989 0.05319 0.0023 0.0265 1.0000
13.000 1.1037 0.06408 0.05766 0.0020 0.0262 1.0000
13.250 1.0902 0.06877 0.06262 0.0010 0.0260 1.0000
13.500 1.0734 0.07415 0.06825 -0.0008 0.0258 1.0000
13.750 1.0539 0.08036 0.07471 -0.0037 0.0258 1.0000
14.000 1.0321 0.08749 0.08207 -0.0077 0.0258 1.0000
14.250 1.0072 0.09603 0.09082 -0.0130 0.0261 1.0000
14.500 0.9796 0.10611 0.10108 -0.0196 0.0265 1.0000
14.750 0.9477 0.11859 0.11370 -0.0280 0.0269 1.0000
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