HQ 1.5/8.5 AIRFOIL (hq1585-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/8.5 AIRFOIL (hq1585-il) Reynolds number: 50,000 Max Cl/Cd: 35.64 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1585-il-50000-n5.txt Download as CSV file: xf-hq1585-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8.5 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5432 0.09249 0.08561 -0.0261 1.0000 0.0489
-9.000 -0.5422 0.08775 0.08092 -0.0283 1.0000 0.0486
-8.750 -0.5430 0.08288 0.07610 -0.0306 1.0000 0.0482
-8.500 -0.5458 0.07843 0.07169 -0.0330 1.0000 0.0477
-8.250 -0.5499 0.07372 0.06700 -0.0357 1.0000 0.0472
-8.000 -0.5533 0.06896 0.06218 -0.0381 1.0000 0.0468
-7.750 -0.5557 0.06414 0.05724 -0.0402 1.0000 0.0464
-7.500 -0.5555 0.05946 0.05236 -0.0418 1.0000 0.0462
-7.250 -0.5522 0.05486 0.04748 -0.0428 1.0000 0.0460
-7.000 -0.5455 0.05047 0.04272 -0.0434 1.0000 0.0461
-6.750 -0.5354 0.04639 0.03820 -0.0434 1.0000 0.0466
-6.500 -0.5226 0.04284 0.03393 -0.0431 1.0000 0.0487
-6.250 -0.5081 0.03978 0.03080 -0.0427 1.0000 0.0517
-6.000 -0.4909 0.03707 0.02777 -0.0419 1.0000 0.0538
-5.750 -0.4720 0.03432 0.02458 -0.0410 1.0000 0.0552
-5.500 -0.4517 0.03189 0.02173 -0.0398 1.0000 0.0570
-5.250 -0.4305 0.02980 0.01924 -0.0386 1.0000 0.0595
-5.000 -0.4100 0.02809 0.01732 -0.0374 1.0000 0.0646
-4.750 -0.3897 0.02671 0.01575 -0.0362 1.0000 0.0706
-4.500 -0.3686 0.02537 0.01414 -0.0345 1.0000 0.0745
-4.250 -0.3494 0.02407 0.01285 -0.0330 1.0000 0.0793
-4.000 -0.3299 0.02303 0.01166 -0.0314 1.0000 0.0868
-3.750 -0.3104 0.02204 0.01067 -0.0303 1.0000 0.1027
-3.500 -0.2903 0.02093 0.00967 -0.0295 1.0000 0.1265
-3.250 -0.2705 0.01938 0.00875 -0.0291 1.0000 0.2161
-3.000 -0.2549 0.01777 0.00837 -0.0275 1.0000 0.4500
-2.750 -0.2432 0.01737 0.00857 -0.0235 1.0000 0.6144
-2.500 -0.2348 0.01735 0.00880 -0.0184 1.0000 0.7283
-2.250 -0.2272 0.01725 0.00880 -0.0130 1.0000 0.8099
-2.000 -0.2103 0.01708 0.00859 -0.0099 1.0000 0.8640
-1.750 -0.1760 0.01696 0.00828 -0.0108 1.0000 0.9151
-1.500 -0.1234 0.01685 0.00786 -0.0161 1.0000 0.9814
-1.250 -0.1040 0.01681 0.00757 -0.0162 1.0000 1.0000
-1.000 -0.0727 0.01695 0.00746 -0.0183 0.9942 1.0000
-0.750 -0.0349 0.01718 0.00741 -0.0213 0.9861 1.0000
-0.500 0.0020 0.01742 0.00744 -0.0241 0.9776 1.0000
-0.250 0.0385 0.01767 0.00752 -0.0267 0.9688 1.0000
0.000 0.0759 0.01794 0.00764 -0.0294 0.9602 1.0000
0.250 0.1150 0.01823 0.00781 -0.0323 0.9514 1.0000
0.500 0.1512 0.01848 0.00798 -0.0345 0.9407 1.0000
0.750 0.1891 0.01873 0.00818 -0.0370 0.9302 1.0000
1.000 0.2293 0.01897 0.00840 -0.0397 0.9200 1.0000
1.250 0.2700 0.01919 0.00863 -0.0425 0.9098 1.0000
1.500 0.3073 0.01936 0.00885 -0.0443 0.8968 1.0000
1.750 0.3452 0.01947 0.00901 -0.0461 0.8824 1.0000
2.000 0.3822 0.01953 0.00914 -0.0474 0.8674 1.0000
2.250 0.4144 0.01960 0.00932 -0.0479 0.8509 1.0000
2.500 0.4454 0.01964 0.00946 -0.0479 0.8332 1.0000
2.750 0.4771 0.01963 0.00955 -0.0478 0.8149 1.0000
3.000 0.5088 0.01958 0.00966 -0.0476 0.7969 1.0000
3.250 0.5342 0.01967 0.00988 -0.0466 0.7755 1.0000
3.500 0.5625 0.01965 0.01000 -0.0457 0.7549 1.0000
3.750 0.5868 0.01971 0.01022 -0.0444 0.7296 1.0000
4.000 0.6115 0.01970 0.01043 -0.0429 0.7017 1.0000
4.250 0.6365 0.01964 0.01050 -0.0413 0.6691 1.0000
4.500 0.6612 0.01957 0.01051 -0.0395 0.6286 1.0000
4.750 0.6847 0.01963 0.01058 -0.0376 0.5778 1.0000
5.000 0.7074 0.01987 0.01073 -0.0356 0.5192 1.0000
5.250 0.7274 0.02041 0.01102 -0.0334 0.4481 1.0000
5.500 0.7435 0.02139 0.01149 -0.0313 0.3648 1.0000
5.750 0.7592 0.02257 0.01223 -0.0297 0.2902 1.0000
6.000 0.7769 0.02374 0.01318 -0.0285 0.2290 1.0000
6.250 0.7946 0.02509 0.01432 -0.0274 0.1735 1.0000
6.500 0.8122 0.02667 0.01567 -0.0263 0.1354 1.0000
6.750 0.8316 0.02829 0.01726 -0.0252 0.1148 1.0000
7.000 0.8526 0.02992 0.01894 -0.0241 0.1022 1.0000
7.250 0.8737 0.03150 0.02068 -0.0233 0.0903 1.0000
7.500 0.8956 0.03330 0.02248 -0.0224 0.0823 1.0000
7.750 0.9202 0.03535 0.02487 -0.0216 0.0746 1.0000
8.000 0.9390 0.03727 0.02698 -0.0207 0.0642 1.0000
8.250 0.9567 0.03952 0.02946 -0.0197 0.0555 1.0000
8.500 0.9740 0.04209 0.03209 -0.0189 0.0497 1.0000
8.750 0.9885 0.04548 0.03610 -0.0175 0.0450 1.0000
9.000 0.9992 0.04844 0.03945 -0.0163 0.0408 1.0000
9.250 1.0079 0.05152 0.04274 -0.0152 0.0382 1.0000
9.500 1.0128 0.05560 0.04717 -0.0140 0.0369 1.0000
9.750 1.0104 0.05986 0.05182 -0.0126 0.0363 1.0000
10.000 1.0024 0.06397 0.05635 -0.0111 0.0360 1.0000
10.250 0.9890 0.06792 0.06060 -0.0097 0.0359 1.0000
10.500 0.9726 0.07208 0.06499 -0.0091 0.0359 1.0000
10.750 0.9546 0.07681 0.06992 -0.0100 0.0360 1.0000
11.000 0.9367 0.08213 0.07538 -0.0121 0.0361 1.0000
11.250 0.9178 0.08835 0.08173 -0.0156 0.0363 1.0000
11.500 0.9006 0.09528 0.08865 -0.0199 0.0365 1.0000
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Polar data table (+)
Polar graphs
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