HQ 1.5/12 AIRFOIL (hq1512-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/12 AIRFOIL (hq1512-il) Reynolds number: 500,000 Max Cl/Cd: 73.2 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1512-il-500000-n5.txt Download as CSV file: xf-hq1512-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/12 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.250 -0.8258 0.09075 0.08786 -0.0307 1.0000 0.0065
-15.000 -0.8604 0.07921 0.07615 -0.0376 1.0000 0.0066
-14.750 -0.8885 0.06940 0.06618 -0.0438 1.0000 0.0065
-14.500 -0.9139 0.06066 0.05725 -0.0497 1.0000 0.0065
-14.250 -0.9283 0.05458 0.05101 -0.0536 1.0000 0.0065
-14.000 -0.9432 0.04901 0.04528 -0.0570 1.0000 0.0065
-13.750 -0.9536 0.04458 0.04068 -0.0592 1.0000 0.0066
-13.500 -0.9635 0.04061 0.03655 -0.0607 1.0000 0.0065
-13.250 -0.9651 0.03782 0.03360 -0.0614 1.0000 0.0067
-13.000 -0.9706 0.03491 0.03054 -0.0617 1.0000 0.0068
-12.750 -0.9762 0.03230 0.02782 -0.0613 1.0000 0.0070
-12.500 -0.9770 0.03034 0.02573 -0.0604 1.0000 0.0070
-12.250 -0.9748 0.02889 0.02419 -0.0589 1.0000 0.0072
-12.000 -0.9719 0.02771 0.02292 -0.0568 1.0000 0.0073
-11.750 -0.9683 0.02656 0.02167 -0.0542 1.0000 0.0075
-11.500 -0.9577 0.02559 0.02061 -0.0526 1.0000 0.0077
-11.250 -0.9460 0.02462 0.01955 -0.0510 1.0000 0.0079
-11.000 -0.9332 0.02369 0.01850 -0.0493 1.0000 0.0082
-10.750 -0.9197 0.02279 0.01749 -0.0477 1.0000 0.0085
-10.500 -0.9060 0.02190 0.01649 -0.0460 1.0000 0.0088
-10.250 -0.8908 0.02117 0.01566 -0.0443 1.0000 0.0091
-9.750 -0.8450 0.01930 0.01362 -0.0444 0.9908 0.0100
-9.500 -0.8140 0.01858 0.01284 -0.0459 0.9850 0.0106
-9.250 -0.7835 0.01788 0.01205 -0.0471 0.9779 0.0112
-9.000 -0.7519 0.01715 0.01123 -0.0486 0.9708 0.0119
-8.750 -0.7217 0.01648 0.01046 -0.0496 0.9613 0.0124
-8.500 -0.6908 0.01590 0.00978 -0.0507 0.9514 0.0129
-8.250 -0.6631 0.01507 0.00889 -0.0514 0.9391 0.0139
-8.000 -0.6354 0.01456 0.00831 -0.0517 0.9261 0.0149
-7.750 -0.6094 0.01411 0.00778 -0.0516 0.9136 0.0158
-7.500 -0.5845 0.01370 0.00726 -0.0512 0.9021 0.0166
-7.250 -0.5599 0.01330 0.00677 -0.0507 0.8915 0.0175
-7.000 -0.5359 0.01285 0.00626 -0.0501 0.8815 0.0192
-6.750 -0.5110 0.01251 0.00587 -0.0497 0.8720 0.0211
-6.500 -0.4856 0.01222 0.00548 -0.0493 0.8635 0.0229
-6.250 -0.4608 0.01185 0.00508 -0.0488 0.8552 0.0263
-6.000 -0.4350 0.01157 0.00476 -0.0485 0.8473 0.0300
-5.750 -0.4096 0.01126 0.00444 -0.0481 0.8399 0.0362
-5.500 -0.3837 0.01098 0.00414 -0.0478 0.8322 0.0436
-5.250 -0.3577 0.01074 0.00387 -0.0476 0.8251 0.0523
-5.000 -0.3314 0.01048 0.00362 -0.0473 0.8178 0.0625
-4.750 -0.3053 0.01022 0.00338 -0.0471 0.8108 0.0764
-4.500 -0.2791 0.00994 0.00315 -0.0469 0.8038 0.0952
-4.250 -0.2529 0.00967 0.00294 -0.0467 0.7969 0.1183
-4.000 -0.2267 0.00937 0.00274 -0.0465 0.7901 0.1486
-3.750 -0.2006 0.00907 0.00255 -0.0463 0.7832 0.1861
-3.500 -0.1740 0.00883 0.00238 -0.0461 0.7767 0.2177
-3.250 -0.1476 0.00856 0.00221 -0.0460 0.7697 0.2542
-3.000 -0.1215 0.00824 0.00203 -0.0458 0.7631 0.3035
-2.750 -0.0961 0.00782 0.00187 -0.0456 0.7559 0.3815
-2.500 -0.0703 0.00751 0.00176 -0.0453 0.7493 0.4484
-2.250 -0.0437 0.00731 0.00170 -0.0451 0.7419 0.4997
-2.000 -0.0166 0.00722 0.00165 -0.0449 0.7352 0.5314
-1.750 0.0111 0.00715 0.00161 -0.0448 0.7273 0.5548
-1.500 0.0386 0.00712 0.00158 -0.0446 0.7187 0.5746
-1.250 0.0660 0.00709 0.00156 -0.0444 0.7083 0.5946
-1.000 0.0937 0.00708 0.00154 -0.0443 0.6992 0.6114
-0.750 0.1214 0.00709 0.00153 -0.0442 0.6906 0.6241
-0.500 0.1492 0.00708 0.00153 -0.0441 0.6812 0.6378
-0.250 0.1769 0.00709 0.00153 -0.0439 0.6726 0.6502
0.000 0.2047 0.00713 0.00153 -0.0438 0.6632 0.6613
0.250 0.2324 0.00713 0.00155 -0.0437 0.6526 0.6708
0.500 0.2601 0.00716 0.00156 -0.0436 0.6414 0.6797
0.750 0.2877 0.00720 0.00158 -0.0435 0.6306 0.6890
1.000 0.3153 0.00724 0.00161 -0.0434 0.6196 0.6963
1.250 0.3430 0.00728 0.00164 -0.0433 0.6072 0.7030
1.500 0.3706 0.00733 0.00168 -0.0432 0.5953 0.7093
1.750 0.3981 0.00739 0.00172 -0.0430 0.5823 0.7161
2.000 0.4253 0.00747 0.00177 -0.0429 0.5660 0.7228
2.250 0.4521 0.00756 0.00184 -0.0426 0.5474 0.7298
2.500 0.4788 0.00768 0.00191 -0.0424 0.5278 0.7371
2.750 0.5049 0.00782 0.00201 -0.0420 0.5026 0.7443
3.000 0.5307 0.00802 0.00212 -0.0416 0.4741 0.7519
3.250 0.5557 0.00825 0.00226 -0.0411 0.4417 0.7593
3.500 0.5804 0.00854 0.00243 -0.0406 0.4037 0.7672
3.750 0.6042 0.00889 0.00263 -0.0400 0.3607 0.7753
4.000 0.6276 0.00929 0.00285 -0.0393 0.3152 0.7836
4.250 0.6506 0.00971 0.00311 -0.0386 0.2764 0.7927
4.500 0.6745 0.01004 0.00336 -0.0380 0.2500 0.8018
4.750 0.6995 0.01027 0.00356 -0.0375 0.2344 0.8118
5.000 0.7245 0.01047 0.00377 -0.0371 0.2227 0.8229
5.250 0.7494 0.01064 0.00398 -0.0366 0.2107 0.8346
5.500 0.7740 0.01083 0.00420 -0.0360 0.1985 0.8481
5.750 0.7978 0.01102 0.00442 -0.0353 0.1846 0.8645
6.000 0.8204 0.01126 0.00467 -0.0343 0.1659 0.8881
6.250 0.8469 0.01157 0.00498 -0.0343 0.1408 0.9304
6.500 0.8828 0.01208 0.00538 -0.0365 0.1135 1.0000
6.750 0.9049 0.01257 0.00577 -0.0358 0.0934 1.0000
7.000 0.9267 0.01307 0.00617 -0.0350 0.0763 1.0000
7.250 0.9488 0.01353 0.00658 -0.0342 0.0646 1.0000
7.500 0.9711 0.01397 0.00699 -0.0334 0.0567 1.0000
7.750 0.9927 0.01446 0.00743 -0.0326 0.0486 1.0000
8.000 1.0151 0.01485 0.00783 -0.0319 0.0442 1.0000
8.250 1.0365 0.01531 0.00827 -0.0310 0.0398 1.0000
8.500 1.0580 0.01574 0.00873 -0.0301 0.0369 1.0000
8.750 1.0798 0.01613 0.00917 -0.0293 0.0351 1.0000
9.000 1.1006 0.01657 0.00963 -0.0284 0.0329 1.0000
9.250 1.1203 0.01708 0.01014 -0.0273 0.0303 1.0000
9.500 1.1397 0.01758 0.01067 -0.0262 0.0285 1.0000
9.750 1.1595 0.01801 0.01116 -0.0251 0.0270 1.0000
10.000 1.1780 0.01848 0.01168 -0.0238 0.0250 1.0000
10.250 1.1936 0.01901 0.01222 -0.0221 0.0226 1.0000
10.500 1.2087 0.01956 0.01279 -0.0203 0.0206 1.0000
10.750 1.2235 0.02012 0.01339 -0.0186 0.0187 1.0000
11.000 1.2368 0.02079 0.01406 -0.0168 0.0164 1.0000
11.250 1.2502 0.02147 0.01479 -0.0150 0.0146 1.0000
11.500 1.2620 0.02227 0.01562 -0.0132 0.0123 1.0000
11.750 1.2716 0.02325 0.01661 -0.0113 0.0083 1.0000
12.000 1.2792 0.02440 0.01780 -0.0093 0.0060 1.0000
12.250 1.2870 0.02559 0.01905 -0.0075 0.0053 1.0000
12.500 1.2956 0.02676 0.02029 -0.0060 0.0050 1.0000
12.750 1.3022 0.02814 0.02176 -0.0044 0.0046 1.0000
13.000 1.3089 0.02956 0.02327 -0.0030 0.0043 1.0000
13.250 1.3155 0.03104 0.02486 -0.0018 0.0041 1.0000
13.500 1.3220 0.03259 0.02650 -0.0007 0.0041 1.0000
13.750 1.3264 0.03438 0.02841 0.0003 0.0039 1.0000
14.000 1.3295 0.03636 0.03051 0.0011 0.0038 1.0000
14.250 1.3323 0.03845 0.03272 0.0018 0.0037 1.0000
14.500 1.3338 0.04075 0.03514 0.0023 0.0036 1.0000
14.750 1.3324 0.04346 0.03798 0.0025 0.0034 1.0000
15.000 1.3322 0.04614 0.04079 0.0026 0.0034 1.0000
15.250 1.3288 0.04931 0.04408 0.0023 0.0033 1.0000
15.500 1.3239 0.05283 0.04773 0.0017 0.0033 1.0000
15.750 1.3153 0.05703 0.05208 0.0007 0.0032 1.0000
16.000 1.3056 0.06166 0.05686 -0.0008 0.0031 1.0000
16.250 1.2989 0.06610 0.06143 -0.0025 0.0032 1.0000
16.500 1.2855 0.07184 0.06734 -0.0050 0.0031 1.0000
16.750 1.2693 0.07842 0.07407 -0.0082 0.0031 1.0000
17.000 1.2572 0.08452 0.08032 -0.0113 0.0031 1.0000
17.250 1.2408 0.09172 0.08768 -0.0150 0.0031 1.0000
17.500 1.2165 0.10076 0.09690 -0.0200 0.0030 1.0000
17.750 1.1999 0.10844 0.10472 -0.0241 0.0031 1.0000
18.000 1.1832 0.11631 0.11272 -0.0285 0.0031 1.0000
18.250 1.1622 0.12520 0.12176 -0.0335 0.0031 1.0000
18.500 1.1399 0.13461 0.13131 -0.0389 0.0032 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/12 AIRFOIL (hq1512-il)