HQ 1.5/8 AIRFOIL (hq158-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/8 AIRFOIL (hq158-il) Reynolds number: 1,000,000 Max Cl/Cd: 77.48 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq158-il-1000000-n5.txt Download as CSV file: xf-hq158-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5263 0.08272 0.08116 -0.0183 1.0000 0.0025
-8.750 -0.5322 0.07736 0.07583 -0.0215 1.0000 0.0024
-8.500 -0.5365 0.07266 0.07116 -0.0246 1.0000 0.0024
-8.250 -0.5485 0.06703 0.06556 -0.0296 1.0000 0.0024
-8.000 -0.5557 0.05923 0.05769 -0.0377 1.0000 0.0023
-7.750 -0.5590 0.05263 0.05097 -0.0411 1.0000 0.0023
-7.500 -0.5525 0.04459 0.04269 -0.0454 0.9887 0.0022
-7.250 -0.5526 0.02411 0.02108 -0.0505 0.9647 0.0020
-7.000 -0.5347 0.01797 0.01413 -0.0504 0.9476 0.0020
-6.750 -0.5130 0.01528 0.01096 -0.0497 0.9304 0.0020
-6.500 -0.4901 0.01375 0.00912 -0.0490 0.9146 0.0020
-6.250 -0.4667 0.01254 0.00764 -0.0482 0.9005 0.0021
-6.000 -0.4425 0.01163 0.00652 -0.0475 0.8877 0.0023
-5.750 -0.4176 0.01096 0.00568 -0.0470 0.8760 0.0025
-5.500 -0.3918 0.01056 0.00517 -0.0466 0.8648 0.0026
-5.250 -0.3668 0.00980 0.00425 -0.0461 0.8543 0.0031
-5.000 -0.3405 0.00948 0.00384 -0.0459 0.8447 0.0037
-4.750 -0.3141 0.00919 0.00348 -0.0456 0.8352 0.0042
-4.500 -0.2874 0.00890 0.00310 -0.0454 0.8258 0.0049
-4.250 -0.2604 0.00870 0.00284 -0.0452 0.8167 0.0055
-4.000 -0.2337 0.00840 0.00246 -0.0450 0.8077 0.0072
-3.750 -0.2067 0.00817 0.00218 -0.0448 0.7983 0.0088
-3.500 -0.1795 0.00799 0.00194 -0.0447 0.7892 0.0100
-3.250 -0.1523 0.00782 0.00169 -0.0445 0.7781 0.0141
-3.000 -0.1255 0.00762 0.00150 -0.0443 0.7639 0.0287
-2.750 -0.0987 0.00753 0.00135 -0.0441 0.7446 0.0379
-2.500 -0.0718 0.00743 0.00122 -0.0440 0.7244 0.0516
-2.250 -0.0448 0.00733 0.00110 -0.0438 0.7074 0.0691
-2.000 -0.0180 0.00707 0.00096 -0.0437 0.6950 0.1209
-1.750 0.0083 0.00667 0.00084 -0.0437 0.6851 0.2164
-1.500 0.0351 0.00644 0.00074 -0.0436 0.6719 0.2781
-1.250 0.0614 0.00618 0.00066 -0.0435 0.6532 0.3594
-1.000 0.0877 0.00595 0.00060 -0.0433 0.6341 0.4420
-0.750 0.1139 0.00571 0.00057 -0.0432 0.6189 0.5312
-0.500 0.1404 0.00559 0.00058 -0.0429 0.5999 0.5945
-0.250 0.1669 0.00557 0.00060 -0.0427 0.5769 0.6404
0.000 0.1938 0.00561 0.00062 -0.0425 0.5520 0.6653
0.250 0.2206 0.00571 0.00065 -0.0423 0.5203 0.6843
0.500 0.2472 0.00581 0.00069 -0.0421 0.4887 0.7041
0.750 0.2738 0.00593 0.00075 -0.0419 0.4579 0.7222
1.250 0.3274 0.00622 0.00088 -0.0416 0.3987 0.7444
1.500 0.3542 0.00639 0.00095 -0.0414 0.3677 0.7550
2.000 0.4072 0.00674 0.00113 -0.0410 0.3118 0.7778
2.250 0.4341 0.00684 0.00126 -0.0409 0.3042 0.7900
2.500 0.4607 0.00697 0.00140 -0.0407 0.2966 0.8031
2.750 0.4882 0.00697 0.00147 -0.0405 0.2823 0.8170
3.000 0.5142 0.00715 0.00155 -0.0403 0.2440 0.8323
3.250 0.5399 0.00734 0.00169 -0.0399 0.2161 0.8491
3.500 0.5648 0.00754 0.00186 -0.0394 0.1846 0.8702
3.750 0.5890 0.00766 0.00202 -0.0387 0.1549 0.9146
4.000 0.6284 0.00811 0.00230 -0.0416 0.1005 1.0000
4.250 0.6528 0.00861 0.00259 -0.0412 0.0581 1.0000
4.500 0.6787 0.00891 0.00282 -0.0409 0.0430 1.0000
4.750 0.7044 0.00921 0.00306 -0.0406 0.0303 1.0000
5.000 0.7302 0.00951 0.00334 -0.0404 0.0212 1.0000
5.250 0.7556 0.00987 0.00364 -0.0400 0.0110 1.0000
5.500 0.7805 0.01032 0.00404 -0.0396 0.0036 1.0000
5.750 0.8063 0.01062 0.00438 -0.0393 0.0031 1.0000
6.000 0.8317 0.01097 0.00476 -0.0389 0.0027 1.0000
6.250 0.8569 0.01136 0.00521 -0.0385 0.0023 1.0000
6.500 0.8806 0.01203 0.00601 -0.0378 0.0017 1.0000
6.750 0.9054 0.01246 0.00652 -0.0373 0.0017 1.0000
7.000 0.9295 0.01299 0.00712 -0.0368 0.0016 1.0000
7.250 0.9531 0.01361 0.00783 -0.0361 0.0015 1.0000
7.500 0.9759 0.01435 0.00867 -0.0354 0.0015 1.0000
7.750 0.9979 0.01520 0.00964 -0.0345 0.0014 1.0000
8.000 1.0193 0.01612 0.01070 -0.0335 0.0014 1.0000
8.250 1.0397 0.01720 0.01193 -0.0324 0.0013 1.0000
8.500 1.0593 0.01839 0.01328 -0.0313 0.0013 1.0000
8.750 1.0778 0.01977 0.01484 -0.0300 0.0013 1.0000
9.000 1.0954 0.02126 0.01652 -0.0286 0.0013 1.0000
9.250 1.1114 0.02297 0.01849 -0.0271 0.0013 1.0000
9.500 1.1248 0.02500 0.02077 -0.0253 0.0013 1.0000
9.750 1.1356 0.02726 0.02332 -0.0234 0.0013 1.0000
10.000 1.1432 0.02975 0.02608 -0.0211 0.0013 1.0000
10.250 1.1479 0.03224 0.02885 -0.0188 0.0013 1.0000
10.500 1.1433 0.03548 0.03240 -0.0156 0.0013 1.0000
10.750 1.1332 0.03806 0.03520 -0.0117 0.0013 1.0000
11.000 1.1199 0.04080 0.03813 -0.0084 0.0014 1.0000
11.250 1.1040 0.04415 0.04168 -0.0063 0.0014 1.0000
11.500 1.0838 0.04857 0.04630 -0.0058 0.0014 1.0000
11.750 1.0644 0.05370 0.05161 -0.0071 0.0014 1.0000
12.000 1.0432 0.06023 0.05831 -0.0106 0.0014 1.0000
12.250 1.0246 0.06801 0.06624 -0.0161 0.0014 1.0000
12.500 1.0047 0.07860 0.07697 -0.0242 0.0014 1.0000
12.750 0.9888 0.08935 0.08782 -0.0312 0.0014 1.0000
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Polar data table (+)
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