HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 0/9 AIRFOIL (hq09-il) Reynolds number: 200,000 Max Cl/Cd: 46.27 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq09-il-200000.txt Download as CSV file: xf-hq09-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 0/9 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6586 0.10774 0.10429 0.0120 1.0000 0.0394
-10.000 -0.6605 0.10266 0.09924 0.0091 1.0000 0.0404
-7.750 -0.7538 0.04290 0.03796 -0.0144 1.0000 0.0223
-7.500 -0.7425 0.03919 0.03384 -0.0131 1.0000 0.0220
-7.250 -0.7336 0.03476 0.02913 -0.0124 1.0000 0.0237
-7.000 -0.7134 0.03411 0.02848 -0.0122 1.0000 0.0280
-6.750 -0.6979 0.02945 0.02322 -0.0105 1.0000 0.0271
-6.500 -0.6780 0.02592 0.01917 -0.0090 1.0000 0.0277
-6.250 -0.6553 0.02334 0.01614 -0.0077 1.0000 0.0301
-6.000 -0.6330 0.02108 0.01361 -0.0068 1.0000 0.0355
-5.750 -0.6089 0.01994 0.01234 -0.0062 1.0000 0.0406
-5.500 -0.5833 0.01915 0.01126 -0.0054 1.0000 0.0443
-5.250 -0.5611 0.01682 0.00893 -0.0044 1.0000 0.0478
-5.000 -0.5377 0.01578 0.00788 -0.0036 1.0000 0.0521
-4.750 -0.5146 0.01483 0.00688 -0.0027 1.0000 0.0589
-4.500 -0.4921 0.01393 0.00599 -0.0018 1.0000 0.0684
-4.250 -0.4709 0.01282 0.00498 -0.0007 1.0000 0.0884
-4.000 -0.4531 0.01123 0.00411 0.0005 1.0000 0.2097
-3.750 -0.4370 0.00987 0.00364 0.0019 1.0000 0.3911
-3.500 -0.4200 0.00907 0.00344 0.0039 1.0000 0.5351
-3.250 -0.4010 0.00877 0.00345 0.0059 1.0000 0.6288
-3.000 -0.3814 0.00864 0.00346 0.0078 1.0000 0.6847
-2.750 -0.3623 0.00859 0.00346 0.0097 1.0000 0.7234
-2.500 -0.3441 0.00857 0.00349 0.0118 1.0000 0.7530
-2.250 -0.3272 0.00859 0.00352 0.0141 1.0000 0.7813
-2.000 -0.3118 0.00864 0.00362 0.0167 1.0000 0.8099
-1.750 -0.2882 0.00872 0.00373 0.0176 0.9968 0.8370
-1.500 -0.2469 0.00876 0.00374 0.0149 0.9872 0.8548
-1.250 -0.2046 0.00879 0.00371 0.0119 0.9787 0.8696
-1.000 -0.1647 0.00881 0.00372 0.0096 0.9708 0.8858
-0.750 -0.1252 0.00885 0.00377 0.0075 0.9638 0.9012
-0.500 -0.0844 0.00890 0.00382 0.0052 0.9569 0.9161
-0.250 -0.0437 0.00893 0.00384 0.0029 0.9484 0.9293
0.000 -0.0001 0.00893 0.00384 0.0000 0.9402 0.9403
0.250 0.0436 0.00893 0.00384 -0.0029 0.9293 0.9484
0.500 0.0844 0.00890 0.00382 -0.0052 0.9162 0.9569
0.750 0.1252 0.00885 0.00377 -0.0075 0.9013 0.9639
1.000 0.1647 0.00881 0.00372 -0.0096 0.8859 0.9708
1.250 0.2044 0.00878 0.00371 -0.0119 0.8695 0.9788
1.500 0.2468 0.00876 0.00374 -0.0148 0.8548 0.9873
1.750 0.2882 0.00872 0.00373 -0.0176 0.8367 0.9969
2.000 0.3115 0.00864 0.00362 -0.0166 0.8099 1.0000
2.250 0.3268 0.00858 0.00352 -0.0140 0.7814 1.0000
2.500 0.3437 0.00857 0.00349 -0.0118 0.7532 1.0000
2.750 0.3619 0.00859 0.00346 -0.0097 0.7235 1.0000
3.000 0.3811 0.00864 0.00346 -0.0077 0.6850 1.0000
3.250 0.4006 0.00877 0.00345 -0.0058 0.6285 1.0000
3.500 0.4197 0.00907 0.00344 -0.0038 0.5357 1.0000
3.750 0.4367 0.00986 0.00363 -0.0019 0.3926 1.0000
4.000 0.4529 0.01122 0.00410 -0.0005 0.2114 1.0000
4.250 0.4707 0.01282 0.00498 0.0007 0.0884 1.0000
4.500 0.4920 0.01393 0.00599 0.0018 0.0685 1.0000
4.750 0.5146 0.01482 0.00687 0.0027 0.0590 1.0000
5.000 0.5376 0.01578 0.00788 0.0036 0.0522 1.0000
5.250 0.5611 0.01681 0.00893 0.0044 0.0478 1.0000
5.500 0.5834 0.01914 0.01126 0.0054 0.0443 1.0000
5.750 0.6090 0.01995 0.01235 0.0062 0.0406 1.0000
6.000 0.6332 0.02108 0.01360 0.0068 0.0354 1.0000
6.250 0.6554 0.02335 0.01615 0.0077 0.0302 1.0000
6.500 0.6782 0.02593 0.01918 0.0089 0.0277 1.0000
6.750 0.6981 0.02947 0.02325 0.0104 0.0271 1.0000
7.000 0.7137 0.03416 0.02853 0.0121 0.0282 1.0000
7.250 0.7276 0.03800 0.03284 0.0135 0.0270 1.0000
7.500 0.7281 0.04789 0.04353 0.0156 0.0411 1.0000
11.000 0.5646 0.11838 0.11516 -0.0138 0.0427 1.0000
11.250 0.5595 0.12342 0.12018 -0.0170 0.0407 1.0000
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