HQ 1.0/9 AIRFOIL (hq109-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.0/9 AIRFOIL (hq109-il) Reynolds number: 50,000 Max Cl/Cd: 33.15 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq109-il-50000-n5.txt Download as CSV file: xf-hq109-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/9 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.5703 0.10111 0.09417 -0.0155 1.0000 0.0369
-9.750 -0.5703 0.09625 0.08936 -0.0177 1.0000 0.0367
-9.500 -0.5728 0.09065 0.08382 -0.0211 1.0000 0.0366
-9.250 -0.5758 0.08512 0.07833 -0.0246 1.0000 0.0363
-9.000 -0.5824 0.07970 0.07295 -0.0283 1.0000 0.0361
-8.750 -0.5918 0.07496 0.06822 -0.0312 1.0000 0.0358
-8.500 -0.6019 0.07050 0.06373 -0.0331 1.0000 0.0356
-8.250 -0.6098 0.06587 0.05897 -0.0346 1.0000 0.0356
-8.000 -0.6153 0.06127 0.05415 -0.0356 1.0000 0.0359
-7.750 -0.6165 0.05684 0.04940 -0.0360 1.0000 0.0363
-7.500 -0.6143 0.05268 0.04475 -0.0360 1.0000 0.0371
-7.250 -0.6067 0.04896 0.04093 -0.0357 1.0000 0.0387
-7.000 -0.5951 0.04586 0.03761 -0.0351 1.0000 0.0404
-6.750 -0.5824 0.04230 0.03363 -0.0343 1.0000 0.0412
-6.500 -0.5671 0.03894 0.02982 -0.0333 1.0000 0.0425
-6.250 -0.5491 0.03585 0.02624 -0.0322 1.0000 0.0443
-6.000 -0.5289 0.03311 0.02299 -0.0309 1.0000 0.0473
-5.750 -0.5092 0.03100 0.02067 -0.0299 1.0000 0.0533
-5.500 -0.4875 0.02906 0.01845 -0.0286 1.0000 0.0597
-5.250 -0.4661 0.02715 0.01642 -0.0271 1.0000 0.0668
-5.000 -0.4459 0.02575 0.01487 -0.0258 1.0000 0.0806
-4.750 -0.4261 0.02432 0.01337 -0.0242 1.0000 0.0947
-4.500 -0.4071 0.02310 0.01213 -0.0227 1.0000 0.1170
-4.250 -0.3884 0.02183 0.01093 -0.0213 1.0000 0.1432
-4.000 -0.3701 0.02048 0.00989 -0.0204 1.0000 0.1969
-3.750 -0.3555 0.01884 0.00897 -0.0189 1.0000 0.3044
-3.500 -0.3450 0.01770 0.00889 -0.0156 1.0000 0.5052
-3.250 -0.3322 0.01750 0.00891 -0.0117 1.0000 0.6244
-3.000 -0.3198 0.01748 0.00896 -0.0076 1.0000 0.6996
-2.750 -0.3079 0.01749 0.00897 -0.0034 1.0000 0.7588
-2.500 -0.2948 0.01746 0.00888 0.0005 1.0000 0.8073
-2.250 -0.2773 0.01739 0.00870 0.0033 1.0000 0.8473
-2.000 -0.2505 0.01736 0.00854 0.0043 1.0000 0.8892
-1.750 -0.1983 0.01745 0.00834 0.0004 1.0000 0.9337
-1.500 -0.1438 0.01745 0.00806 -0.0052 1.0000 0.9632
-1.250 -0.0963 0.01736 0.00775 -0.0102 1.0000 0.9862
-1.000 -0.0800 0.01724 0.00747 -0.0101 1.0000 1.0000
-0.750 -0.0676 0.01723 0.00732 -0.0092 0.9957 1.0000
-0.500 -0.0293 0.01740 0.00733 -0.0126 0.9848 1.0000
-0.250 0.0083 0.01759 0.00737 -0.0157 0.9737 1.0000
0.000 0.0451 0.01778 0.00743 -0.0185 0.9625 1.0000
0.250 0.0823 0.01799 0.00755 -0.0212 0.9515 1.0000
0.500 0.1206 0.01820 0.00770 -0.0240 0.9410 1.0000
0.750 0.1614 0.01840 0.00789 -0.0271 0.9310 1.0000
1.000 0.1981 0.01859 0.00806 -0.0294 0.9192 1.0000
1.250 0.2349 0.01878 0.00827 -0.0315 0.9074 1.0000
1.500 0.2700 0.01897 0.00850 -0.0332 0.8955 1.0000
1.750 0.3042 0.01916 0.00878 -0.0347 0.8834 1.0000
2.000 0.3401 0.01924 0.00894 -0.0360 0.8676 1.0000
2.250 0.3784 0.01916 0.00895 -0.0371 0.8490 1.0000
2.500 0.4072 0.01915 0.00908 -0.0367 0.8272 1.0000
2.750 0.4374 0.01916 0.00919 -0.0364 0.8087 1.0000
3.000 0.4637 0.01924 0.00939 -0.0356 0.7892 1.0000
3.250 0.4900 0.01929 0.00957 -0.0346 0.7689 1.0000
3.500 0.5155 0.01933 0.00975 -0.0334 0.7465 1.0000
3.750 0.5410 0.01932 0.00994 -0.0321 0.7220 1.0000
4.000 0.5649 0.01933 0.01010 -0.0304 0.6928 1.0000
4.250 0.5885 0.01932 0.01021 -0.0286 0.6583 1.0000
4.500 0.6114 0.01933 0.01030 -0.0266 0.6155 1.0000
4.750 0.6331 0.01943 0.01040 -0.0244 0.5616 1.0000
5.000 0.6537 0.01972 0.01060 -0.0222 0.4941 1.0000
5.250 0.6724 0.02034 0.01091 -0.0200 0.4183 1.0000
5.500 0.6901 0.02128 0.01154 -0.0182 0.3461 1.0000
5.750 0.7071 0.02245 0.01239 -0.0166 0.2860 1.0000
6.000 0.7245 0.02374 0.01348 -0.0153 0.2326 1.0000
6.250 0.7420 0.02510 0.01467 -0.0142 0.1839 1.0000
6.500 0.7597 0.02658 0.01596 -0.0130 0.1467 1.0000
6.750 0.7781 0.02812 0.01753 -0.0119 0.1160 1.0000
7.000 0.7970 0.02984 0.01921 -0.0108 0.0951 1.0000
7.250 0.8166 0.03168 0.02110 -0.0097 0.0782 1.0000
7.500 0.8374 0.03383 0.02330 -0.0087 0.0669 1.0000
7.750 0.8562 0.03580 0.02544 -0.0077 0.0558 1.0000
8.000 0.8789 0.03879 0.02887 -0.0066 0.0499 1.0000
8.250 0.8946 0.04105 0.03123 -0.0057 0.0436 1.0000
8.500 0.9094 0.04429 0.03500 -0.0044 0.0395 1.0000
8.750 0.9207 0.04801 0.03924 -0.0030 0.0375 1.0000
9.000 0.9274 0.05162 0.04327 -0.0016 0.0357 1.0000
9.250 0.9310 0.05492 0.04688 -0.0003 0.0340 1.0000
9.500 0.9331 0.05787 0.04995 0.0007 0.0321 1.0000
9.750 0.9280 0.06184 0.05410 0.0017 0.0313 1.0000
10.000 0.9138 0.06567 0.05827 0.0031 0.0309 1.0000
10.250 0.8986 0.06967 0.06249 0.0037 0.0309 1.0000
10.500 0.8810 0.07430 0.06731 0.0027 0.0309 1.0000
10.750 0.8654 0.07943 0.07257 0.0006 0.0311 1.0000
11.000 0.8497 0.08522 0.07845 -0.0026 0.0313 1.0000
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Polar data table (+)
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