HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.5/9 AIRFOIL (hq159-il) Reynolds number: 50,000 Max Cl/Cd: 35.14 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq159-il-50000.txt Download as CSV file: xf-hq159-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/9 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4886 0.11068 0.10385 -0.0046 1.0000 0.2491
-9.000 -0.5024 0.10934 0.10262 -0.0062 1.0000 0.2603
-8.750 -0.4966 0.10571 0.09905 -0.0058 1.0000 0.2741
-8.500 -0.4912 0.10229 0.09568 -0.0055 1.0000 0.2879
-8.250 -0.4804 0.09815 0.09152 -0.0047 1.0000 0.3030
-8.000 -0.4654 0.09402 0.08741 -0.0035 1.0000 0.3193
-7.750 -0.4587 0.09102 0.08445 -0.0024 1.0000 0.3402
-7.500 -0.4481 0.08767 0.08113 -0.0009 1.0000 0.3642
-6.250 -0.5171 0.05375 0.04665 -0.0396 1.0000 0.1344
-6.000 -0.5052 0.04850 0.04094 -0.0398 1.0000 0.1199
-5.750 -0.4915 0.04406 0.03612 -0.0393 1.0000 0.1129
-5.500 -0.4757 0.04051 0.03159 -0.0384 1.0000 0.1083
-5.250 -0.4591 0.03732 0.02841 -0.0373 1.0000 0.1134
-5.000 -0.4399 0.03440 0.02505 -0.0361 1.0000 0.1159
-4.750 -0.4185 0.03158 0.02172 -0.0347 1.0000 0.1177
-4.500 -0.3964 0.02908 0.01876 -0.0333 1.0000 0.1241
-4.250 -0.3753 0.02707 0.01667 -0.0320 1.0000 0.1397
-4.000 -0.3528 0.02512 0.01465 -0.0304 1.0000 0.1587
-3.750 -0.3308 0.02336 0.01292 -0.0287 1.0000 0.1889
-3.500 -0.3097 0.02154 0.01142 -0.0272 1.0000 0.2388
-3.250 -0.2938 0.01850 0.00994 -0.0250 1.0000 0.4122
-3.000 -0.3067 0.01805 0.01094 -0.0123 1.0000 0.7471
-2.750 -0.0695 0.01906 0.01025 -0.0363 1.0000 1.0000
-2.500 -0.0711 0.01870 0.00983 -0.0334 1.0000 1.0000
-2.250 -0.0774 0.01840 0.00948 -0.0297 1.0000 1.0000
-2.000 -0.0880 0.01811 0.00917 -0.0253 1.0000 1.0000
-1.750 -0.1010 0.01781 0.00884 -0.0205 1.0000 1.0000
-1.500 -0.1107 0.01753 0.00848 -0.0162 1.0000 1.0000
-1.250 -0.1082 0.01740 0.00817 -0.0136 1.0000 1.0000
-1.000 -0.0954 0.01741 0.00797 -0.0125 1.0000 1.0000
-0.750 -0.0785 0.01752 0.00783 -0.0120 1.0000 1.0000
-0.500 -0.0598 0.01770 0.00781 -0.0116 1.0000 1.0000
-0.250 -0.0403 0.01794 0.00787 -0.0114 1.0000 1.0000
0.000 -0.0205 0.01822 0.00800 -0.0112 1.0000 1.0000
0.250 -0.0006 0.01855 0.00818 -0.0110 1.0000 1.0000
0.500 0.0192 0.01891 0.00841 -0.0108 1.0000 1.0000
0.750 0.0390 0.01932 0.00872 -0.0107 1.0000 1.0000
1.000 0.0586 0.01977 0.00909 -0.0106 1.0000 1.0000
1.250 0.0780 0.02027 0.00952 -0.0105 1.0000 1.0000
1.500 0.0972 0.02081 0.01002 -0.0104 1.0000 1.0000
1.750 0.1160 0.02141 0.01059 -0.0104 1.0000 1.0000
2.000 0.1346 0.02206 0.01122 -0.0104 1.0000 1.0000
2.250 0.1528 0.02276 0.01192 -0.0105 1.0000 1.0000
2.500 0.2026 0.02413 0.01335 -0.0166 0.9845 1.0000
2.750 0.2690 0.02555 0.01493 -0.0252 0.9579 1.0000
3.000 0.3262 0.02658 0.01612 -0.0316 0.9324 1.0000
3.250 0.3805 0.02740 0.01714 -0.0369 0.9075 1.0000
3.500 0.4348 0.02801 0.01804 -0.0418 0.8817 1.0000
3.750 0.4901 0.02831 0.01864 -0.0461 0.8541 1.0000
4.000 0.5479 0.02820 0.01889 -0.0499 0.8241 1.0000
4.250 0.6058 0.02752 0.01866 -0.0524 0.7924 1.0000
4.500 0.6541 0.02652 0.01801 -0.0523 0.7571 1.0000
4.750 0.6962 0.02515 0.01700 -0.0502 0.7169 1.0000
5.000 0.7288 0.02385 0.01588 -0.0466 0.6685 1.0000
5.250 0.7560 0.02279 0.01482 -0.0424 0.6113 1.0000
5.500 0.7772 0.02236 0.01422 -0.0382 0.5439 1.0000
5.750 0.7953 0.02263 0.01409 -0.0344 0.4696 1.0000
6.000 0.8104 0.02371 0.01466 -0.0311 0.3895 1.0000
6.250 0.8249 0.02553 0.01598 -0.0282 0.3089 1.0000
6.500 0.8427 0.02780 0.01778 -0.0262 0.2448 1.0000
6.750 0.8665 0.03022 0.01999 -0.0251 0.2059 1.0000
7.000 0.8876 0.03220 0.02198 -0.0238 0.1771 1.0000
7.250 0.9126 0.03486 0.02484 -0.0228 0.1598 1.0000
7.500 0.9343 0.03722 0.02725 -0.0218 0.1428 1.0000
7.750 0.9539 0.04034 0.03078 -0.0204 0.1313 1.0000
8.000 0.9710 0.04437 0.03526 -0.0189 0.1246 1.0000
8.250 0.9861 0.04763 0.03875 -0.0176 0.1133 1.0000
8.500 0.9900 0.05211 0.04399 -0.0154 0.1098 1.0000
8.750 0.9916 0.05675 0.04916 -0.0135 0.1072 1.0000
9.000 1.0000 0.06059 0.05306 -0.0123 0.0997 1.0000
9.250 0.9918 0.06540 0.05833 -0.0107 0.0992 1.0000
9.500 0.9815 0.07042 0.06367 -0.0096 0.0995 1.0000
9.750 0.9727 0.07562 0.06905 -0.0089 0.1003 1.0000
10.000 0.9654 0.08095 0.07454 -0.0085 0.1016 1.0000
10.250 0.8763 0.09124 0.08501 -0.0154 0.1241 1.0000
10.500 0.8681 0.09793 0.09168 -0.0181 0.1265 1.0000
10.750 0.8695 0.10424 0.09798 -0.0194 0.1277 1.0000
11.000 0.7663 0.13411 0.12742 -0.0544 0.2928 1.0000
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