HQ 0/9 AIRFOIL (hq09-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 0/9 AIRFOIL (hq09-il) Reynolds number: 50,000 Max Cl/Cd: 27.26 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq09-il-50000.txt Download as CSV file: xf-hq09-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.6630 0.05546 0.04929 -0.0137 1.0000 0.1482
-7.250 -0.6693 0.04795 0.04141 -0.0153 1.0000 0.1345
-7.000 -0.7097 0.05257 0.04475 -0.0148 1.0000 0.1221
-6.750 -0.6941 0.04753 0.03944 -0.0144 1.0000 0.1180
-6.500 -0.6791 0.04274 0.03401 -0.0137 1.0000 0.1144
-6.250 -0.6601 0.03869 0.02935 -0.0128 1.0000 0.1140
-6.000 -0.6396 0.03579 0.02582 -0.0117 1.0000 0.1217
-5.750 -0.6159 0.03261 0.02233 -0.0109 1.0000 0.1277
-5.500 -0.5894 0.02984 0.01912 -0.0099 1.0000 0.1336
-5.250 -0.5621 0.02735 0.01649 -0.0091 1.0000 0.1438
-5.000 -0.5372 0.02551 0.01469 -0.0081 1.0000 0.1677
-4.750 -0.5107 0.02345 0.01273 -0.0069 1.0000 0.1989
-4.500 -0.4904 0.02085 0.01082 -0.0052 1.0000 0.2830
-4.250 -0.4925 0.01806 0.01053 0.0026 1.0000 0.5930
-4.000 -0.4856 0.01885 0.01152 0.0118 1.0000 0.7348
-3.750 -0.4656 0.01993 0.01235 0.0193 1.0000 0.8114
-3.500 -0.3490 0.02219 0.01357 0.0128 1.0000 0.9117
-3.250 -0.2594 0.02149 0.01211 0.0014 1.0000 0.9521
-3.000 -0.1776 0.02022 0.01034 -0.0101 1.0000 0.9849
-2.750 -0.1299 0.01916 0.00901 -0.0160 1.0000 1.0000
-2.500 -0.1137 0.01858 0.00834 -0.0158 1.0000 1.0000
-2.250 -0.0973 0.01808 0.00771 -0.0154 1.0000 1.0000
-2.000 -0.0809 0.01763 0.00720 -0.0149 1.0000 1.0000
-1.750 -0.0648 0.01723 0.00676 -0.0142 1.0000 1.0000
-1.500 -0.0491 0.01688 0.00639 -0.0134 1.0000 1.0000
-1.250 -0.0342 0.01658 0.00608 -0.0124 1.0000 1.0000
-1.000 -0.0206 0.01633 0.00584 -0.0110 1.0000 1.0000
-0.750 -0.0092 0.01612 0.00564 -0.0093 1.0000 1.0000
-0.500 -0.0014 0.01596 0.00552 -0.0070 1.0000 1.0000
-0.250 0.0013 0.01586 0.00546 -0.0038 1.0000 1.0000
0.000 0.0000 0.01582 0.00545 0.0000 1.0000 1.0000
0.250 -0.0013 0.01586 0.00546 0.0038 1.0000 1.0000
0.500 0.0014 0.01596 0.00552 0.0070 1.0000 1.0000
0.750 0.0092 0.01612 0.00564 0.0093 1.0000 1.0000
1.000 0.0207 0.01632 0.00583 0.0110 1.0000 1.0000
1.250 0.0343 0.01658 0.00608 0.0124 1.0000 1.0000
1.500 0.0492 0.01688 0.00638 0.0134 1.0000 1.0000
1.750 0.0649 0.01723 0.00675 0.0142 1.0000 1.0000
2.000 0.0811 0.01762 0.00719 0.0149 1.0000 1.0000
2.250 0.0975 0.01807 0.00770 0.0154 1.0000 1.0000
2.500 0.1139 0.01858 0.00833 0.0158 1.0000 1.0000
2.750 0.1302 0.01916 0.00901 0.0160 1.0000 1.0000
3.000 0.1779 0.02021 0.01033 0.0101 0.9849 1.0000
3.250 0.2597 0.02148 0.01210 -0.0014 0.9522 1.0000
3.500 0.3494 0.02218 0.01356 -0.0129 0.9117 1.0000
3.750 0.4656 0.01994 0.01235 -0.0193 0.8117 1.0000
4.000 0.4855 0.01885 0.01152 -0.0118 0.7348 1.0000
4.250 0.4924 0.01806 0.01053 -0.0026 0.5931 1.0000
4.500 0.4904 0.02085 0.01081 0.0052 0.2831 1.0000
4.750 0.5106 0.02344 0.01273 0.0069 0.1990 1.0000
5.000 0.5372 0.02551 0.01469 0.0081 0.1678 1.0000
5.250 0.5621 0.02735 0.01650 0.0091 0.1439 1.0000
5.500 0.5894 0.02985 0.01912 0.0099 0.1336 1.0000
5.750 0.6159 0.03261 0.02234 0.0109 0.1277 1.0000
6.000 0.6397 0.03580 0.02583 0.0117 0.1217 1.0000
6.250 0.6602 0.03870 0.02935 0.0128 0.1140 1.0000
6.500 0.6793 0.04275 0.03403 0.0137 0.1144 1.0000
6.750 0.6943 0.04755 0.03946 0.0143 0.1180 1.0000
7.000 0.7099 0.05259 0.04476 0.0148 0.1221 1.0000
7.250 0.7145 0.05880 0.05172 0.0140 0.1342 1.0000
7.500 0.6635 0.05538 0.04921 0.0137 0.1483 1.0000
7.750 0.6514 0.06461 0.05872 0.0099 0.1766 1.0000
8.000 0.6027 0.07264 0.06677 0.0024 0.1928 1.0000
8.250 0.6040 0.09581 0.08918 -0.0276 0.3627 1.0000
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