HQ 0/10 AIRFOIL (hq010-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 0/10 AIRFOIL (hq010-il) Reynolds number: 1,000,000 Max Cl/Cd: 70.74 at α=8° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq010-il-1000000-n5.txt Download as CSV file: xf-hq010-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 0/10 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
------ -------- --------- --------- -------- -------- --------
-16.000 -0.9739 0.11124 0.10919 0.0102 1.0000 0.0014
-15.750 -1.0348 0.09340 0.09120 0.0006 1.0000 0.0014
-15.500 -1.0998 0.07450 0.07204 -0.0107 1.0000 0.0014
-15.250 -1.1293 0.06405 0.06142 -0.0176 1.0000 0.0014
-15.000 -1.1476 0.05681 0.05403 -0.0223 1.0000 0.0013
-14.750 -1.1729 0.04948 0.04650 -0.0265 1.0000 0.0013
-14.500 -1.1860 0.04472 0.04158 -0.0286 1.0000 0.0013
-14.250 -1.1934 0.04112 0.03785 -0.0296 1.0000 0.0013
-14.000 -1.2042 0.03751 0.03408 -0.0299 1.0000 0.0013
-13.750 -1.2112 0.03463 0.03105 -0.0295 1.0000 0.0013
-13.500 -1.2139 0.03242 0.02871 -0.0285 1.0000 0.0013
-13.250 -1.2179 0.03032 0.02646 -0.0268 1.0000 0.0013
-13.000 -1.2202 0.02857 0.02456 -0.0245 1.0000 0.0014
-12.750 -1.2196 0.02718 0.02305 -0.0217 1.0000 0.0014
-12.500 -1.2168 0.02595 0.02169 -0.0189 1.0000 0.0014
-12.250 -1.2078 0.02473 0.02033 -0.0169 1.0000 0.0014
-12.000 -1.1961 0.02354 0.01901 -0.0153 1.0000 0.0014
-11.750 -1.1824 0.02242 0.01776 -0.0138 1.0000 0.0014
-11.500 -1.1662 0.02147 0.01670 -0.0126 1.0000 0.0014
-11.250 -1.1489 0.02055 0.01565 -0.0114 1.0000 0.0014
-11.000 -1.1309 0.01965 0.01464 -0.0104 1.0000 0.0015
-10.750 -1.1115 0.01885 0.01374 -0.0094 1.0000 0.0015
-10.500 -1.0914 0.01809 0.01288 -0.0086 1.0000 0.0015
-10.250 -1.0705 0.01738 0.01209 -0.0078 1.0000 0.0015
-10.000 -1.0491 0.01671 0.01133 -0.0070 1.0000 0.0016
-9.750 -1.0271 0.01609 0.01064 -0.0063 1.0000 0.0016
-9.500 -1.0045 0.01553 0.01001 -0.0057 1.0000 0.0016
-9.250 -0.9821 0.01491 0.00930 -0.0050 1.0000 0.0017
-9.000 -0.9593 0.01432 0.00861 -0.0044 1.0000 0.0020
-8.750 -0.9360 0.01381 0.00804 -0.0038 1.0000 0.0022
-8.500 -0.9127 0.01328 0.00747 -0.0032 1.0000 0.0027
-8.250 -0.8893 0.01275 0.00693 -0.0026 1.0000 0.0038
-8.000 -0.8659 0.01225 0.00645 -0.0020 1.0000 0.0069
-7.750 -0.8412 0.01192 0.00611 -0.0016 1.0000 0.0086
-7.500 -0.8165 0.01158 0.00577 -0.0012 1.0000 0.0099
-7.250 -0.7882 0.01124 0.00542 -0.0016 0.9899 0.0114
-7.000 -0.7525 0.01092 0.00509 -0.0036 0.9713 0.0129
-6.750 -0.7179 0.01065 0.00477 -0.0052 0.9515 0.0136
-6.500 -0.6903 0.01042 0.00447 -0.0052 0.9246 0.0149
-6.250 -0.6660 0.01019 0.00420 -0.0046 0.8982 0.0183
-6.000 -0.6410 0.01000 0.00395 -0.0040 0.8790 0.0206
-5.750 -0.6155 0.00982 0.00368 -0.0036 0.8638 0.0227
-5.500 -0.5898 0.00959 0.00343 -0.0032 0.8513 0.0273
-5.250 -0.5636 0.00941 0.00320 -0.0029 0.8404 0.0307
-4.750 -0.5109 0.00903 0.00278 -0.0024 0.8172 0.0433
-4.500 -0.4845 0.00883 0.00258 -0.0022 0.8056 0.0531
-4.250 -0.4580 0.00862 0.00238 -0.0020 0.7947 0.0662
-4.000 -0.4315 0.00841 0.00219 -0.0018 0.7838 0.0819
-3.750 -0.4051 0.00818 0.00200 -0.0016 0.7712 0.1025
-3.500 -0.3796 0.00783 0.00179 -0.0013 0.7582 0.1459
-3.250 -0.3532 0.00758 0.00164 -0.0012 0.7466 0.1876
-3.000 -0.3261 0.00741 0.00152 -0.0011 0.7375 0.2081
-2.750 -0.2988 0.00728 0.00140 -0.0010 0.7300 0.2250
-2.500 -0.2716 0.00708 0.00129 -0.0009 0.7227 0.2520
-2.250 -0.2450 0.00682 0.00116 -0.0008 0.7153 0.2951
-2.000 -0.2195 0.00638 0.00101 -0.0006 0.7074 0.3823
-1.750 -0.1925 0.00618 0.00093 -0.0005 0.7009 0.4265
-1.500 -0.1653 0.00598 0.00086 -0.0004 0.6944 0.4648
-1.250 -0.1392 0.00568 0.00079 -0.0002 0.6878 0.5349
-1.000 -0.1117 0.00558 0.00076 -0.0001 0.6783 0.5664
-0.750 -0.0838 0.00553 0.00074 -0.0001 0.6673 0.5852
-0.500 -0.0560 0.00551 0.00072 0.0000 0.6542 0.6024
-0.250 -0.0280 0.00550 0.00071 0.0000 0.6410 0.6163
0.000 0.0000 0.00550 0.00071 0.0000 0.6295 0.6296
0.250 0.0280 0.00550 0.00071 0.0000 0.6163 0.6412
0.500 0.0560 0.00551 0.00072 0.0000 0.6024 0.6541
0.750 0.0838 0.00553 0.00074 0.0001 0.5852 0.6672
1.000 0.1117 0.00558 0.00076 0.0001 0.5663 0.6783
1.250 0.1392 0.00568 0.00079 0.0002 0.5347 0.6878
1.500 0.1653 0.00598 0.00086 0.0004 0.4645 0.6944
1.750 0.1925 0.00618 0.00093 0.0005 0.4263 0.7010
2.000 0.2195 0.00638 0.00101 0.0006 0.3834 0.7075
2.250 0.2450 0.00682 0.00116 0.0008 0.2960 0.7152
2.500 0.2716 0.00708 0.00129 0.0009 0.2518 0.7227
2.750 0.2988 0.00728 0.00140 0.0010 0.2248 0.7300
3.000 0.3261 0.00741 0.00152 0.0011 0.2079 0.7375
3.250 0.3533 0.00758 0.00164 0.0011 0.1878 0.7465
3.500 0.3796 0.00783 0.00179 0.0013 0.1458 0.7583
3.750 0.4052 0.00818 0.00200 0.0016 0.1026 0.7711
4.000 0.4315 0.00841 0.00219 0.0018 0.0822 0.7839
4.250 0.4580 0.00862 0.00238 0.0020 0.0658 0.7947
4.500 0.4845 0.00882 0.00258 0.0022 0.0532 0.8056
4.750 0.5109 0.00903 0.00278 0.0024 0.0432 0.8172
5.250 0.5636 0.00941 0.00320 0.0029 0.0306 0.8403
5.500 0.5898 0.00959 0.00343 0.0032 0.0274 0.8514
5.750 0.6155 0.00982 0.00368 0.0036 0.0227 0.8639
6.000 0.6411 0.01000 0.00394 0.0040 0.0202 0.8790
6.500 0.6903 0.01043 0.00448 0.0053 0.0146 0.9248
6.750 0.7180 0.01065 0.00477 0.0052 0.0136 0.9517
7.000 0.7526 0.01092 0.00509 0.0035 0.0128 0.9714
7.250 0.7882 0.01124 0.00542 0.0016 0.0113 0.9899
7.500 0.8165 0.01159 0.00577 0.0012 0.0097 1.0000
7.750 0.8412 0.01192 0.00611 0.0016 0.0085 1.0000
8.000 0.8659 0.01224 0.00645 0.0020 0.0072 1.0000
8.250 0.8889 0.01281 0.00697 0.0026 0.0032 1.0000
8.500 0.9126 0.01328 0.00747 0.0032 0.0027 1.0000
8.750 0.9359 0.01381 0.00804 0.0038 0.0022 1.0000
9.000 0.9592 0.01432 0.00862 0.0044 0.0020 1.0000
9.250 0.9817 0.01495 0.00936 0.0051 0.0017 1.0000
9.500 1.0044 0.01553 0.01000 0.0057 0.0016 1.0000
9.750 1.0269 0.01610 0.01064 0.0064 0.0016 1.0000
10.000 1.0490 0.01671 0.01134 0.0070 0.0016 1.0000
10.250 1.0705 0.01737 0.01207 0.0078 0.0015 1.0000
10.500 1.0913 0.01807 0.01287 0.0086 0.0015 1.0000
10.750 1.1115 0.01883 0.01372 0.0094 0.0015 1.0000
11.000 1.1307 0.01965 0.01464 0.0104 0.0015 1.0000
11.250 1.1491 0.02051 0.01561 0.0114 0.0014 1.0000
11.500 1.1659 0.02149 0.01672 0.0126 0.0014 1.0000
11.750 1.1819 0.02246 0.01781 0.0139 0.0014 1.0000
12.000 1.1955 0.02359 0.01907 0.0153 0.0014 1.0000
12.250 1.2075 0.02473 0.02034 0.0170 0.0014 1.0000
12.500 1.2166 0.02595 0.02169 0.0189 0.0014 1.0000
12.750 1.2201 0.02714 0.02300 0.0217 0.0014 1.0000
13.000 1.2177 0.02874 0.02475 0.0247 0.0013 1.0000
13.250 1.2189 0.03024 0.02636 0.0267 0.0014 1.0000
13.500 1.2151 0.03231 0.02859 0.0284 0.0013 1.0000
13.750 1.2100 0.03475 0.03118 0.0295 0.0013 1.0000
14.000 1.2061 0.03732 0.03388 0.0299 0.0013 1.0000
14.250 1.1937 0.04110 0.03784 0.0295 0.0013 1.0000
14.500 1.1847 0.04491 0.04178 0.0285 0.0013 1.0000
14.750 1.1717 0.04969 0.04672 0.0264 0.0013 1.0000
15.000 1.1530 0.05597 0.05317 0.0228 0.0013 1.0000
15.250 1.1260 0.06478 0.06217 0.0170 0.0013 1.0000
15.500 1.0846 0.07805 0.07566 0.0082 0.0014 1.0000
15.750 1.0341 0.09369 0.09149 -0.0009 0.0014 1.0000
16.000 0.9657 0.11326 0.11122 -0.0114 0.0014 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 0/10 AIRFOIL (hq010-il)