HQ 0/10 AIRFOIL (hq010-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 0/10 AIRFOIL (hq010-il) Reynolds number: 500,000 Max Cl/Cd: 56.92 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq010-il-500000-n5.txt Download as CSV file: xf-hq010-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-15.500 -0.8276 0.13902 0.13665 0.0252 1.0000 0.0029
-15.250 -0.9098 0.11351 0.11102 0.0121 1.0000 0.0026
-15.000 -0.9716 0.09462 0.09200 0.0021 1.0000 0.0024
-14.750 -1.0258 0.07753 0.07469 -0.0081 1.0000 0.0023
-14.500 -1.0569 0.06632 0.06327 -0.0156 1.0000 0.0023
-14.250 -1.0778 0.05847 0.05524 -0.0208 1.0000 0.0022
-14.000 -1.0971 0.05191 0.04849 -0.0247 1.0000 0.0022
-13.750 -1.1120 0.04680 0.04320 -0.0270 1.0000 0.0022
-13.500 -1.1223 0.04282 0.03904 -0.0282 1.0000 0.0022
-13.250 -1.1312 0.03938 0.03542 -0.0285 1.0000 0.0023
-13.000 -1.1386 0.03641 0.03227 -0.0280 1.0000 0.0023
-12.750 -1.1428 0.03402 0.02972 -0.0269 1.0000 0.0023
-12.500 -1.1448 0.03203 0.02756 -0.0253 1.0000 0.0023
-12.250 -1.1453 0.03031 0.02569 -0.0231 1.0000 0.0023
-12.000 -1.1443 0.02880 0.02401 -0.0204 1.0000 0.0024
-11.750 -1.1390 0.02736 0.02240 -0.0181 1.0000 0.0025
-11.500 -1.1281 0.02609 0.02095 -0.0165 1.0000 0.0025
-11.250 -1.1151 0.02488 0.01959 -0.0150 1.0000 0.0025
-11.000 -1.1009 0.02367 0.01822 -0.0137 1.0000 0.0027
-10.750 -1.0846 0.02262 0.01703 -0.0125 1.0000 0.0027
-10.500 -1.0672 0.02162 0.01588 -0.0114 1.0000 0.0028
-10.250 -1.0486 0.02069 0.01482 -0.0104 1.0000 0.0030
-10.000 -1.0292 0.01981 0.01381 -0.0094 1.0000 0.0032
-9.750 -1.0097 0.01889 0.01279 -0.0085 1.0000 0.0036
-9.500 -0.9892 0.01809 0.01193 -0.0077 1.0000 0.0042
-9.250 -0.9680 0.01736 0.01113 -0.0069 1.0000 0.0056
-9.000 -0.9455 0.01679 0.01055 -0.0063 1.0000 0.0077
-8.750 -0.9225 0.01629 0.00998 -0.0058 1.0000 0.0090
-8.500 -0.8993 0.01580 0.00948 -0.0052 1.0000 0.0105
-8.250 -0.8753 0.01541 0.00906 -0.0048 1.0000 0.0120
-8.000 -0.8514 0.01500 0.00858 -0.0043 1.0000 0.0130
-7.750 -0.8269 0.01468 0.00822 -0.0039 1.0000 0.0138
-7.500 -0.8038 0.01414 0.00768 -0.0033 1.0000 0.0162
-7.250 -0.7794 0.01380 0.00732 -0.0029 1.0000 0.0182
-7.000 -0.7546 0.01354 0.00704 -0.0026 1.0000 0.0202
-6.750 -0.7301 0.01322 0.00668 -0.0021 1.0000 0.0211
-6.500 -0.7074 0.01266 0.00605 -0.0013 1.0000 0.0229
-6.250 -0.6829 0.01220 0.00558 -0.0008 0.9986 0.0252
-6.000 -0.6496 0.01180 0.00516 -0.0023 0.9821 0.0281
-5.750 -0.6144 0.01147 0.00475 -0.0041 0.9675 0.0309
-5.500 -0.5810 0.01107 0.00436 -0.0055 0.9510 0.0376
-5.250 -0.5504 0.01077 0.00404 -0.0062 0.9322 0.0453
-5.000 -0.5232 0.01048 0.00375 -0.0062 0.9122 0.0567
-4.750 -0.4979 0.01021 0.00348 -0.0057 0.8937 0.0709
-4.500 -0.4728 0.00994 0.00320 -0.0052 0.8784 0.0870
-4.250 -0.4478 0.00964 0.00294 -0.0047 0.8654 0.1110
-4.000 -0.4239 0.00916 0.00268 -0.0042 0.8533 0.1678
-3.750 -0.3982 0.00894 0.00250 -0.0038 0.8416 0.2016
-3.500 -0.3724 0.00870 0.00231 -0.0035 0.8298 0.2278
-3.000 -0.3211 0.00812 0.00193 -0.0028 0.8069 0.3034
-2.750 -0.2966 0.00768 0.00175 -0.0024 0.7949 0.3816
-2.500 -0.2715 0.00735 0.00160 -0.0020 0.7827 0.4447
-2.250 -0.2459 0.00709 0.00149 -0.0016 0.7723 0.4960
-2.000 -0.2200 0.00686 0.00141 -0.0012 0.7632 0.5496
-1.750 -0.1937 0.00671 0.00139 -0.0008 0.7549 0.5967
-1.500 -0.1664 0.00667 0.00135 -0.0007 0.7462 0.6175
-1.250 -0.1386 0.00664 0.00131 -0.0006 0.7374 0.6311
-1.000 -0.1109 0.00662 0.00128 -0.0004 0.7302 0.6445
-0.750 -0.0830 0.00660 0.00127 -0.0004 0.7225 0.6568
-0.500 -0.0553 0.00659 0.00125 -0.0003 0.7143 0.6689
-0.250 -0.0276 0.00658 0.00124 -0.0001 0.7044 0.6801
0.000 0.0000 0.00658 0.00124 0.0000 0.6920 0.6919
0.250 0.0276 0.00658 0.00124 0.0001 0.6801 0.7044
0.500 0.0553 0.00658 0.00125 0.0003 0.6690 0.7143
0.750 0.0831 0.00660 0.00127 0.0004 0.6568 0.7225
1.000 0.1110 0.00662 0.00128 0.0004 0.6445 0.7302
1.250 0.1387 0.00664 0.00131 0.0006 0.6311 0.7374
1.500 0.1664 0.00667 0.00135 0.0007 0.6175 0.7462
1.750 0.1938 0.00671 0.00139 0.0008 0.5964 0.7549
2.000 0.2200 0.00686 0.00141 0.0012 0.5496 0.7631
2.250 0.2459 0.00709 0.00148 0.0016 0.4956 0.7721
2.500 0.2715 0.00735 0.00160 0.0020 0.4450 0.7827
2.750 0.2966 0.00769 0.00175 0.0024 0.3813 0.7949
3.000 0.3212 0.00812 0.00193 0.0028 0.3036 0.8069
3.250 0.3466 0.00845 0.00212 0.0032 0.2589 0.8181
3.500 0.3725 0.00870 0.00231 0.0035 0.2281 0.8298
3.750 0.3982 0.00894 0.00250 0.0038 0.2017 0.8416
4.000 0.4239 0.00917 0.00268 0.0042 0.1672 0.8533
4.250 0.4478 0.00964 0.00294 0.0047 0.1112 0.8654
4.500 0.4728 0.00995 0.00320 0.0052 0.0867 0.8784
4.750 0.4979 0.01021 0.00348 0.0057 0.0708 0.8938
5.250 0.5504 0.01077 0.00404 0.0062 0.0451 0.9322
5.500 0.5810 0.01107 0.00436 0.0055 0.0376 0.9510
5.750 0.6145 0.01147 0.00475 0.0041 0.0308 0.9675
6.000 0.6496 0.01180 0.00516 0.0023 0.0279 0.9823
6.250 0.6829 0.01220 0.00558 0.0008 0.0251 0.9985
6.500 0.7073 0.01267 0.00607 0.0013 0.0227 1.0000
6.750 0.7301 0.01323 0.00669 0.0021 0.0211 1.0000
7.000 0.7548 0.01351 0.00700 0.0025 0.0200 1.0000
7.250 0.7795 0.01379 0.00731 0.0029 0.0181 1.0000
7.500 0.8038 0.01414 0.00767 0.0033 0.0160 1.0000
7.750 0.8270 0.01467 0.00820 0.0039 0.0138 1.0000
8.000 0.8513 0.01501 0.00860 0.0043 0.0131 1.0000
8.250 0.8754 0.01540 0.00905 0.0048 0.0120 1.0000
8.500 0.8993 0.01580 0.00948 0.0052 0.0105 1.0000
8.750 0.9225 0.01629 0.00997 0.0058 0.0089 1.0000
9.000 0.9455 0.01679 0.01055 0.0063 0.0076 1.0000
9.250 0.9678 0.01738 0.01113 0.0069 0.0053 1.0000
9.500 0.9893 0.01806 0.01191 0.0077 0.0044 1.0000
9.750 1.0097 0.01887 0.01277 0.0085 0.0037 1.0000
10.000 1.0291 0.01980 0.01380 0.0095 0.0032 1.0000
10.250 1.0485 0.02067 0.01480 0.0104 0.0030 1.0000
10.500 1.0671 0.02160 0.01586 0.0114 0.0029 1.0000
10.750 1.0846 0.02259 0.01699 0.0125 0.0027 1.0000
11.000 1.1007 0.02367 0.01822 0.0137 0.0026 1.0000
11.250 1.1153 0.02482 0.01952 0.0150 0.0026 1.0000
11.500 1.1276 0.02611 0.02097 0.0166 0.0025 1.0000
11.750 1.1383 0.02739 0.02244 0.0182 0.0024 1.0000
12.000 1.1434 0.02884 0.02405 0.0205 0.0024 1.0000
12.250 1.1453 0.03029 0.02565 0.0231 0.0023 1.0000
12.500 1.1450 0.03199 0.02751 0.0253 0.0023 1.0000
12.750 1.1431 0.03396 0.02966 0.0270 0.0023 1.0000
13.000 1.1369 0.03656 0.03243 0.0281 0.0022 1.0000
13.250 1.1331 0.03915 0.03518 0.0285 0.0023 1.0000
13.500 1.1232 0.04270 0.03892 0.0282 0.0022 1.0000
13.750 1.1122 0.04678 0.04317 0.0270 0.0022 1.0000
14.000 1.0991 0.05161 0.04818 0.0248 0.0023 1.0000
14.250 1.0778 0.05848 0.05525 0.0208 0.0022 1.0000
14.500 1.0568 0.06640 0.06336 0.0155 0.0023 1.0000
14.750 1.0248 0.07786 0.07502 0.0078 0.0023 1.0000
15.000 0.9782 0.09323 0.09060 -0.0014 0.0024 1.0000
15.250 0.9050 0.11478 0.11230 -0.0129 0.0026 1.0000
15.500 0.8147 0.14285 0.14049 -0.0271 0.0028 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 0/10 AIRFOIL (hq010-il)