HQ 3.0/11 AIRFOIL (hq3011-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.0/11 AIRFOIL (hq3011-il) Reynolds number: 1,000,000 Max Cl/Cd: 101.18 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq3011-il-1000000-n5.txt Download as CSV file: xf-hq3011-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.0/11 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
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7075 0.04979 0.04805 -0.0705 1.0000 0.0032
-11.750 -0.7618 0.03900 0.03700 -0.0777 0.9999 0.0032
-11.500 -0.7551 0.03421 0.03197 -0.0837 0.9846 0.0034
-11.250 -0.7424 0.02937 0.02675 -0.0887 0.9748 0.0034
-11.000 -0.7224 0.02634 0.02342 -0.0916 0.9651 0.0035
-10.750 -0.6997 0.02441 0.02126 -0.0931 0.9548 0.0037
-10.500 -0.6799 0.02275 0.01936 -0.0936 0.9422 0.0038
-10.000 -0.6428 0.02002 0.01619 -0.0927 0.9188 0.0040
-9.750 -0.6223 0.01898 0.01497 -0.0922 0.9107 0.0041
-9.500 -0.6008 0.01807 0.01388 -0.0917 0.9033 0.0043
-9.250 -0.5786 0.01718 0.01284 -0.0913 0.8969 0.0044
-9.000 -0.5553 0.01648 0.01200 -0.0909 0.8904 0.0045
-8.750 -0.5316 0.01579 0.01118 -0.0906 0.8846 0.0045
-8.500 -0.5071 0.01522 0.01048 -0.0903 0.8787 0.0046
-8.000 -0.4611 0.01327 0.00821 -0.0895 0.8672 0.0051
-7.750 -0.4358 0.01274 0.00760 -0.0892 0.8616 0.0053
-7.500 -0.4101 0.01229 0.00706 -0.0890 0.8564 0.0055
-7.250 -0.3839 0.01190 0.00661 -0.0889 0.8508 0.0058
-7.000 -0.3577 0.01154 0.00618 -0.0887 0.8453 0.0061
-6.750 -0.3312 0.01116 0.00573 -0.0886 0.8400 0.0065
-6.500 -0.3047 0.01081 0.00530 -0.0885 0.8342 0.0067
-6.250 -0.2781 0.01047 0.00488 -0.0883 0.8288 0.0069
-6.000 -0.2511 0.01016 0.00451 -0.0882 0.8228 0.0070
-5.750 -0.2244 0.00981 0.00408 -0.0881 0.8169 0.0072
-5.500 -0.1974 0.00946 0.00367 -0.0880 0.8111 0.0076
-5.250 -0.1703 0.00918 0.00332 -0.0879 0.8048 0.0080
-5.000 -0.1429 0.00894 0.00303 -0.0878 0.7990 0.0086
-4.750 -0.1154 0.00873 0.00278 -0.0878 0.7926 0.0093
-4.500 -0.0879 0.00855 0.00254 -0.0877 0.7866 0.0101
-4.250 -0.0602 0.00834 0.00231 -0.0877 0.7802 0.0127
-4.000 -0.0330 0.00811 0.00212 -0.0876 0.7737 0.0228
-3.750 -0.0051 0.00798 0.00198 -0.0876 0.7669 0.0283
-3.250 0.0504 0.00776 0.00172 -0.0876 0.7507 0.0374
-3.000 0.0781 0.00769 0.00160 -0.0875 0.7423 0.0415
-2.750 0.1059 0.00758 0.00150 -0.0876 0.7343 0.0498
-2.500 0.1338 0.00746 0.00140 -0.0876 0.7277 0.0626
-2.250 0.1616 0.00732 0.00131 -0.0876 0.7206 0.0842
-2.000 0.1893 0.00717 0.00122 -0.0877 0.7140 0.1117
-1.750 0.2169 0.00701 0.00114 -0.0877 0.7053 0.1512
-1.500 0.2443 0.00681 0.00106 -0.0877 0.6953 0.2072
-1.250 0.2714 0.00661 0.00100 -0.0877 0.6839 0.2697
-1.000 0.2976 0.00620 0.00092 -0.0877 0.6728 0.3988
-0.750 0.3249 0.00603 0.00091 -0.0877 0.6626 0.4659
-0.500 0.3526 0.00598 0.00091 -0.0877 0.6533 0.5015
-0.250 0.3800 0.00593 0.00095 -0.0876 0.6440 0.5483
0.000 0.4076 0.00590 0.00099 -0.0876 0.6341 0.5863
0.250 0.4353 0.00592 0.00101 -0.0875 0.6229 0.6023
0.500 0.4628 0.00598 0.00103 -0.0875 0.6086 0.6144
1.000 0.5166 0.00618 0.00111 -0.0871 0.5670 0.6315
1.250 0.5436 0.00629 0.00116 -0.0869 0.5476 0.6386
1.500 0.5703 0.00643 0.00122 -0.0867 0.5256 0.6451
2.000 0.6230 0.00675 0.00140 -0.0863 0.4773 0.6598
2.250 0.6489 0.00696 0.00151 -0.0859 0.4497 0.6687
2.500 0.6747 0.00716 0.00163 -0.0856 0.4231 0.6780
2.750 0.7005 0.00736 0.00176 -0.0853 0.3983 0.6886
3.250 0.7514 0.00778 0.00205 -0.0846 0.3468 0.7154
3.500 0.7766 0.00798 0.00221 -0.0842 0.3219 0.7343
3.750 0.8013 0.00818 0.00240 -0.0837 0.2993 0.7612
4.000 0.8259 0.00833 0.00260 -0.0831 0.2819 0.8004
4.250 0.8505 0.00851 0.00279 -0.0826 0.2660 0.8331
4.500 0.8754 0.00871 0.00297 -0.0821 0.2503 0.8515
4.750 0.9003 0.00891 0.00315 -0.0816 0.2359 0.8647
5.000 0.9248 0.00914 0.00334 -0.0811 0.2194 0.8771
5.250 0.9479 0.00943 0.00357 -0.0803 0.1960 0.8910
5.500 0.9692 0.00975 0.00383 -0.0792 0.1696 0.9112
5.750 0.9973 0.01014 0.00414 -0.0796 0.1386 1.0000
6.000 1.0197 0.01061 0.00446 -0.0789 0.1127 1.0000
6.250 1.0425 0.01103 0.00477 -0.0782 0.0946 1.0000
6.500 1.0651 0.01147 0.00510 -0.0775 0.0763 1.0000
6.750 1.0867 0.01197 0.00550 -0.0766 0.0581 1.0000
7.000 1.1085 0.01244 0.00589 -0.0757 0.0434 1.0000
7.250 1.1288 0.01302 0.00637 -0.0747 0.0278 1.0000
7.500 1.1497 0.01354 0.00681 -0.0737 0.0176 1.0000
7.750 1.1682 0.01423 0.00742 -0.0723 0.0051 1.0000
8.000 1.1894 0.01467 0.00789 -0.0713 0.0036 1.0000
8.250 1.2105 0.01511 0.00836 -0.0703 0.0032 1.0000
8.500 1.2312 0.01554 0.00883 -0.0693 0.0029 1.0000
8.750 1.2511 0.01601 0.00935 -0.0681 0.0026 1.0000
9.000 1.2708 0.01645 0.00983 -0.0669 0.0025 1.0000
9.250 1.2888 0.01691 0.01034 -0.0654 0.0024 1.0000
9.500 1.3050 0.01741 0.01089 -0.0636 0.0023 1.0000
9.750 1.3205 0.01794 0.01148 -0.0618 0.0022 1.0000
10.000 1.3352 0.01852 0.01212 -0.0598 0.0021 1.0000
10.250 1.3493 0.01914 0.01280 -0.0579 0.0020 1.0000
10.500 1.3628 0.01982 0.01354 -0.0559 0.0019 1.0000
10.750 1.3758 0.02054 0.01431 -0.0540 0.0018 1.0000
11.000 1.3876 0.02136 0.01520 -0.0520 0.0017 1.0000
11.250 1.3987 0.02224 0.01615 -0.0500 0.0016 1.0000
11.500 1.4088 0.02321 0.01719 -0.0480 0.0015 1.0000
11.750 1.4176 0.02430 0.01836 -0.0460 0.0015 1.0000
12.000 1.4220 0.02575 0.01994 -0.0437 0.0014 1.0000
12.250 1.4263 0.02726 0.02155 -0.0415 0.0014 1.0000
12.500 1.4294 0.02891 0.02331 -0.0395 0.0014 1.0000
12.750 1.4345 0.03046 0.02494 -0.0378 0.0014 1.0000
13.000 1.4378 0.03223 0.02681 -0.0361 0.0014 1.0000
13.250 1.4438 0.03384 0.02850 -0.0349 0.0013 1.0000
13.500 1.4482 0.03564 0.03039 -0.0337 0.0013 1.0000
13.750 1.4518 0.03760 0.03244 -0.0327 0.0013 1.0000
14.000 1.4525 0.03992 0.03487 -0.0317 0.0013 1.0000
14.250 1.4522 0.04245 0.03750 -0.0309 0.0013 1.0000
14.500 1.4506 0.04524 0.04041 -0.0304 0.0013 1.0000
14.750 1.4487 0.04819 0.04347 -0.0302 0.0013 1.0000
15.000 1.4461 0.05140 0.04679 -0.0302 0.0013 1.0000
15.250 1.4385 0.05542 0.05096 -0.0306 0.0012 1.0000
15.500 1.4355 0.05905 0.05469 -0.0312 0.0012 1.0000
15.750 1.4265 0.06372 0.05950 -0.0324 0.0012 1.0000
16.000 1.4237 0.06771 0.06360 -0.0337 0.0012 1.0000
16.250 1.4064 0.07421 0.07025 -0.0361 0.0012 1.0000
16.500 1.4007 0.07911 0.07527 -0.0381 0.0012 1.0000
16.750 1.3835 0.08628 0.08259 -0.0413 0.0012 1.0000
17.000 1.3683 0.09337 0.08982 -0.0447 0.0012 1.0000
17.250 1.3558 0.10017 0.09675 -0.0482 0.0012 1.0000
17.500 1.3374 0.10836 0.10509 -0.0524 0.0012 1.0000
17.750 1.3178 0.11697 0.11384 -0.0570 0.0012 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.0/11 AIRFOIL (hq3011-il)