HQ 3.5/9 AIRFOIL (hq359-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.5/9 AIRFOIL (hq359-il) Reynolds number: 100,000 Max Cl/Cd: 63.24 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq359-il-100000.txt Download as CSV file: xf-hq359-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.5/9 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3769 0.09940 0.09497 -0.0369 1.0000 0.0667
-7.750 -0.3946 0.09834 0.09403 -0.0356 1.0000 0.0668
-7.500 -0.4073 0.09645 0.09222 -0.0368 1.0000 0.0669
-7.250 -0.4176 0.09410 0.08992 -0.0393 1.0000 0.0671
-7.000 -0.4011 0.08885 0.08473 -0.0281 1.0000 0.0698
-6.750 -0.4051 0.08667 0.08260 -0.0260 1.0000 0.0713
-6.500 -0.4112 0.08442 0.08042 -0.0251 1.0000 0.0728
-6.250 -0.4166 0.08188 0.07792 -0.0253 1.0000 0.0748
-6.000 -0.4195 0.07892 0.07500 -0.0273 1.0000 0.0767
-5.750 -0.4158 0.07505 0.07092 -0.0385 1.0000 0.0804
-5.500 -0.4118 0.07047 0.06632 -0.0399 1.0000 0.0817
-5.250 -0.4088 0.06752 0.06351 -0.0360 1.0000 0.0838
-5.000 -0.4013 0.06472 0.06071 -0.0356 1.0000 0.0875
-4.750 -0.3793 0.05999 0.05557 -0.0439 1.0000 0.0958
-4.500 -0.3718 0.05699 0.05272 -0.0417 1.0000 0.0990
-4.250 -0.3490 0.05320 0.04862 -0.0462 1.0000 0.1104
-4.000 -0.3304 0.05020 0.04548 -0.0478 1.0000 0.1244
-3.750 -0.3141 0.04738 0.04263 -0.0481 1.0000 0.1400
-3.500 -0.2931 0.04478 0.03989 -0.0498 1.0000 0.1666
-2.750 -0.1325 0.03110 0.02356 -0.0631 0.9898 0.0660
-2.500 -0.0916 0.02801 0.02008 -0.0656 0.9860 0.0601
-2.250 -0.0475 0.02597 0.01747 -0.0679 0.9825 0.0579
-2.000 -0.0128 0.02482 0.01611 -0.0691 0.9766 0.0626
-1.750 0.0265 0.02351 0.01463 -0.0710 0.9719 0.0718
-1.500 0.0648 0.02232 0.01348 -0.0728 0.9672 0.0959
-1.250 0.0992 0.02109 0.01266 -0.0746 0.9610 0.1709
-1.000 0.1190 0.01906 0.01304 -0.0711 0.9568 0.8207
-0.750 0.1532 0.01875 0.01269 -0.0720 0.9482 1.0000
-0.500 0.1995 0.01919 0.01276 -0.0761 0.9421 1.0000
-0.250 0.2318 0.01947 0.01279 -0.0776 0.9326 1.0000
0.000 0.2707 0.01981 0.01291 -0.0802 0.9249 1.0000
0.250 0.3096 0.02011 0.01300 -0.0827 0.9169 1.0000
0.500 0.3424 0.02041 0.01316 -0.0842 0.9077 1.0000
0.750 0.3897 0.02060 0.01322 -0.0879 0.9013 1.0000
1.000 0.4242 0.02070 0.01322 -0.0892 0.8897 1.0000
1.250 0.4652 0.02061 0.01305 -0.0913 0.8777 1.0000
1.500 0.5091 0.02033 0.01272 -0.0937 0.8659 1.0000
1.750 0.5574 0.01995 0.01234 -0.0968 0.8571 1.0000
2.000 0.5987 0.01969 0.01207 -0.0988 0.8484 1.0000
2.250 0.6314 0.01959 0.01198 -0.0992 0.8374 1.0000
2.500 0.6677 0.01935 0.01176 -0.1002 0.8277 1.0000
2.750 0.7097 0.01881 0.01131 -0.1017 0.8198 1.0000
3.000 0.7409 0.01857 0.01111 -0.1016 0.8075 1.0000
3.250 0.7728 0.01823 0.01083 -0.1015 0.7949 1.0000
3.500 0.8050 0.01784 0.01050 -0.1012 0.7816 1.0000
3.750 0.8372 0.01739 0.01018 -0.1009 0.7672 1.0000
4.000 0.8697 0.01690 0.00975 -0.1005 0.7515 1.0000
4.250 0.8961 0.01664 0.00957 -0.0993 0.7306 1.0000
4.500 0.9262 0.01623 0.00919 -0.0985 0.7085 1.0000
4.750 0.9525 0.01599 0.00905 -0.0971 0.6799 1.0000
5.000 0.9777 0.01585 0.00890 -0.0956 0.6442 1.0000
5.250 1.0018 0.01587 0.00882 -0.0940 0.6003 1.0000
5.500 1.0232 0.01618 0.00895 -0.0921 0.5480 1.0000
5.750 1.0437 0.01677 0.00928 -0.0903 0.4974 1.0000
6.000 1.0642 0.01754 0.00986 -0.0888 0.4554 1.0000
6.250 1.0852 0.01838 0.01050 -0.0875 0.4218 1.0000
6.500 1.1064 0.01922 0.01123 -0.0864 0.3927 1.0000
6.750 1.1275 0.02004 0.01200 -0.0853 0.3666 1.0000
7.000 1.1482 0.02085 0.01284 -0.0841 0.3421 1.0000
7.250 1.1685 0.02169 0.01371 -0.0829 0.3187 1.0000
7.500 1.1878 0.02257 0.01465 -0.0816 0.2950 1.0000
7.750 1.2029 0.02336 0.01546 -0.0796 0.2667 1.0000
8.000 1.2104 0.02384 0.01596 -0.0768 0.2303 1.0000
8.250 1.2188 0.02436 0.01651 -0.0741 0.1874 1.0000
8.500 1.2202 0.02621 0.01778 -0.0707 0.0980 1.0000
8.750 1.2119 0.02973 0.02079 -0.0657 0.0529 1.0000
9.000 1.2118 0.03207 0.02316 -0.0618 0.0445 1.0000
9.250 1.2161 0.03417 0.02532 -0.0587 0.0395 1.0000
9.500 1.2230 0.03696 0.02810 -0.0562 0.0362 1.0000
9.750 1.2376 0.03898 0.03035 -0.0544 0.0328 1.0000
10.000 1.2504 0.04116 0.03266 -0.0528 0.0299 1.0000
10.250 1.2691 0.04425 0.03585 -0.0520 0.0281 1.0000
10.500 1.2934 0.04853 0.04047 -0.0519 0.0274 1.0000
10.750 1.3061 0.05275 0.04505 -0.0506 0.0272 1.0000
11.000 1.3086 0.05624 0.04891 -0.0482 0.0273 1.0000
11.250 1.3047 0.05954 0.05255 -0.0455 0.0274 1.0000
11.500 1.2964 0.06273 0.05609 -0.0427 0.0277 1.0000
11.750 1.2838 0.06622 0.05994 -0.0403 0.0280 1.0000
12.000 1.2680 0.07023 0.06430 -0.0386 0.0284 1.0000
12.250 1.2490 0.07481 0.06923 -0.0377 0.0288 1.0000
12.500 1.2263 0.08006 0.07480 -0.0380 0.0292 1.0000
12.750 1.2008 0.08607 0.08110 -0.0395 0.0296 1.0000
13.000 1.1747 0.09292 0.08821 -0.0425 0.0300 1.0000
13.250 1.1483 0.10072 0.09624 -0.0470 0.0304 1.0000
13.500 1.1210 0.10985 0.10557 -0.0531 0.0308 1.0000
13.750 1.0923 0.12089 0.11670 -0.0611 0.0314 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.5/9 AIRFOIL (hq359-il)