HQ 1.5/12 AIRFOIL (hq1512-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.5/12 AIRFOIL (hq1512-il) Reynolds number: 100,000 Max Cl/Cd: 49.93 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1512-il-100000.txt Download as CSV file: xf-hq1512-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/12 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4739 0.09689 0.09194 -0.0275 1.0000 0.1499
-9.500 -0.4788 0.06534 0.06083 -0.0481 1.0000 0.0867
-9.250 -0.5318 0.05497 0.05036 -0.0536 1.0000 0.0751
-9.000 -0.6118 0.06230 0.05738 -0.0515 1.0000 0.0766
-8.750 -0.6536 0.05501 0.04957 -0.0503 1.0000 0.0678
-8.500 -0.6596 0.05091 0.04529 -0.0483 1.0000 0.0660
-8.250 -0.6661 0.04688 0.04097 -0.0458 1.0000 0.0642
-8.000 -0.6709 0.04316 0.03688 -0.0428 1.0000 0.0631
-7.750 -0.6707 0.04040 0.03381 -0.0398 1.0000 0.0636
-7.500 -0.6680 0.03806 0.03115 -0.0368 1.0000 0.0651
-7.250 -0.6631 0.03563 0.02834 -0.0339 1.0000 0.0663
-7.000 -0.6548 0.03318 0.02549 -0.0313 1.0000 0.0670
-6.750 -0.6434 0.03104 0.02294 -0.0289 1.0000 0.0682
-6.500 -0.6299 0.02942 0.02087 -0.0267 1.0000 0.0703
-6.250 -0.6145 0.02743 0.01895 -0.0253 1.0000 0.0748
-6.000 -0.5977 0.02617 0.01749 -0.0236 1.0000 0.0799
-5.750 -0.5793 0.02464 0.01574 -0.0220 1.0000 0.0846
-5.500 -0.5619 0.02360 0.01475 -0.0207 1.0000 0.0932
-5.250 -0.5439 0.02244 0.01364 -0.0193 1.0000 0.1035
-5.000 -0.5262 0.02146 0.01270 -0.0180 1.0000 0.1178
-4.750 -0.5085 0.02060 0.01187 -0.0166 1.0000 0.1385
-4.500 -0.4914 0.01969 0.01114 -0.0154 1.0000 0.1674
-4.250 -0.4746 0.01875 0.01053 -0.0142 1.0000 0.2126
-4.000 -0.4432 0.01730 0.00992 -0.0161 0.9946 0.3247
-3.750 -0.4161 0.01655 0.01023 -0.0163 0.9881 0.5233
-3.500 -0.3826 0.01707 0.01090 -0.0168 0.9815 0.6181
-3.250 -0.3534 0.01756 0.01137 -0.0165 0.9737 0.6679
-3.000 -0.3212 0.01813 0.01190 -0.0166 0.9670 0.7059
-2.750 -0.2939 0.01853 0.01222 -0.0159 0.9592 0.7354
-2.500 -0.2629 0.01890 0.01252 -0.0159 0.9524 0.7622
-2.250 -0.2373 0.01915 0.01274 -0.0148 0.9445 0.7838
-2.000 -0.2077 0.01938 0.01289 -0.0147 0.9376 0.8073
-1.750 -0.1832 0.01951 0.01298 -0.0134 0.9296 0.8271
-1.500 -0.1538 0.01964 0.01305 -0.0133 0.9229 0.8467
-1.250 -0.1282 0.01968 0.01303 -0.0127 0.9147 0.8656
-1.000 -0.0947 0.01979 0.01309 -0.0132 0.9085 0.8836
-0.750 -0.0640 0.01993 0.01319 -0.0132 0.9007 0.9041
-0.500 -0.0092 0.02012 0.01331 -0.0178 0.8967 0.9214
-0.250 0.0344 0.02031 0.01345 -0.0211 0.8890 0.9343
0.000 0.0930 0.02036 0.01343 -0.0272 0.8841 0.9426
0.250 0.1633 0.02032 0.01333 -0.0353 0.8813 0.9477
0.500 0.2094 0.02038 0.01338 -0.0394 0.8721 0.9570
0.750 0.2839 0.01986 0.01283 -0.0475 0.8667 0.9598
1.000 0.3334 0.01950 0.01247 -0.0513 0.8547 0.9679
1.250 0.3893 0.01899 0.01197 -0.0561 0.8432 0.9739
1.500 0.4380 0.01855 0.01155 -0.0597 0.8332 0.9819
1.750 0.4883 0.01811 0.01115 -0.0639 0.8231 0.9893
2.000 0.5328 0.01782 0.01091 -0.0673 0.8109 0.9987
2.250 0.5512 0.01773 0.01085 -0.0658 0.7979 1.0000
2.500 0.5648 0.01765 0.01079 -0.0634 0.7850 1.0000
2.750 0.5785 0.01753 0.01069 -0.0609 0.7722 1.0000
3.000 0.5925 0.01738 0.01055 -0.0583 0.7595 1.0000
3.250 0.6076 0.01719 0.01037 -0.0557 0.7467 1.0000
3.500 0.6219 0.01706 0.01025 -0.0530 0.7325 1.0000
3.750 0.6395 0.01692 0.01010 -0.0507 0.7172 1.0000
4.000 0.6583 0.01684 0.01004 -0.0487 0.6995 1.0000
4.250 0.6805 0.01668 0.00988 -0.0470 0.6803 1.0000
4.500 0.7058 0.01641 0.00961 -0.0457 0.6602 1.0000
4.750 0.7285 0.01625 0.00946 -0.0440 0.6347 1.0000
5.000 0.7518 0.01607 0.00924 -0.0424 0.6051 1.0000
5.250 0.7748 0.01595 0.00903 -0.0406 0.5702 1.0000
5.500 0.7954 0.01603 0.00898 -0.0388 0.5276 1.0000
5.750 0.8153 0.01633 0.00905 -0.0369 0.4832 1.0000
6.000 0.8345 0.01682 0.00931 -0.0352 0.4409 1.0000
6.250 0.8537 0.01742 0.00972 -0.0337 0.4039 1.0000
6.500 0.8728 0.01806 0.01021 -0.0322 0.3707 1.0000
6.750 0.8918 0.01875 0.01077 -0.0308 0.3406 1.0000
7.000 0.9101 0.01948 0.01138 -0.0294 0.3116 1.0000
7.250 0.9274 0.02031 0.01209 -0.0279 0.2820 1.0000
7.500 0.9429 0.02126 0.01291 -0.0262 0.2500 1.0000
7.750 0.9564 0.02233 0.01381 -0.0242 0.2158 1.0000
8.000 0.9692 0.02347 0.01483 -0.0221 0.1820 1.0000
8.250 0.9832 0.02469 0.01585 -0.0202 0.1579 1.0000
8.500 0.9995 0.02592 0.01702 -0.0187 0.1400 1.0000
8.750 1.0188 0.02734 0.01839 -0.0176 0.1280 1.0000
9.000 1.0410 0.02893 0.01981 -0.0171 0.1183 1.0000
9.250 1.0606 0.03021 0.02127 -0.0160 0.1102 1.0000
9.500 1.0864 0.03216 0.02315 -0.0160 0.1040 1.0000
9.750 1.1076 0.03387 0.02513 -0.0152 0.0990 1.0000
10.000 1.1314 0.03569 0.02692 -0.0150 0.0941 1.0000
10.250 1.1506 0.03806 0.02955 -0.0142 0.0902 1.0000
10.500 1.1655 0.04024 0.03210 -0.0127 0.0872 1.0000
10.750 1.1807 0.04264 0.03477 -0.0114 0.0846 1.0000
11.000 1.1951 0.04504 0.03735 -0.0103 0.0819 1.0000
11.250 1.2072 0.04887 0.04131 -0.0095 0.0791 1.0000
11.500 1.2009 0.05129 0.04423 -0.0061 0.0777 1.0000
11.750 1.1914 0.05437 0.04773 -0.0029 0.0767 1.0000
12.000 1.1761 0.05743 0.05113 0.0007 0.0760 1.0000
12.250 1.1568 0.06072 0.05471 0.0037 0.0754 1.0000
12.500 1.1344 0.06445 0.05871 0.0057 0.0751 1.0000
12.750 1.1085 0.06887 0.06338 0.0066 0.0751 1.0000
13.000 1.0794 0.07439 0.06912 0.0062 0.0759 1.0000
13.250 1.0487 0.08074 0.07566 0.0043 0.0768 1.0000
13.500 1.0181 0.08808 0.08315 0.0012 0.0777 1.0000
13.750 0.9889 0.09640 0.09156 -0.0030 0.0786 1.0000
14.000 0.9662 0.10515 0.10036 -0.0072 0.0794 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.5/12 AIRFOIL (hq1512-il)