HQ 3.5/12 AIRFOIL (hq3512-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.5/12 AIRFOIL (hq3512-il) Reynolds number: 100,000 Max Cl/Cd: 57.82 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq3512-il-100000-n5.txt Download as CSV file: xf-hq3512-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.5/12 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3561 0.09346 0.08880 -0.0460 1.0000 0.0266
-9.000 -0.3670 0.09098 0.08641 -0.0447 1.0000 0.0263
-8.750 -0.3810 0.08863 0.08416 -0.0430 1.0000 0.0261
-8.500 -0.3813 0.08427 0.07985 -0.0461 0.9961 0.0257
-8.250 -0.3757 0.07814 0.07373 -0.0528 0.9883 0.0252
-8.000 -0.3751 0.07041 0.06601 -0.0624 0.9778 0.0248
-7.750 -0.3722 0.06135 0.05679 -0.0744 0.9662 0.0242
-7.500 -0.3645 0.05343 0.04853 -0.0828 0.9580 0.0239
-7.250 -0.3581 0.04763 0.04235 -0.0863 0.9479 0.0237
-7.000 -0.3408 0.04197 0.03615 -0.0901 0.9418 0.0238
-6.750 -0.3243 0.03777 0.03144 -0.0915 0.9341 0.0242
-6.500 -0.2965 0.03388 0.02692 -0.0939 0.9302 0.0256
-6.250 -0.2751 0.03111 0.02349 -0.0939 0.9232 0.0274
-6.000 -0.2469 0.02883 0.02089 -0.0950 0.9187 0.0290
-5.750 -0.2146 0.02707 0.01889 -0.0965 0.9157 0.0308
-5.500 -0.1912 0.02565 0.01723 -0.0960 0.9091 0.0328
-5.250 -0.1613 0.02406 0.01533 -0.0963 0.9047 0.0364
-5.000 -0.1291 0.02289 0.01411 -0.0975 0.9015 0.0416
-4.750 -0.1050 0.02222 0.01330 -0.0970 0.8950 0.0500
-4.500 -0.0751 0.02174 0.01267 -0.0976 0.8902 0.0611
-4.250 -0.0433 0.02090 0.01186 -0.0987 0.8869 0.0721
-4.000 -0.0209 0.02028 0.01121 -0.0978 0.8799 0.0786
-3.750 0.0084 0.01976 0.01058 -0.0981 0.8752 0.0888
-3.500 0.0403 0.01907 0.00989 -0.0990 0.8719 0.0995
-3.250 0.0621 0.01869 0.00950 -0.0981 0.8644 0.1126
-3.000 0.0917 0.01808 0.00899 -0.0987 0.8600 0.1454
-2.750 0.1199 0.01702 0.00853 -0.0995 0.8559 0.2806
-2.500 0.1382 0.01633 0.00876 -0.0979 0.8485 0.5086
-2.250 0.1666 0.01626 0.00879 -0.0975 0.8443 0.5920
-2.000 0.1907 0.01636 0.00885 -0.0964 0.8381 0.6361
-1.750 0.2150 0.01646 0.00896 -0.0951 0.8322 0.6771
-1.500 0.2412 0.01652 0.00905 -0.0937 0.8282 0.7243
-1.250 0.2585 0.01669 0.00924 -0.0911 0.8201 0.7540
-1.000 0.2870 0.01660 0.00905 -0.0908 0.8153 0.7699
-0.750 0.3120 0.01658 0.00897 -0.0901 0.8086 0.7796
-0.500 0.3397 0.01650 0.00881 -0.0899 0.8025 0.7893
-0.250 0.3699 0.01639 0.00860 -0.0901 0.7973 0.7996
0.000 0.3939 0.01640 0.00857 -0.0893 0.7895 0.8098
0.250 0.4256 0.01621 0.00832 -0.0897 0.7847 0.8200
0.500 0.4485 0.01621 0.00830 -0.0887 0.7753 0.8314
0.750 0.4785 0.01600 0.00803 -0.0887 0.7682 0.8433
1.000 0.5044 0.01585 0.00787 -0.0880 0.7582 0.8564
1.250 0.5299 0.01571 0.00773 -0.0872 0.7481 0.8707
1.500 0.5597 0.01547 0.00748 -0.0871 0.7390 0.8871
1.750 0.5897 0.01530 0.00732 -0.0872 0.7281 0.9083
2.000 0.6253 0.01520 0.00726 -0.0887 0.7169 0.9435
2.250 0.6594 0.01516 0.00719 -0.0901 0.7067 1.0000
2.500 0.6898 0.01516 0.00712 -0.0906 0.6976 1.0000
2.750 0.7159 0.01527 0.00723 -0.0905 0.6858 1.0000
3.000 0.7428 0.01537 0.00730 -0.0904 0.6740 1.0000
3.250 0.7699 0.01545 0.00735 -0.0903 0.6617 1.0000
3.500 0.7968 0.01554 0.00742 -0.0901 0.6484 1.0000
3.750 0.8229 0.01565 0.00753 -0.0897 0.6339 1.0000
4.000 0.8484 0.01578 0.00764 -0.0893 0.6179 1.0000
4.250 0.8736 0.01591 0.00775 -0.0887 0.6004 1.0000
4.500 0.8988 0.01605 0.00786 -0.0881 0.5813 1.0000
4.750 0.9222 0.01625 0.00805 -0.0873 0.5587 1.0000
5.000 0.9458 0.01647 0.00820 -0.0864 0.5342 1.0000
5.250 0.9684 0.01676 0.00839 -0.0854 0.5070 1.0000
5.500 0.9899 0.01712 0.00866 -0.0843 0.4777 1.0000
5.750 1.0104 0.01757 0.00898 -0.0830 0.4488 1.0000
6.000 1.0299 0.01810 0.00937 -0.0817 0.4213 1.0000
6.250 1.0491 0.01867 0.00983 -0.0803 0.3957 1.0000
6.500 1.0680 0.01927 0.01036 -0.0790 0.3733 1.0000
6.750 1.0869 0.01987 0.01091 -0.0777 0.3538 1.0000
7.000 1.1057 0.02050 0.01148 -0.0764 0.3367 1.0000
7.250 1.1245 0.02113 0.01208 -0.0752 0.3219 1.0000
7.500 1.1433 0.02177 0.01273 -0.0739 0.3085 1.0000
7.750 1.1620 0.02242 0.01339 -0.0727 0.2965 1.0000
8.000 1.1802 0.02310 0.01408 -0.0714 0.2851 1.0000
8.250 1.1974 0.02381 0.01478 -0.0700 0.2745 1.0000
8.500 1.2139 0.02450 0.01552 -0.0685 0.2637 1.0000
8.750 1.2303 0.02520 0.01630 -0.0670 0.2532 1.0000
9.000 1.2457 0.02595 0.01714 -0.0654 0.2432 1.0000
9.250 1.2604 0.02673 0.01798 -0.0638 0.2335 1.0000
9.500 1.2751 0.02748 0.01888 -0.0621 0.2234 1.0000
9.750 1.2885 0.02831 0.01980 -0.0604 0.2139 1.0000
10.000 1.3002 0.02919 0.02076 -0.0586 0.2042 1.0000
10.250 1.3117 0.03008 0.02181 -0.0568 0.1931 1.0000
10.500 1.3216 0.03107 0.02295 -0.0549 0.1816 1.0000
10.750 1.3294 0.03217 0.02414 -0.0530 0.1691 1.0000
11.000 1.3358 0.03340 0.02544 -0.0511 0.1557 1.0000
11.250 1.3415 0.03478 0.02688 -0.0494 0.1424 1.0000
11.500 1.3457 0.03636 0.02848 -0.0477 0.1282 1.0000
11.750 1.3492 0.03811 0.03025 -0.0461 0.1149 1.0000
12.000 1.3510 0.04010 0.03224 -0.0446 0.1008 1.0000
12.250 1.3518 0.04229 0.03444 -0.0433 0.0873 1.0000
12.500 1.3517 0.04467 0.03684 -0.0422 0.0750 1.0000
12.750 1.3502 0.04730 0.03950 -0.0412 0.0635 1.0000
13.250 1.3420 0.05345 0.04579 -0.0397 0.0443 1.0000
13.500 1.3344 0.05712 0.04954 -0.0394 0.0378 1.0000
13.750 1.3254 0.06115 0.05368 -0.0394 0.0321 1.0000
14.000 1.3138 0.06575 0.05839 -0.0399 0.0288 1.0000
14.250 1.3026 0.07058 0.06340 -0.0407 0.0260 1.0000
14.500 1.2886 0.07612 0.06908 -0.0422 0.0241 1.0000
14.750 1.2731 0.08222 0.07534 -0.0442 0.0230 1.0000
15.000 1.2571 0.08877 0.08204 -0.0468 0.0223 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.5/12 AIRFOIL (hq3512-il)