HQ 2.5/12 AIRFOIL (hq2512-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/12 AIRFOIL (hq2512-il) Reynolds number: 100,000 Max Cl/Cd: 55.19 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2512-il-100000.txt Download as CSV file: xf-hq2512-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/12 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 -0.3814 0.10197 0.09708 -0.0381 1.0000 0.1151
-9.500 -0.4028 0.09879 0.09405 -0.0425 1.0000 0.1203
-9.250 -0.4379 0.09527 0.09071 -0.0481 1.0000 0.1212
-9.000 -0.3896 0.09172 0.08704 -0.0404 1.0000 0.1269
-8.750 -0.3948 0.08877 0.08417 -0.0409 1.0000 0.1313
-8.500 -0.4412 0.08580 0.08147 -0.0448 1.0000 0.1349
-8.250 -0.4807 0.08336 0.07919 -0.0448 1.0000 0.1351
-8.000 -0.4289 0.08017 0.07591 -0.0389 1.0000 0.1421
-7.750 -0.4552 0.07852 0.07441 -0.0359 1.0000 0.1440
-7.500 -0.4843 0.07648 0.07248 -0.0341 1.0000 0.1446
-7.250 -0.5234 0.07363 0.06969 -0.0346 1.0000 0.1478
-6.500 -0.5769 0.05016 0.04493 -0.0395 1.0000 0.0814
-6.250 -0.5681 0.04311 0.03686 -0.0387 1.0000 0.0648
-6.000 -0.5553 0.03918 0.03258 -0.0379 1.0000 0.0640
-5.750 -0.5390 0.03675 0.02948 -0.0370 1.0000 0.0655
-5.500 -0.5222 0.03342 0.02607 -0.0368 1.0000 0.0692
-5.250 -0.4893 0.03091 0.02317 -0.0384 0.9964 0.0727
-5.000 -0.4526 0.02875 0.02030 -0.0402 0.9919 0.0801
-4.750 -0.4184 0.02704 0.01851 -0.0420 0.9870 0.0891
-4.500 -0.3806 0.02564 0.01694 -0.0443 0.9827 0.1007
-4.250 -0.3489 0.02451 0.01572 -0.0454 0.9770 0.1121
-4.000 -0.3131 0.02355 0.01467 -0.0471 0.9720 0.1242
-3.750 -0.2787 0.02273 0.01392 -0.0486 0.9668 0.1388
-3.500 -0.2486 0.02198 0.01332 -0.0494 0.9606 0.1582
-3.250 -0.2118 0.02110 0.01276 -0.0516 0.9559 0.2063
-3.000 -0.1951 0.01912 0.01266 -0.0503 0.9494 0.5244
-2.750 -0.1679 0.01962 0.01338 -0.0490 0.9429 0.6537
-2.500 -0.1451 0.02009 0.01383 -0.0473 0.9353 0.7020
-2.250 -0.1173 0.02050 0.01419 -0.0464 0.9287 0.7396
-2.000 -0.0958 0.02081 0.01444 -0.0445 0.9212 0.7693
-1.750 -0.0706 0.02103 0.01463 -0.0431 0.9143 0.7998
-1.500 -0.0538 0.02119 0.01478 -0.0401 0.9066 0.8299
-1.250 -0.0314 0.02123 0.01481 -0.0379 0.8997 0.8620
-1.000 -0.0128 0.02127 0.01481 -0.0355 0.8917 0.8891
-0.750 0.0213 0.02129 0.01478 -0.0360 0.8855 0.9162
-0.500 0.0679 0.02146 0.01485 -0.0393 0.8799 0.9413
-0.250 0.1223 0.02162 0.01491 -0.0449 0.8735 0.9568
0.000 0.1886 0.02165 0.01482 -0.0524 0.8698 0.9666
0.250 0.2464 0.02176 0.01486 -0.0589 0.8634 0.9770
0.500 0.3092 0.02171 0.01474 -0.0660 0.8581 0.9857
0.750 0.3776 0.02139 0.01436 -0.0736 0.8548 0.9922
1.000 0.4157 0.02123 0.01418 -0.0763 0.8437 1.0000
1.250 0.4352 0.02095 0.01388 -0.0750 0.8315 1.0000
1.500 0.4864 0.02014 0.01303 -0.0783 0.8247 1.0000
1.750 0.4908 0.02017 0.01304 -0.0747 0.8116 1.0000
2.000 0.5050 0.02024 0.01308 -0.0727 0.7994 1.0000
2.250 0.5316 0.02017 0.01298 -0.0725 0.7890 1.0000
2.500 0.5712 0.01967 0.01247 -0.0738 0.7818 1.0000
2.750 0.5962 0.01966 0.01244 -0.0732 0.7696 1.0000
3.000 0.6246 0.01952 0.01233 -0.0729 0.7581 1.0000
3.250 0.6566 0.01920 0.01201 -0.0729 0.7474 1.0000
3.500 0.6926 0.01868 0.01148 -0.0732 0.7375 1.0000
3.750 0.7204 0.01847 0.01129 -0.0726 0.7238 1.0000
4.000 0.7490 0.01820 0.01104 -0.0720 0.7095 1.0000
4.250 0.7783 0.01789 0.01073 -0.0714 0.6945 1.0000
4.500 0.8046 0.01769 0.01058 -0.0704 0.6768 1.0000
4.750 0.8305 0.01750 0.01040 -0.0693 0.6569 1.0000
5.000 0.8593 0.01718 0.01006 -0.0686 0.6366 1.0000
5.250 0.8829 0.01708 0.00996 -0.0672 0.6106 1.0000
5.500 0.9070 0.01699 0.00986 -0.0658 0.5817 1.0000
5.750 0.9307 0.01701 0.00977 -0.0644 0.5495 1.0000
6.000 0.9514 0.01724 0.00990 -0.0628 0.5120 1.0000
6.250 0.9724 0.01765 0.01009 -0.0612 0.4754 1.0000
6.500 0.9918 0.01824 0.01052 -0.0596 0.4394 1.0000
6.750 1.0114 0.01893 0.01102 -0.0582 0.4077 1.0000
7.000 1.0308 0.01968 0.01163 -0.0568 0.3790 1.0000
7.250 1.0500 0.02047 0.01230 -0.0555 0.3531 1.0000
7.500 1.0698 0.02131 0.01299 -0.0543 0.3302 1.0000
7.750 1.0889 0.02217 0.01384 -0.0530 0.3081 1.0000
8.000 1.1094 0.02315 0.01466 -0.0520 0.2883 1.0000
8.250 1.1278 0.02410 0.01565 -0.0507 0.2688 1.0000
8.500 1.1471 0.02512 0.01666 -0.0496 0.2510 1.0000
8.750 1.1660 0.02614 0.01765 -0.0484 0.2342 1.0000
9.000 1.1843 0.02717 0.01869 -0.0472 0.2185 1.0000
9.250 1.2020 0.02822 0.01976 -0.0459 0.2038 1.0000
9.500 1.2185 0.02927 0.02086 -0.0444 0.1901 1.0000
9.750 1.2347 0.03041 0.02210 -0.0429 0.1774 1.0000
10.000 1.2504 0.03163 0.02344 -0.0414 0.1657 1.0000
10.250 1.2625 0.03273 0.02462 -0.0395 0.1539 1.0000
10.500 1.2687 0.03357 0.02550 -0.0369 0.1423 1.0000
10.750 1.2669 0.03418 0.02609 -0.0333 0.1313 1.0000
11.000 1.2624 0.03494 0.02694 -0.0296 0.1209 1.0000
11.250 1.2585 0.03607 0.02825 -0.0265 0.1105 1.0000
11.500 1.2567 0.03756 0.02986 -0.0241 0.0999 1.0000
11.750 1.2533 0.03941 0.03179 -0.0218 0.0887 1.0000
12.000 1.2460 0.04183 0.03427 -0.0197 0.0758 1.0000
12.250 1.2371 0.04483 0.03729 -0.0178 0.0634 1.0000
12.500 1.2289 0.04814 0.04060 -0.0164 0.0538 1.0000
12.750 1.2252 0.05106 0.04354 -0.0154 0.0480 1.0000
13.000 1.2237 0.05429 0.04691 -0.0145 0.0436 1.0000
13.250 1.2229 0.05736 0.05015 -0.0139 0.0403 1.0000
13.500 1.2226 0.06041 0.05322 -0.0136 0.0379 1.0000
13.750 1.2264 0.06401 0.05689 -0.0128 0.0361 1.0000
14.000 1.2230 0.06789 0.06107 -0.0127 0.0352 1.0000
14.250 1.2166 0.07220 0.06567 -0.0130 0.0345 1.0000
14.500 1.2074 0.07691 0.07065 -0.0137 0.0340 1.0000
14.750 1.1947 0.08223 0.07624 -0.0151 0.0337 1.0000
15.000 1.1794 0.08809 0.08237 -0.0172 0.0336 1.0000
15.250 1.1604 0.09485 0.08939 -0.0202 0.0336 1.0000
15.500 1.1400 0.10228 0.09705 -0.0240 0.0338 1.0000
15.750 1.1167 0.11078 0.10578 -0.0290 0.0341 1.0000
16.000 1.0927 0.12009 0.11529 -0.0349 0.0347 1.0000
16.250 1.0666 0.13065 0.12601 -0.0417 0.0353 1.0000
16.500 1.0412 0.14180 0.13726 -0.0489 0.0360 1.0000
16.750 0.9396 0.18433 0.17970 -0.0734 0.0480 1.0000
17.000 0.9412 0.18996 0.18531 -0.0761 0.0475 1.0000
17.250 0.9431 0.19571 0.19103 -0.0789 0.0473 1.0000
17.500 0.9494 0.19990 0.19523 -0.0805 0.0467 1.0000
17.750 0.7393 0.20731 0.20338 -0.0773 0.0934 1.0000
18.000 0.7262 0.20974 0.20577 -0.0813 0.0926 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/12 AIRFOIL (hq2512-il)