HQ 1.5/10 AIRFOIL (hq1510-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.5/10 AIRFOIL (hq1510-il) Reynolds number: 200,000 Max Cl/Cd: 68.43 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1510-il-200000.txt Download as CSV file: xf-hq1510-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5123 0.08911 0.08567 -0.0329 1.0000 0.0524
-9.250 -0.5272 0.08208 0.07868 -0.0407 1.0000 0.0526
-9.000 -0.5442 0.07759 0.07416 -0.0444 1.0000 0.0527
-8.750 -0.5606 0.07397 0.07045 -0.0458 1.0000 0.0528
-8.500 -0.5712 0.07037 0.06670 -0.0466 1.0000 0.0529
-8.250 -0.5857 0.06275 0.05902 -0.0477 1.0000 0.0538
-8.000 -0.5753 0.05866 0.05504 -0.0471 1.0000 0.0549
-7.750 -0.5708 0.05551 0.05186 -0.0464 1.0000 0.0557
-7.500 -0.5672 0.05245 0.04875 -0.0454 1.0000 0.0570
-7.250 -0.5648 0.04937 0.04554 -0.0441 1.0000 0.0584
-7.000 -0.5859 0.03540 0.03022 -0.0392 1.0000 0.0289
-6.750 -0.5826 0.03182 0.02630 -0.0364 1.0000 0.0282
-6.500 -0.5753 0.02881 0.02290 -0.0337 1.0000 0.0280
-6.250 -0.5644 0.02626 0.01996 -0.0313 1.0000 0.0281
-6.000 -0.5501 0.02421 0.01755 -0.0292 1.0000 0.0287
-5.750 -0.5340 0.02304 0.01606 -0.0273 1.0000 0.0302
-5.500 -0.5144 0.02075 0.01365 -0.0267 0.9994 0.0334
-5.250 -0.4778 0.01920 0.01189 -0.0285 0.9956 0.0364
-5.000 -0.4420 0.01771 0.01019 -0.0300 0.9917 0.0418
-4.750 -0.4055 0.01684 0.00928 -0.0321 0.9868 0.0527
-4.500 -0.3691 0.01553 0.00801 -0.0343 0.9829 0.0682
-4.250 -0.3353 0.01481 0.00733 -0.0359 0.9771 0.0845
-4.000 -0.2984 0.01418 0.00670 -0.0382 0.9721 0.1047
-3.750 -0.2602 0.01324 0.00597 -0.0408 0.9684 0.1413
-3.500 -0.2334 0.01174 0.00541 -0.0418 0.9610 0.3229
-3.250 -0.2003 0.01085 0.00539 -0.0432 0.9566 0.5464
-3.000 -0.1678 0.01075 0.00541 -0.0439 0.9502 0.6197
-2.750 -0.1318 0.01072 0.00543 -0.0452 0.9449 0.6669
-2.500 -0.0932 0.01072 0.00542 -0.0470 0.9413 0.7025
-2.250 -0.0644 0.01074 0.00541 -0.0469 0.9332 0.7263
-2.000 -0.0294 0.01072 0.00540 -0.0477 0.9283 0.7505
-1.750 -0.0021 0.01076 0.00544 -0.0470 0.9204 0.7750
-1.500 0.0292 0.01071 0.00540 -0.0471 0.9145 0.7952
-1.250 0.0562 0.01067 0.00533 -0.0465 0.9061 0.8096
-1.000 0.0857 0.01059 0.00523 -0.0463 0.8998 0.8256
-0.750 0.1094 0.01055 0.00521 -0.0450 0.8906 0.8426
-0.500 0.1368 0.01045 0.00509 -0.0442 0.8842 0.8576
-0.250 0.1606 0.01038 0.00502 -0.0430 0.8740 0.8709
0.000 0.1857 0.01027 0.00491 -0.0420 0.8639 0.8836
0.250 0.2116 0.01013 0.00475 -0.0409 0.8540 0.8963
0.500 0.2376 0.01001 0.00462 -0.0400 0.8434 0.9098
0.750 0.2653 0.00994 0.00454 -0.0395 0.8322 0.9240
1.000 0.2969 0.00986 0.00446 -0.0398 0.8212 0.9385
1.250 0.3333 0.00978 0.00437 -0.0413 0.8110 0.9520
1.500 0.3732 0.00972 0.00430 -0.0436 0.8008 0.9644
1.750 0.4148 0.00967 0.00426 -0.0464 0.7894 0.9764
2.000 0.4578 0.00961 0.00422 -0.0496 0.7773 0.9878
2.250 0.4984 0.00955 0.00415 -0.0524 0.7644 1.0000
2.500 0.5140 0.00954 0.00412 -0.0503 0.7510 1.0000
2.750 0.5342 0.00957 0.00412 -0.0489 0.7367 1.0000
3.000 0.5573 0.00961 0.00415 -0.0479 0.7212 1.0000
3.250 0.5816 0.00965 0.00416 -0.0471 0.7040 1.0000
3.500 0.6062 0.00970 0.00418 -0.0462 0.6842 1.0000
3.750 0.6307 0.00976 0.00420 -0.0453 0.6600 1.0000
4.000 0.6548 0.00985 0.00425 -0.0443 0.6298 1.0000
4.250 0.6781 0.00999 0.00430 -0.0432 0.5902 1.0000
4.500 0.7007 0.01024 0.00439 -0.0420 0.5399 1.0000
4.750 0.7219 0.01067 0.00455 -0.0407 0.4797 1.0000
5.000 0.7423 0.01124 0.00483 -0.0394 0.4176 1.0000
5.250 0.7628 0.01188 0.00522 -0.0382 0.3619 1.0000
5.500 0.7842 0.01248 0.00564 -0.0373 0.3189 1.0000
5.750 0.8058 0.01308 0.00610 -0.0364 0.2833 1.0000
6.000 0.8273 0.01372 0.00659 -0.0356 0.2490 1.0000
6.250 0.8487 0.01435 0.00710 -0.0347 0.2082 1.0000
6.500 0.8680 0.01523 0.00772 -0.0336 0.1474 1.0000
6.750 0.8852 0.01643 0.00862 -0.0322 0.1075 1.0000
7.000 0.9036 0.01749 0.00957 -0.0309 0.0899 1.0000
7.250 0.9229 0.01847 0.01056 -0.0296 0.0799 1.0000
7.500 0.9406 0.01972 0.01177 -0.0282 0.0731 1.0000
7.750 0.9616 0.02063 0.01278 -0.0272 0.0675 1.0000
8.000 0.9799 0.02215 0.01425 -0.0260 0.0615 1.0000
8.250 1.0016 0.02261 0.01489 -0.0252 0.0554 1.0000
8.500 1.0195 0.02386 0.01611 -0.0242 0.0489 1.0000
8.750 1.0397 0.02442 0.01687 -0.0232 0.0435 1.0000
9.000 1.0557 0.02628 0.01873 -0.0219 0.0377 1.0000
9.250 1.0723 0.02681 0.01944 -0.0205 0.0309 1.0000
9.500 1.0857 0.02861 0.02137 -0.0188 0.0252 1.0000
9.750 1.1004 0.02996 0.02284 -0.0172 0.0221 1.0000
10.000 1.1122 0.03232 0.02528 -0.0157 0.0202 1.0000
10.250 1.1227 0.03548 0.02880 -0.0139 0.0191 1.0000
10.500 1.1313 0.03790 0.03155 -0.0118 0.0184 1.0000
10.750 1.1339 0.04060 0.03457 -0.0092 0.0180 1.0000
11.000 1.1312 0.04343 0.03771 -0.0063 0.0177 1.0000
11.250 1.1232 0.04656 0.04116 -0.0035 0.0176 1.0000
11.500 1.1122 0.04993 0.04482 -0.0013 0.0176 1.0000
11.750 1.0974 0.05371 0.04887 0.0002 0.0176 1.0000
12.000 1.0805 0.05790 0.05333 0.0008 0.0176 1.0000
12.250 1.0613 0.06270 0.05838 0.0004 0.0176 1.0000
12.500 1.0422 0.06797 0.06386 -0.0012 0.0178 1.0000
12.750 1.0198 0.07431 0.07040 -0.0042 0.0179 1.0000
13.000 0.9981 0.08133 0.07760 -0.0084 0.0181 1.0000
13.250 0.9767 0.08922 0.08564 -0.0136 0.0184 1.0000
13.500 0.9498 0.09953 0.09610 -0.0208 0.0186 1.0000
13.750 0.9266 0.11002 0.10668 -0.0278 0.0191 1.0000
14.000 0.8992 0.12259 0.11926 -0.0353 0.0196 1.0000
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