HQ 1.5/11 AIRFOIL (hq1511-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.5/11 AIRFOIL (hq1511-il) Reynolds number: 200,000 Max Cl/Cd: 67.84 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1511-il-200000.txt Download as CSV file: xf-hq1511-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/11 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6361 0.05343 0.04915 -0.0503 1.0000 0.0359
-8.500 -0.6720 0.04181 0.03683 -0.0473 1.0000 0.0283
-8.250 -0.6653 0.03912 0.03395 -0.0453 1.0000 0.0276
-8.000 -0.6652 0.03551 0.03007 -0.0426 1.0000 0.0272
-7.750 -0.6631 0.03264 0.02691 -0.0395 1.0000 0.0269
-7.500 -0.6595 0.03007 0.02402 -0.0364 1.0000 0.0268
-7.250 -0.6529 0.02779 0.02140 -0.0335 1.0000 0.0268
-7.000 -0.6428 0.02581 0.01911 -0.0310 1.0000 0.0271
-6.750 -0.6300 0.02417 0.01717 -0.0287 1.0000 0.0276
-6.500 -0.6018 0.02172 0.01448 -0.0295 0.9973 0.0290
-6.250 -0.5657 0.02042 0.01313 -0.0317 0.9933 0.0317
-6.000 -0.5313 0.01885 0.01141 -0.0333 0.9885 0.0376
-5.750 -0.4935 0.01806 0.01062 -0.0359 0.9842 0.0497
-5.500 -0.4587 0.01745 0.00994 -0.0377 0.9788 0.0602
-5.250 -0.4215 0.01706 0.00946 -0.0400 0.9739 0.0692
-5.000 -0.3834 0.01614 0.00859 -0.0426 0.9705 0.0781
-4.750 -0.3518 0.01547 0.00791 -0.0437 0.9638 0.0877
-4.500 -0.3153 0.01476 0.00722 -0.0459 0.9594 0.1020
-4.250 -0.2760 0.01395 0.00649 -0.0487 0.9566 0.1264
-4.000 -0.2492 0.01282 0.00595 -0.0494 0.9489 0.2254
-3.750 -0.2194 0.01146 0.00559 -0.0508 0.9440 0.4474
-3.500 -0.1909 0.01112 0.00566 -0.0508 0.9378 0.5578
-3.250 -0.1592 0.01108 0.00567 -0.0513 0.9314 0.6137
-3.000 -0.1231 0.01106 0.00567 -0.0524 0.9271 0.6510
-2.750 -0.0974 0.01111 0.00569 -0.0516 0.9177 0.6773
-2.500 -0.0649 0.01112 0.00566 -0.0519 0.9126 0.7036
-2.250 -0.0404 0.01117 0.00573 -0.0506 0.9032 0.7235
-2.000 -0.0111 0.01117 0.00570 -0.0501 0.8974 0.7448
-1.750 0.0126 0.01122 0.00574 -0.0488 0.8880 0.7628
-1.500 0.0407 0.01117 0.00565 -0.0483 0.8821 0.7747
-1.250 0.0654 0.01116 0.00562 -0.0473 0.8732 0.7857
-1.000 0.0923 0.01111 0.00554 -0.0466 0.8668 0.7979
-0.750 0.1163 0.01111 0.00553 -0.0454 0.8577 0.8113
-0.500 0.1418 0.01107 0.00546 -0.0445 0.8501 0.8237
-0.250 0.1667 0.01103 0.00542 -0.0435 0.8418 0.8347
0.000 0.1917 0.01098 0.00537 -0.0425 0.8337 0.8450
0.250 0.2172 0.01090 0.00527 -0.0417 0.8253 0.8555
0.500 0.2416 0.01084 0.00521 -0.0406 0.8153 0.8672
0.750 0.2666 0.01074 0.00511 -0.0395 0.8070 0.8782
1.000 0.2910 0.01066 0.00506 -0.0384 0.7973 0.8895
1.250 0.3163 0.01059 0.00500 -0.0376 0.7881 0.9012
1.500 0.3431 0.01049 0.00489 -0.0369 0.7793 0.9131
1.750 0.3698 0.01042 0.00485 -0.0364 0.7684 0.9264
2.000 0.4010 0.01035 0.00481 -0.0368 0.7582 0.9387
2.250 0.4366 0.01028 0.00472 -0.0381 0.7482 0.9506
2.500 0.4755 0.01023 0.00471 -0.0402 0.7363 0.9627
2.750 0.5168 0.01019 0.00471 -0.0429 0.7233 0.9747
3.000 0.5595 0.01015 0.00468 -0.0460 0.7092 0.9860
3.250 0.6021 0.01009 0.00464 -0.0491 0.6927 0.9982
3.500 0.6198 0.01008 0.00461 -0.0475 0.6761 1.0000
3.750 0.6377 0.01009 0.00459 -0.0457 0.6532 1.0000
4.000 0.6595 0.01012 0.00454 -0.0445 0.6234 1.0000
4.250 0.6825 0.01022 0.00456 -0.0434 0.5901 1.0000
4.500 0.7055 0.01040 0.00462 -0.0424 0.5496 1.0000
4.750 0.7271 0.01074 0.00473 -0.0412 0.5002 1.0000
5.000 0.7477 0.01124 0.00497 -0.0399 0.4499 1.0000
5.250 0.7687 0.01180 0.00533 -0.0388 0.4037 1.0000
5.500 0.7902 0.01232 0.00568 -0.0378 0.3625 1.0000
5.750 0.8119 0.01286 0.00606 -0.0369 0.3297 1.0000
6.000 0.8332 0.01344 0.00650 -0.0360 0.3023 1.0000
6.250 0.8545 0.01404 0.00698 -0.0351 0.2768 1.0000
6.500 0.8765 0.01455 0.00744 -0.0344 0.2509 1.0000
6.750 0.8981 0.01507 0.00792 -0.0336 0.2234 1.0000
7.000 0.9189 0.01566 0.00841 -0.0326 0.1899 1.0000
7.250 0.9386 0.01637 0.00893 -0.0316 0.1426 1.0000
7.500 0.9545 0.01749 0.00981 -0.0301 0.1108 1.0000
7.750 0.9704 0.01863 0.01084 -0.0285 0.0877 1.0000
8.000 0.9856 0.01981 0.01194 -0.0268 0.0720 1.0000
8.250 1.0011 0.02097 0.01310 -0.0251 0.0619 1.0000
8.500 1.0152 0.02229 0.01438 -0.0234 0.0553 1.0000
8.750 1.0328 0.02325 0.01544 -0.0220 0.0503 1.0000
9.000 1.0470 0.02485 0.01698 -0.0203 0.0465 1.0000
9.250 1.0652 0.02620 0.01845 -0.0191 0.0440 1.0000
9.500 1.0833 0.02735 0.01972 -0.0178 0.0412 1.0000
9.750 1.1000 0.02855 0.02095 -0.0166 0.0386 1.0000
10.000 1.1213 0.03126 0.02369 -0.0163 0.0365 1.0000
10.250 1.1393 0.03317 0.02585 -0.0152 0.0354 1.0000
10.500 1.1539 0.03474 0.02766 -0.0136 0.0340 1.0000
10.750 1.1664 0.03640 0.02954 -0.0120 0.0325 1.0000
11.000 1.1766 0.03811 0.03142 -0.0102 0.0312 1.0000
11.250 1.1846 0.04020 0.03370 -0.0083 0.0304 1.0000
11.500 1.1905 0.04243 0.03610 -0.0063 0.0297 1.0000
11.750 1.1935 0.04511 0.03897 -0.0043 0.0292 1.0000
12.000 1.1918 0.04853 0.04260 -0.0024 0.0286 1.0000
12.250 1.1820 0.05245 0.04680 -0.0003 0.0283 1.0000
12.750 1.1463 0.06069 0.05558 0.0028 0.0279 1.0000
13.000 1.1288 0.06468 0.05980 0.0032 0.0279 1.0000
13.250 1.1108 0.06850 0.06384 0.0028 0.0277 1.0000
13.500 1.0898 0.07473 0.07027 0.0014 0.0280 1.0000
13.750 1.0684 0.07985 0.07560 -0.0011 0.0279 1.0000
14.250 0.8428 0.14562 0.14219 -0.0434 0.0415 1.0000
14.500 0.8518 0.14900 0.14560 -0.0432 0.0470 1.0000
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Polar data table (+)
Polar graphs
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