HQ 2.5/11 AIRFOIL (hq2511-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.5/11 AIRFOIL (hq2511-il) Reynolds number: 200,000 Max Cl/Cd: 78.14 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2511-il-200000.txt Download as CSV file: xf-hq2511-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/11 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4007 0.08920 0.08577 -0.0423 1.0000 0.0490
-9.000 -0.4010 0.08584 0.08246 -0.0431 1.0000 0.0498
-8.750 -0.4033 0.08249 0.07917 -0.0439 1.0000 0.0506
-8.500 -0.4100 0.07888 0.07562 -0.0449 1.0000 0.0514
-8.250 -0.4226 0.07537 0.07221 -0.0456 1.0000 0.0519
-8.000 -0.4482 0.07279 0.06975 -0.0442 1.0000 0.0518
-7.750 -0.4832 0.07047 0.06752 -0.0422 1.0000 0.0515
-7.500 -0.5076 0.06815 0.06522 -0.0402 1.0000 0.0515
-7.250 -0.4955 0.06169 0.05857 -0.0483 0.9946 0.0532
-6.750 -0.4696 0.03560 0.03036 -0.0626 0.9775 0.0277
-6.500 -0.4420 0.03081 0.02520 -0.0650 0.9733 0.0266
-6.250 -0.4090 0.02677 0.02060 -0.0674 0.9704 0.0260
-6.000 -0.3831 0.02410 0.01750 -0.0676 0.9642 0.0258
-5.750 -0.3494 0.02184 0.01483 -0.0689 0.9605 0.0265
-5.500 -0.3120 0.02020 0.01291 -0.0708 0.9579 0.0275
-5.250 -0.2819 0.01863 0.01135 -0.0717 0.9531 0.0304
-5.000 -0.2493 0.01778 0.01036 -0.0726 0.9480 0.0357
-4.750 -0.2127 0.01684 0.00950 -0.0748 0.9449 0.0536
-4.500 -0.1723 0.01643 0.00901 -0.0775 0.9428 0.0694
-4.250 -0.1463 0.01586 0.00844 -0.0775 0.9357 0.0777
-4.000 -0.1116 0.01533 0.00786 -0.0790 0.9316 0.0879
-3.750 -0.0749 0.01475 0.00728 -0.0810 0.9288 0.1017
-3.500 -0.0494 0.01424 0.00686 -0.0809 0.9214 0.1212
-3.250 -0.0188 0.01308 0.00628 -0.0821 0.9163 0.2326
-3.000 0.0033 0.01175 0.00624 -0.0814 0.9103 0.5531
-2.750 0.0301 0.01171 0.00631 -0.0807 0.9034 0.6243
-2.500 0.0626 0.01166 0.00623 -0.0809 0.8994 0.6639
-2.250 0.0865 0.01173 0.00627 -0.0797 0.8911 0.6919
-2.000 0.1161 0.01169 0.00621 -0.0794 0.8861 0.7162
-1.750 0.1410 0.01174 0.00624 -0.0783 0.8791 0.7373
-1.500 0.1670 0.01175 0.00625 -0.0773 0.8733 0.7579
-1.250 0.1939 0.01175 0.00623 -0.0764 0.8689 0.7809
-1.000 0.2153 0.01182 0.00633 -0.0747 0.8610 0.7991
-0.750 0.2423 0.01175 0.00625 -0.0740 0.8563 0.8142
-0.500 0.2660 0.01177 0.00625 -0.0729 0.8492 0.8278
-0.250 0.2926 0.01167 0.00613 -0.0723 0.8428 0.8405
0.000 0.3173 0.01160 0.00604 -0.0713 0.8350 0.8526
0.250 0.3434 0.01144 0.00587 -0.0705 0.8276 0.8629
0.500 0.3680 0.01136 0.00580 -0.0695 0.8191 0.8739
0.750 0.3946 0.01121 0.00563 -0.0688 0.8120 0.8857
1.000 0.4186 0.01113 0.00556 -0.0678 0.8026 0.8987
1.250 0.4459 0.01095 0.00536 -0.0671 0.7949 0.9124
1.500 0.4729 0.01081 0.00524 -0.0666 0.7843 0.9281
1.750 0.5054 0.01070 0.00515 -0.0673 0.7749 0.9450
2.000 0.5452 0.01056 0.00500 -0.0695 0.7669 0.9620
2.250 0.5871 0.01048 0.00495 -0.0725 0.7558 0.9833
2.500 0.6206 0.01045 0.00490 -0.0741 0.7450 1.0000
2.750 0.6464 0.01044 0.00485 -0.0738 0.7330 1.0000
3.000 0.6730 0.01041 0.00478 -0.0736 0.7183 1.0000
3.250 0.6994 0.01040 0.00472 -0.0732 0.7015 1.0000
3.500 0.7257 0.01041 0.00471 -0.0728 0.6838 1.0000
3.750 0.7521 0.01045 0.00470 -0.0723 0.6659 1.0000
4.000 0.7779 0.01052 0.00475 -0.0718 0.6457 1.0000
4.250 0.8033 0.01062 0.00481 -0.0712 0.6229 1.0000
4.500 0.8280 0.01075 0.00491 -0.0705 0.5962 1.0000
4.750 0.8521 0.01093 0.00502 -0.0697 0.5653 1.0000
5.000 0.8752 0.01120 0.00517 -0.0687 0.5295 1.0000
5.250 0.8968 0.01160 0.00539 -0.0675 0.4895 1.0000
5.500 0.9176 0.01211 0.00571 -0.0663 0.4502 1.0000
5.750 0.9382 0.01267 0.00611 -0.0651 0.4132 1.0000
6.000 0.9587 0.01321 0.00652 -0.0640 0.3740 1.0000
6.250 0.9783 0.01380 0.00698 -0.0627 0.3363 1.0000
6.500 0.9967 0.01453 0.00752 -0.0613 0.3033 1.0000
6.750 1.0148 0.01528 0.00811 -0.0599 0.2723 1.0000
7.000 1.0336 0.01595 0.00867 -0.0587 0.2427 1.0000
7.250 1.0530 0.01657 0.00923 -0.0576 0.2168 1.0000
7.500 1.0718 0.01723 0.00980 -0.0563 0.1957 1.0000
7.750 1.0915 0.01782 0.01038 -0.0553 0.1772 1.0000
8.000 1.1104 0.01843 0.01095 -0.0541 0.1592 1.0000
8.250 1.1283 0.01909 0.01156 -0.0528 0.1422 1.0000
8.500 1.1475 0.01967 0.01217 -0.0517 0.1201 1.0000
8.750 1.1593 0.02088 0.01306 -0.0497 0.0691 1.0000
9.000 1.1646 0.02248 0.01435 -0.0466 0.0425 1.0000
9.250 1.1733 0.02371 0.01561 -0.0439 0.0368 1.0000
9.500 1.1788 0.02512 0.01709 -0.0409 0.0339 1.0000
9.750 1.1870 0.02637 0.01848 -0.0384 0.0322 1.0000
10.000 1.1938 0.02777 0.01999 -0.0360 0.0307 1.0000
10.250 1.1999 0.02929 0.02159 -0.0337 0.0293 1.0000
10.500 1.2027 0.03116 0.02352 -0.0314 0.0276 1.0000
10.750 1.2048 0.03360 0.02600 -0.0291 0.0264 1.0000
11.000 1.2153 0.03520 0.02774 -0.0276 0.0259 1.0000
11.250 1.2262 0.03701 0.02969 -0.0261 0.0253 1.0000
11.500 1.2377 0.03897 0.03179 -0.0248 0.0247 1.0000
11.750 1.2495 0.04107 0.03406 -0.0235 0.0243 1.0000
12.000 1.2600 0.04338 0.03656 -0.0223 0.0239 1.0000
12.250 1.2672 0.04573 0.03909 -0.0210 0.0232 1.0000
12.500 1.2723 0.04828 0.04183 -0.0198 0.0227 1.0000
12.750 1.2743 0.05088 0.04462 -0.0187 0.0220 1.0000
13.000 1.2748 0.05368 0.04757 -0.0179 0.0214 1.0000
13.250 1.2726 0.05704 0.05116 -0.0170 0.0213 1.0000
13.500 1.2667 0.06079 0.05516 -0.0165 0.0212 1.0000
13.750 1.2575 0.06488 0.05950 -0.0163 0.0212 1.0000
14.000 1.2453 0.06943 0.06431 -0.0167 0.0212 1.0000
14.250 1.2297 0.07457 0.06971 -0.0177 0.0214 1.0000
14.500 1.2116 0.08035 0.07576 -0.0195 0.0215 1.0000
14.750 1.1881 0.08727 0.08297 -0.0225 0.0219 1.0000
15.000 1.1624 0.09528 0.09126 -0.0268 0.0222 1.0000
15.250 1.1350 0.10452 0.10075 -0.0325 0.0225 1.0000
15.500 1.1058 0.11513 0.11159 -0.0396 0.0229 1.0000
15.750 1.0730 0.12788 0.12454 -0.0484 0.0234 1.0000
16.000 1.0310 0.14482 0.14165 -0.0598 0.0245 1.0000
16.250 0.9947 0.16162 0.15847 -0.0697 0.0259 1.0000
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