HQ 1.5/10 AIRFOIL (hq1510-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.5/10 AIRFOIL (hq1510-il) Reynolds number: 500,000 Max Cl/Cd: 88.38 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1510-il-500000.txt Download as CSV file: xf-hq1510-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5109 0.08349 0.08126 -0.0314 1.0000 0.0211
-9.500 -0.5183 0.07477 0.07256 -0.0377 1.0000 0.0211
-9.250 -0.5341 0.06678 0.06451 -0.0451 1.0000 0.0211
-9.000 -0.5481 0.06229 0.05996 -0.0483 1.0000 0.0211
-8.750 -0.6224 0.04803 0.04549 -0.0494 1.0000 0.0139
-8.500 -0.6962 0.03040 0.02657 -0.0450 1.0000 0.0105
-8.250 -0.6916 0.02728 0.02310 -0.0426 1.0000 0.0106
-8.000 -0.6837 0.02505 0.02061 -0.0402 1.0000 0.0108
-7.750 -0.6765 0.02292 0.01819 -0.0372 1.0000 0.0107
-7.500 -0.6541 0.02150 0.01662 -0.0373 0.9984 0.0113
-7.250 -0.6200 0.02032 0.01528 -0.0393 0.9953 0.0121
-7.000 -0.5871 0.01876 0.01348 -0.0410 0.9918 0.0127
-6.750 -0.5549 0.01690 0.01138 -0.0423 0.9878 0.0131
-6.500 -0.5207 0.01551 0.00980 -0.0439 0.9846 0.0138
-6.250 -0.4845 0.01456 0.00871 -0.0458 0.9823 0.0146
-6.000 -0.4552 0.01287 0.00688 -0.0467 0.9764 0.0164
-5.750 -0.4202 0.01223 0.00621 -0.0484 0.9725 0.0186
-5.500 -0.3848 0.01157 0.00546 -0.0502 0.9690 0.0228
-5.250 -0.3545 0.01112 0.00501 -0.0507 0.9611 0.0293
-5.000 -0.3218 0.01063 0.00451 -0.0518 0.9555 0.0407
-4.750 -0.2942 0.01023 0.00412 -0.0519 0.9459 0.0509
-4.500 -0.2658 0.00989 0.00376 -0.0520 0.9377 0.0600
-4.250 -0.2388 0.00959 0.00343 -0.0518 0.9286 0.0706
-4.000 -0.2129 0.00926 0.00315 -0.0514 0.9190 0.0896
-3.750 -0.1875 0.00880 0.00284 -0.0511 0.9104 0.1313
-3.500 -0.1632 0.00823 0.00257 -0.0506 0.9011 0.2177
-3.250 -0.1394 0.00758 0.00231 -0.0502 0.8919 0.3295
-3.000 -0.1158 0.00699 0.00215 -0.0496 0.8836 0.4621
-2.750 -0.0904 0.00674 0.00208 -0.0490 0.8744 0.5334
-2.500 -0.0642 0.00662 0.00204 -0.0486 0.8662 0.5781
-2.250 -0.0375 0.00656 0.00198 -0.0481 0.8578 0.6103
-2.000 -0.0108 0.00652 0.00197 -0.0477 0.8491 0.6403
-1.750 0.0161 0.00650 0.00194 -0.0473 0.8415 0.6625
-1.500 0.0433 0.00648 0.00191 -0.0471 0.8326 0.6785
-1.250 0.0702 0.00649 0.00190 -0.0466 0.8234 0.7001
-1.000 0.0966 0.00648 0.00189 -0.0461 0.8137 0.7194
-0.750 0.1236 0.00646 0.00188 -0.0457 0.8031 0.7328
-0.500 0.1509 0.00646 0.00185 -0.0454 0.7937 0.7449
-0.250 0.1781 0.00648 0.00183 -0.0451 0.7848 0.7574
0.000 0.2052 0.00646 0.00184 -0.0448 0.7748 0.7699
0.250 0.2321 0.00646 0.00184 -0.0444 0.7646 0.7821
0.500 0.2591 0.00646 0.00183 -0.0441 0.7547 0.7924
0.750 0.2865 0.00646 0.00182 -0.0439 0.7449 0.8019
1.000 0.3138 0.00646 0.00183 -0.0436 0.7345 0.8118
1.250 0.3408 0.00645 0.00185 -0.0433 0.7244 0.8215
1.500 0.3678 0.00646 0.00185 -0.0430 0.7138 0.8321
1.750 0.3946 0.00646 0.00187 -0.0427 0.7020 0.8437
2.000 0.4212 0.00646 0.00190 -0.0423 0.6896 0.8563
2.250 0.4473 0.00645 0.00194 -0.0418 0.6761 0.8703
2.500 0.4729 0.00644 0.00197 -0.0411 0.6613 0.8868
2.750 0.4982 0.00645 0.00201 -0.0404 0.6446 0.9081
3.000 0.5266 0.00645 0.00206 -0.0403 0.6229 0.9378
3.250 0.5642 0.00656 0.00212 -0.0423 0.5910 0.9707
3.500 0.6010 0.00680 0.00218 -0.0445 0.5422 1.0000
3.750 0.6239 0.00716 0.00230 -0.0437 0.4874 1.0000
4.000 0.6467 0.00760 0.00250 -0.0429 0.4281 1.0000
4.250 0.6699 0.00807 0.00272 -0.0423 0.3739 1.0000
4.500 0.6935 0.00852 0.00297 -0.0417 0.3250 1.0000
4.750 0.7175 0.00895 0.00323 -0.0412 0.2886 1.0000
5.000 0.7419 0.00935 0.00351 -0.0407 0.2606 1.0000
5.250 0.7664 0.00973 0.00380 -0.0403 0.2356 1.0000
5.500 0.7912 0.01009 0.00407 -0.0399 0.2099 1.0000
5.750 0.8155 0.01049 0.00436 -0.0394 0.1792 1.0000
6.000 0.8386 0.01103 0.00471 -0.0388 0.1376 1.0000
6.250 0.8601 0.01175 0.00521 -0.0380 0.0954 1.0000
6.500 0.8823 0.01241 0.00573 -0.0373 0.0713 1.0000
6.750 0.9053 0.01295 0.00623 -0.0366 0.0587 1.0000
7.000 0.9275 0.01359 0.00683 -0.0358 0.0502 1.0000
7.250 0.9519 0.01394 0.00723 -0.0354 0.0456 1.0000
7.500 0.9741 0.01453 0.00780 -0.0346 0.0393 1.0000
7.750 0.9982 0.01489 0.00823 -0.0341 0.0360 1.0000
8.000 1.0218 0.01530 0.00867 -0.0336 0.0328 1.0000
8.250 1.0425 0.01601 0.00939 -0.0326 0.0282 1.0000
8.500 1.0677 0.01620 0.00964 -0.0324 0.0252 1.0000
8.750 1.0900 0.01671 0.01012 -0.0317 0.0180 1.0000
9.000 1.1087 0.01762 0.01102 -0.0304 0.0123 1.0000
9.250 1.1278 0.01845 0.01188 -0.0292 0.0100 1.0000
9.500 1.1419 0.01974 0.01332 -0.0273 0.0088 1.0000
9.750 1.1586 0.02069 0.01439 -0.0258 0.0085 1.0000
10.000 1.1736 0.02173 0.01556 -0.0241 0.0081 1.0000
10.250 1.1876 0.02280 0.01674 -0.0223 0.0077 1.0000
10.500 1.2015 0.02377 0.01781 -0.0206 0.0073 1.0000
10.750 1.2116 0.02481 0.01895 -0.0182 0.0070 1.0000
11.000 1.2183 0.02603 0.02028 -0.0156 0.0069 1.0000
11.250 1.2251 0.02722 0.02155 -0.0132 0.0065 1.0000
11.500 1.2286 0.02874 0.02320 -0.0107 0.0063 1.0000
11.750 1.2277 0.03075 0.02536 -0.0081 0.0062 1.0000
12.000 1.2257 0.03300 0.02779 -0.0058 0.0061 1.0000
12.250 1.2194 0.03584 0.03084 -0.0037 0.0060 1.0000
12.500 1.2152 0.03858 0.03380 -0.0022 0.0059 1.0000
12.750 1.2061 0.04205 0.03749 -0.0010 0.0059 1.0000
13.000 1.1979 0.04555 0.04119 -0.0004 0.0059 1.0000
13.250 1.1933 0.04875 0.04456 -0.0005 0.0059 1.0000
13.500 1.1833 0.05287 0.04888 -0.0011 0.0058 1.0000
13.750 1.1746 0.05708 0.05326 -0.0023 0.0058 1.0000
14.000 1.1570 0.06297 0.05935 -0.0044 0.0058 1.0000
14.250 1.1472 0.06810 0.06465 -0.0070 0.0058 1.0000
14.500 1.1283 0.07522 0.07197 -0.0108 0.0058 1.0000
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