HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.51 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq2510-il-1000000-n5.txt Download as CSV file: xf-hq2510-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.500 -0.4490 0.11444 0.11275 -0.0238 1.0000 0.0023
-10.000 -0.7038 0.02739 0.02474 -0.0744 0.9825 0.0019
-9.750 -0.6837 0.02363 0.02056 -0.0767 0.9753 0.0019
-9.500 -0.6586 0.02144 0.01808 -0.0781 0.9688 0.0021
-9.250 -0.6334 0.01969 0.01604 -0.0791 0.9605 0.0021
-9.000 -0.6091 0.01836 0.01451 -0.0794 0.9511 0.0021
-8.750 -0.5886 0.01634 0.01214 -0.0792 0.9398 0.0023
-8.500 -0.5658 0.01518 0.01077 -0.0788 0.9291 0.0025
-8.250 -0.5413 0.01454 0.01002 -0.0785 0.9188 0.0027
-8.000 -0.5165 0.01398 0.00935 -0.0782 0.9095 0.0028
-7.750 -0.4915 0.01346 0.00872 -0.0779 0.9013 0.0030
-7.500 -0.4663 0.01295 0.00809 -0.0776 0.8929 0.0032
-7.250 -0.4410 0.01241 0.00744 -0.0774 0.8853 0.0033
-7.000 -0.4156 0.01190 0.00679 -0.0771 0.8775 0.0036
-6.750 -0.3897 0.01144 0.00624 -0.0768 0.8703 0.0038
-6.500 -0.3637 0.01103 0.00572 -0.0766 0.8633 0.0040
-6.250 -0.3372 0.01065 0.00525 -0.0764 0.8565 0.0042
-6.000 -0.3113 0.01016 0.00466 -0.0761 0.8498 0.0046
-5.750 -0.2847 0.00978 0.00424 -0.0760 0.8434 0.0053
-5.500 -0.2578 0.00954 0.00394 -0.0759 0.8367 0.0060
-5.250 -0.2307 0.00927 0.00359 -0.0758 0.8306 0.0067
-5.000 -0.2036 0.00902 0.00328 -0.0757 0.8238 0.0071
-4.750 -0.1763 0.00877 0.00296 -0.0756 0.8176 0.0076
-4.500 -0.1491 0.00849 0.00262 -0.0755 0.8107 0.0095
-4.250 -0.1217 0.00828 0.00237 -0.0754 0.8042 0.0117
-4.000 -0.0943 0.00805 0.00216 -0.0754 0.7972 0.0184
-3.750 -0.0668 0.00788 0.00198 -0.0754 0.7904 0.0276
-3.500 -0.0390 0.00778 0.00183 -0.0753 0.7827 0.0302
-3.250 -0.0114 0.00765 0.00169 -0.0753 0.7749 0.0359
-2.750 0.0438 0.00742 0.00144 -0.0752 0.7565 0.0543
-2.500 0.0714 0.00727 0.00132 -0.0752 0.7481 0.0727
-2.250 0.0990 0.00715 0.00122 -0.0752 0.7400 0.0923
-2.000 0.1267 0.00699 0.00112 -0.0752 0.7319 0.1220
-1.750 0.1539 0.00679 0.00103 -0.0752 0.7221 0.1720
-1.500 0.1811 0.00660 0.00095 -0.0752 0.7111 0.2277
-1.250 0.2082 0.00637 0.00087 -0.0752 0.6993 0.2930
-1.000 0.2348 0.00604 0.00080 -0.0752 0.6869 0.3985
-0.750 0.2620 0.00590 0.00077 -0.0752 0.6742 0.4554
-0.500 0.2891 0.00576 0.00077 -0.0751 0.6626 0.5172
-0.250 0.3158 0.00560 0.00081 -0.0750 0.6524 0.6004
0.000 0.3431 0.00559 0.00084 -0.0748 0.6403 0.6339
0.250 0.3707 0.00563 0.00086 -0.0748 0.6261 0.6488
0.500 0.3981 0.00568 0.00089 -0.0747 0.6090 0.6628
0.750 0.4253 0.00576 0.00092 -0.0745 0.5905 0.6744
1.000 0.4526 0.00586 0.00096 -0.0744 0.5718 0.6831
1.250 0.4798 0.00595 0.00101 -0.0743 0.5522 0.6906
1.500 0.5066 0.00610 0.00107 -0.0741 0.5270 0.6979
1.750 0.5326 0.00631 0.00116 -0.0738 0.4916 0.7057
2.000 0.5584 0.00655 0.00126 -0.0735 0.4536 0.7132
2.250 0.5844 0.00679 0.00139 -0.0732 0.4209 0.7214
2.500 0.6103 0.00701 0.00151 -0.0729 0.3903 0.7297
2.750 0.6365 0.00722 0.00164 -0.0727 0.3645 0.7383
3.000 0.6623 0.00745 0.00178 -0.0724 0.3342 0.7477
3.250 0.6879 0.00770 0.00195 -0.0721 0.3040 0.7571
3.500 0.7135 0.00795 0.00212 -0.0718 0.2784 0.7669
4.000 0.7649 0.00838 0.00245 -0.0712 0.2380 0.7885
4.250 0.7910 0.00855 0.00263 -0.0709 0.2244 0.8004
4.500 0.8154 0.00885 0.00284 -0.0704 0.1943 0.8143
4.750 0.8380 0.00931 0.00312 -0.0697 0.1505 0.8304
5.000 0.8614 0.00964 0.00337 -0.0690 0.1238 0.8505
5.250 0.8834 0.00988 0.00363 -0.0679 0.1025 0.8875
5.500 0.9116 0.01031 0.00401 -0.0684 0.0666 1.0000
5.750 0.9340 0.01090 0.00442 -0.0677 0.0367 1.0000
6.000 0.9541 0.01176 0.00508 -0.0665 0.0043 1.0000
6.250 0.9788 0.01208 0.00543 -0.0660 0.0031 1.0000
6.500 1.0036 0.01239 0.00577 -0.0656 0.0029 1.0000
6.750 1.0281 0.01271 0.00613 -0.0651 0.0027 1.0000
7.000 1.0523 0.01307 0.00652 -0.0646 0.0025 1.0000
7.250 1.0760 0.01346 0.00698 -0.0640 0.0023 1.0000
7.500 1.0992 0.01389 0.00745 -0.0634 0.0021 1.0000
7.750 1.1216 0.01439 0.00800 -0.0626 0.0019 1.0000
8.000 1.1437 0.01490 0.00856 -0.0618 0.0018 1.0000
8.250 1.1646 0.01553 0.00925 -0.0607 0.0016 1.0000
8.500 1.1837 0.01632 0.01013 -0.0594 0.0015 1.0000
8.750 1.1973 0.01762 0.01158 -0.0573 0.0014 1.0000
9.000 1.2148 0.01842 0.01246 -0.0558 0.0014 1.0000
9.250 1.2317 0.01924 0.01338 -0.0543 0.0014 1.0000
9.500 1.2478 0.02007 0.01429 -0.0526 0.0013 1.0000
9.750 1.2619 0.02098 0.01529 -0.0507 0.0013 1.0000
10.000 1.2724 0.02191 0.01631 -0.0482 0.0013 1.0000
10.250 1.2799 0.02297 0.01747 -0.0453 0.0013 1.0000
10.500 1.2870 0.02406 0.01866 -0.0425 0.0013 1.0000
10.750 1.2931 0.02527 0.01998 -0.0398 0.0013 1.0000
11.000 1.2979 0.02664 0.02146 -0.0371 0.0013 1.0000
11.250 1.3020 0.02813 0.02306 -0.0346 0.0013 1.0000
11.500 1.3049 0.02977 0.02482 -0.0323 0.0013 1.0000
11.750 1.3060 0.03167 0.02685 -0.0301 0.0013 1.0000
12.000 1.3082 0.03354 0.02884 -0.0283 0.0013 1.0000
12.250 1.3076 0.03578 0.03122 -0.0265 0.0013 1.0000
12.500 1.3052 0.03831 0.03389 -0.0251 0.0013 1.0000
12.750 1.3021 0.04104 0.03679 -0.0239 0.0013 1.0000
13.000 1.2967 0.04419 0.04009 -0.0231 0.0013 1.0000
13.250 1.2965 0.04684 0.04286 -0.0228 0.0012 1.0000
13.500 1.2896 0.05050 0.04667 -0.0228 0.0012 1.0000
13.750 1.2811 0.05461 0.05094 -0.0233 0.0012 1.0000
14.000 1.2717 0.05911 0.05559 -0.0243 0.0012 1.0000
14.250 1.2617 0.06400 0.06062 -0.0259 0.0012 1.0000
14.500 1.2471 0.07005 0.06684 -0.0283 0.0012 1.0000
14.750 1.2376 0.07565 0.07257 -0.0310 0.0012 1.0000
15.000 1.2295 0.08136 0.07841 -0.0340 0.0012 1.0000
15.250 1.2148 0.08870 0.08590 -0.0380 0.0012 1.0000
15.500 1.1947 0.09752 0.09489 -0.0429 0.0012 1.0000
15.750 1.1841 0.10473 0.10221 -0.0471 0.0012 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/10 AIRFOIL (hq2510-il)