HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il) Reynolds number: 500,000 Max Cl/Cd: 85.26 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq2510-il-500000-n5.txt Download as CSV file: xf-hq2510-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4453 0.08171 0.07951 -0.0380 1.0000 0.0041
-9.250 -0.4553 0.07615 0.07400 -0.0406 1.0000 0.0041
-9.000 -0.4686 0.07042 0.06834 -0.0435 1.0000 0.0040
-8.750 -0.4819 0.06574 0.06371 -0.0459 1.0000 0.0039
-8.500 -0.4937 0.05523 0.05315 -0.0597 0.9890 0.0038
-8.250 -0.5276 0.02578 0.02212 -0.0767 0.9662 0.0038
-8.000 -0.5036 0.02236 0.01818 -0.0783 0.9595 0.0040
-7.750 -0.4793 0.02039 0.01586 -0.0787 0.9513 0.0042
-7.500 -0.4547 0.01810 0.01318 -0.0792 0.9440 0.0044
-7.250 -0.4305 0.01665 0.01146 -0.0792 0.9351 0.0047
-7.000 -0.4045 0.01579 0.01045 -0.0793 0.9274 0.0050
-6.750 -0.3778 0.01533 0.00991 -0.0793 0.9196 0.0055
-6.500 -0.3519 0.01471 0.00917 -0.0792 0.9119 0.0061
-6.250 -0.3263 0.01395 0.00824 -0.0789 0.9045 0.0067
-6.000 -0.3010 0.01325 0.00740 -0.0785 0.8967 0.0071
-5.750 -0.2756 0.01252 0.00651 -0.0781 0.8896 0.0075
-5.500 -0.2508 0.01167 0.00552 -0.0777 0.8817 0.0084
-5.250 -0.2247 0.01121 0.00494 -0.0774 0.8749 0.0094
-5.000 -0.1983 0.01083 0.00448 -0.0772 0.8677 0.0106
-4.750 -0.1715 0.01055 0.00411 -0.0770 0.8613 0.0120
-4.500 -0.1450 0.01010 0.00361 -0.0768 0.8542 0.0157
-4.250 -0.1181 0.00980 0.00330 -0.0766 0.8479 0.0227
-4.000 -0.0908 0.00962 0.00308 -0.0765 0.8409 0.0296
-3.500 -0.0363 0.00923 0.00263 -0.0763 0.8276 0.0447
-3.250 -0.0089 0.00907 0.00244 -0.0762 0.8212 0.0537
-3.000 0.0184 0.00886 0.00226 -0.0762 0.8140 0.0688
-2.500 0.0729 0.00843 0.00191 -0.0761 0.7991 0.1195
-2.250 0.0997 0.00812 0.00175 -0.0760 0.7912 0.1795
-2.000 0.1264 0.00778 0.00161 -0.0760 0.7824 0.2555
-1.750 0.1526 0.00732 0.00147 -0.0760 0.7741 0.3676
-1.500 0.1791 0.00701 0.00139 -0.0759 0.7663 0.4638
-1.250 0.2060 0.00682 0.00138 -0.0758 0.7583 0.5305
-1.000 0.2329 0.00672 0.00137 -0.0755 0.7495 0.5826
-0.750 0.2594 0.00665 0.00141 -0.0751 0.7389 0.6367
-0.500 0.2863 0.00663 0.00142 -0.0748 0.7270 0.6672
-0.250 0.3134 0.00665 0.00141 -0.0746 0.7145 0.6829
0.000 0.3407 0.00667 0.00141 -0.0744 0.7023 0.6970
0.250 0.3681 0.00670 0.00142 -0.0742 0.6914 0.7080
0.500 0.3955 0.00674 0.00143 -0.0741 0.6809 0.7162
0.750 0.4228 0.00679 0.00144 -0.0739 0.6688 0.7243
1.250 0.4770 0.00691 0.00150 -0.0735 0.6400 0.7411
1.750 0.5308 0.00705 0.00159 -0.0731 0.6070 0.7593
2.000 0.5573 0.00715 0.00165 -0.0728 0.5881 0.7687
2.250 0.5833 0.00728 0.00174 -0.0724 0.5633 0.7790
2.500 0.6084 0.00747 0.00183 -0.0719 0.5320 0.7898
2.750 0.6332 0.00768 0.00194 -0.0713 0.4976 0.8012
3.000 0.6576 0.00792 0.00209 -0.0707 0.4626 0.8136
3.250 0.6818 0.00817 0.00226 -0.0701 0.4285 0.8279
3.500 0.7059 0.00840 0.00243 -0.0694 0.3986 0.8448
3.750 0.7290 0.00864 0.00262 -0.0685 0.3659 0.8673
4.000 0.7518 0.00883 0.00281 -0.0675 0.3321 0.9121
4.250 0.7852 0.00921 0.00307 -0.0691 0.2941 1.0000
4.500 0.8097 0.00958 0.00332 -0.0686 0.2672 1.0000
4.750 0.8351 0.00986 0.00355 -0.0683 0.2503 1.0000
5.000 0.8600 0.01019 0.00380 -0.0680 0.2290 1.0000
5.250 0.8831 0.01068 0.00411 -0.0674 0.1910 1.0000
5.500 0.9054 0.01126 0.00445 -0.0667 0.1514 1.0000
5.750 0.9282 0.01177 0.00483 -0.0661 0.1213 1.0000
6.000 0.9510 0.01229 0.00521 -0.0654 0.0958 1.0000
6.250 0.9737 0.01280 0.00562 -0.0647 0.0739 1.0000
6.500 0.9937 0.01359 0.00619 -0.0637 0.0402 1.0000
6.750 1.0114 0.01464 0.00702 -0.0623 0.0072 1.0000
7.000 1.0340 0.01515 0.00757 -0.0615 0.0050 1.0000
7.250 1.0563 0.01565 0.00816 -0.0607 0.0042 1.0000
7.500 1.0782 0.01620 0.00877 -0.0598 0.0037 1.0000
7.750 1.0993 0.01682 0.00949 -0.0588 0.0033 1.0000
8.000 1.1194 0.01752 0.01030 -0.0577 0.0031 1.0000
8.250 1.1388 0.01827 0.01115 -0.0565 0.0029 1.0000
8.500 1.1576 0.01901 0.01199 -0.0552 0.0028 1.0000
8.750 1.1754 0.01982 0.01289 -0.0538 0.0026 1.0000
9.000 1.1918 0.02069 0.01387 -0.0522 0.0026 1.0000
9.250 1.2063 0.02166 0.01494 -0.0504 0.0025 1.0000
9.500 1.2179 0.02268 0.01611 -0.0481 0.0024 1.0000
9.750 1.2270 0.02374 0.01728 -0.0454 0.0024 1.0000
10.000 1.2352 0.02486 0.01851 -0.0428 0.0024 1.0000
10.250 1.2410 0.02620 0.01997 -0.0400 0.0023 1.0000
10.500 1.2470 0.02757 0.02147 -0.0375 0.0023 1.0000
10.750 1.2524 0.02905 0.02306 -0.0352 0.0023 1.0000
11.000 1.2557 0.03078 0.02494 -0.0328 0.0022 1.0000
11.250 1.2590 0.03258 0.02688 -0.0308 0.0022 1.0000
11.500 1.2613 0.03456 0.02901 -0.0289 0.0022 1.0000
11.750 1.2618 0.03681 0.03142 -0.0271 0.0022 1.0000
12.000 1.2627 0.03909 0.03385 -0.0257 0.0022 1.0000
12.250 1.2612 0.04173 0.03666 -0.0245 0.0022 1.0000
12.500 1.2581 0.04468 0.03978 -0.0235 0.0022 1.0000
12.750 1.2536 0.04792 0.04320 -0.0229 0.0022 1.0000
13.000 1.2475 0.05153 0.04698 -0.0227 0.0022 1.0000
13.250 1.2419 0.05521 0.05084 -0.0230 0.0022 1.0000
13.500 1.2321 0.05971 0.05552 -0.0238 0.0022 1.0000
13.750 1.2216 0.06459 0.06062 -0.0251 0.0022 1.0000
14.000 1.2101 0.06996 0.06617 -0.0271 0.0022 1.0000
14.250 1.1982 0.07575 0.07214 -0.0296 0.0022 1.0000
14.500 1.1832 0.08260 0.07917 -0.0330 0.0022 1.0000
14.750 1.1700 0.08956 0.08629 -0.0368 0.0022 1.0000
15.000 1.1565 0.09695 0.09384 -0.0410 0.0023 1.0000
15.250 1.1410 0.10519 0.10224 -0.0458 0.0023 1.0000
15.500 1.1250 0.11391 0.11111 -0.0510 0.0023 1.0000
15.750 1.1100 0.12268 0.12001 -0.0562 0.0023 1.0000
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Polar data table (+)
Polar graphs
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