HQ 2.5/12 AIRFOIL (hq2512-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: HQ 2.5/12 AIRFOIL (hq2512-il) Reynolds number: 1,000,000 Max Cl/Cd: 92.87 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq2512-il-1000000-n5.txt Download as CSV file: xf-hq2512-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/12 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
------ -------- --------- --------- -------- -------- --------
-15.000 -0.6096 0.12108 0.11925 -0.0269 1.0000 0.0040
-14.750 -0.9370 0.05048 0.04786 -0.0689 1.0000 0.0032
-14.500 -0.9570 0.04380 0.04101 -0.0739 1.0000 0.0031
-14.250 -0.9699 0.03904 0.03609 -0.0769 1.0000 0.0031
-14.000 -0.9841 0.03478 0.03167 -0.0788 1.0000 0.0031
-13.750 -0.9895 0.03201 0.02877 -0.0794 1.0000 0.0031
-13.500 -0.9913 0.02998 0.02664 -0.0789 1.0000 0.0032
-13.250 -0.9998 0.02819 0.02474 -0.0768 1.0000 0.0032
-13.000 -1.0030 0.02683 0.02327 -0.0739 1.0000 0.0033
-12.750 -0.9972 0.02570 0.02205 -0.0718 1.0000 0.0034
-12.500 -0.9782 0.02447 0.02072 -0.0722 0.9971 0.0035
-12.250 -0.9559 0.02319 0.01932 -0.0730 0.9899 0.0035
-12.000 -0.9290 0.02205 0.01806 -0.0745 0.9856 0.0037
-11.750 -0.9008 0.02104 0.01695 -0.0760 0.9812 0.0038
-11.500 -0.8727 0.02006 0.01586 -0.0773 0.9755 0.0039
-11.250 -0.8424 0.01924 0.01492 -0.0789 0.9702 0.0042
-11.000 -0.8145 0.01849 0.01408 -0.0798 0.9613 0.0044
-10.750 -0.7877 0.01775 0.01323 -0.0805 0.9511 0.0045
-10.500 -0.7629 0.01707 0.01244 -0.0806 0.9389 0.0046
-10.250 -0.7400 0.01641 0.01166 -0.0802 0.9253 0.0047
-10.000 -0.7192 0.01558 0.01067 -0.0795 0.9128 0.0049
-9.750 -0.6970 0.01493 0.00991 -0.0788 0.9024 0.0052
-9.500 -0.6738 0.01439 0.00928 -0.0783 0.8931 0.0054
-9.250 -0.6500 0.01394 0.00873 -0.0779 0.8844 0.0057
-9.000 -0.6257 0.01350 0.00821 -0.0775 0.8759 0.0059
-8.750 -0.6012 0.01308 0.00770 -0.0771 0.8684 0.0061
-8.500 -0.5762 0.01270 0.00724 -0.0767 0.8608 0.0065
-8.250 -0.5510 0.01233 0.00680 -0.0764 0.8540 0.0067
-8.000 -0.5254 0.01199 0.00638 -0.0761 0.8470 0.0069
-7.750 -0.5002 0.01159 0.00589 -0.0758 0.8407 0.0075
-7.500 -0.4745 0.01121 0.00547 -0.0756 0.8340 0.0082
-7.250 -0.4486 0.01091 0.00511 -0.0753 0.8277 0.0088
-7.000 -0.4221 0.01061 0.00476 -0.0752 0.8214 0.0094
-6.750 -0.3956 0.01036 0.00444 -0.0750 0.8150 0.0100
-6.500 -0.3691 0.01005 0.00411 -0.0749 0.8088 0.0118
-6.250 -0.3424 0.00979 0.00381 -0.0747 0.8022 0.0141
-6.000 -0.3158 0.00953 0.00356 -0.0746 0.7963 0.0187
-5.750 -0.2888 0.00931 0.00336 -0.0745 0.7898 0.0239
-5.500 -0.2614 0.00917 0.00317 -0.0744 0.7839 0.0262
-5.250 -0.2339 0.00900 0.00297 -0.0744 0.7777 0.0284
-5.000 -0.2066 0.00883 0.00277 -0.0743 0.7715 0.0315
-4.750 -0.1790 0.00868 0.00261 -0.0743 0.7656 0.0339
-4.500 -0.1513 0.00856 0.00245 -0.0743 0.7592 0.0359
-4.250 -0.1238 0.00840 0.00229 -0.0743 0.7533 0.0416
-4.000 -0.0960 0.00827 0.00215 -0.0743 0.7468 0.0470
-3.750 -0.0685 0.00813 0.00200 -0.0743 0.7406 0.0541
-3.500 -0.0406 0.00801 0.00187 -0.0743 0.7336 0.0612
-3.250 -0.0131 0.00789 0.00175 -0.0742 0.7263 0.0715
-3.000 0.0147 0.00775 0.00163 -0.0742 0.7183 0.0848
-2.750 0.0423 0.00762 0.00152 -0.0742 0.7114 0.1036
-2.500 0.0699 0.00743 0.00142 -0.0743 0.7048 0.1317
-2.250 0.0973 0.00724 0.00132 -0.0743 0.6986 0.1721
-2.000 0.1249 0.00705 0.00123 -0.0743 0.6912 0.2123
-1.750 0.1521 0.00688 0.00116 -0.0743 0.6830 0.2530
-1.500 0.1790 0.00659 0.00106 -0.0743 0.6733 0.3256
-1.250 0.2057 0.00631 0.00100 -0.0743 0.6641 0.4072
-1.000 0.2325 0.00611 0.00096 -0.0742 0.6545 0.4771
-0.750 0.2598 0.00598 0.00096 -0.0741 0.6438 0.5298
-0.500 0.2871 0.00591 0.00099 -0.0740 0.6325 0.5760
-0.250 0.3148 0.00592 0.00100 -0.0740 0.6221 0.5974
0.250 0.3707 0.00599 0.00104 -0.0739 0.6010 0.6205
0.500 0.3986 0.00603 0.00106 -0.0739 0.5891 0.6304
0.750 0.4261 0.00610 0.00110 -0.0738 0.5736 0.6410
1.000 0.4534 0.00619 0.00114 -0.0737 0.5575 0.6514
1.250 0.4806 0.00629 0.00119 -0.0736 0.5402 0.6583
1.500 0.5077 0.00641 0.00125 -0.0734 0.5195 0.6643
1.750 0.5345 0.00655 0.00132 -0.0733 0.4986 0.6698
2.000 0.5611 0.00671 0.00140 -0.0731 0.4756 0.6755
2.250 0.5872 0.00692 0.00151 -0.0728 0.4473 0.6815
2.500 0.6126 0.00717 0.00164 -0.0724 0.4149 0.6872
2.750 0.6382 0.00742 0.00178 -0.0720 0.3849 0.6936
3.000 0.6640 0.00764 0.00191 -0.0717 0.3587 0.6997
3.250 0.6892 0.00791 0.00207 -0.0714 0.3291 0.7067
3.500 0.7144 0.00818 0.00224 -0.0710 0.3013 0.7132
3.750 0.7400 0.00839 0.00240 -0.0706 0.2831 0.7202
4.000 0.7659 0.00860 0.00255 -0.0704 0.2682 0.7271
4.250 0.7914 0.00881 0.00273 -0.0700 0.2521 0.7343
4.500 0.8169 0.00903 0.00291 -0.0697 0.2362 0.7415
4.750 0.8423 0.00924 0.00309 -0.0693 0.2215 0.7491
5.000 0.8679 0.00943 0.00326 -0.0690 0.2109 0.7570
5.250 0.8934 0.00962 0.00344 -0.0686 0.1993 0.7654
5.500 0.9174 0.00991 0.00368 -0.0681 0.1800 0.7742
5.750 0.9419 0.01016 0.00389 -0.0676 0.1664 0.7838
6.000 0.9658 0.01043 0.00413 -0.0670 0.1513 0.7942
6.250 0.9888 0.01074 0.00440 -0.0663 0.1337 0.8057
6.500 1.0104 0.01115 0.00473 -0.0653 0.1125 0.8194
6.750 1.0331 0.01143 0.00502 -0.0645 0.1013 0.8364
7.000 1.0538 0.01176 0.00535 -0.0634 0.0850 0.8601
7.250 1.0746 0.01199 0.00570 -0.0622 0.0699 0.9433
7.500 1.1012 0.01252 0.00613 -0.0625 0.0534 1.0000
7.750 1.1218 0.01302 0.00656 -0.0614 0.0404 1.0000
8.000 1.1407 0.01362 0.00707 -0.0601 0.0257 1.0000
8.250 1.1538 0.01463 0.00793 -0.0579 0.0049 1.0000
8.500 1.1740 0.01508 0.00840 -0.0568 0.0037 1.0000
8.750 1.1945 0.01547 0.00882 -0.0557 0.0035 1.0000
9.000 1.2141 0.01590 0.00928 -0.0545 0.0032 1.0000
9.250 1.2316 0.01634 0.00978 -0.0529 0.0030 1.0000
9.500 1.2479 0.01680 0.01027 -0.0511 0.0029 1.0000
9.750 1.2627 0.01734 0.01086 -0.0491 0.0027 1.0000
10.000 1.2772 0.01791 0.01147 -0.0471 0.0026 1.0000
10.250 1.2912 0.01852 0.01214 -0.0452 0.0024 1.0000
10.500 1.3037 0.01924 0.01293 -0.0431 0.0023 1.0000
10.750 1.3157 0.02001 0.01376 -0.0411 0.0022 1.0000
11.000 1.3286 0.02073 0.01454 -0.0393 0.0022 1.0000
11.250 1.3412 0.02150 0.01537 -0.0375 0.0022 1.0000
11.500 1.3539 0.02228 0.01621 -0.0358 0.0021 1.0000
11.750 1.3649 0.02318 0.01718 -0.0341 0.0020 1.0000
12.000 1.3749 0.02420 0.01826 -0.0324 0.0020 1.0000
12.250 1.3843 0.02528 0.01942 -0.0307 0.0019 1.0000
12.500 1.3928 0.02646 0.02068 -0.0291 0.0019 1.0000
12.750 1.4002 0.02777 0.02207 -0.0275 0.0019 1.0000
13.000 1.4065 0.02921 0.02359 -0.0259 0.0019 1.0000
13.250 1.4131 0.03069 0.02515 -0.0246 0.0017 1.0000
13.500 1.4173 0.03242 0.02697 -0.0233 0.0017 1.0000
13.750 1.4212 0.03424 0.02889 -0.0221 0.0017 1.0000
14.000 1.4248 0.03616 0.03090 -0.0211 0.0016 1.0000
14.250 1.4257 0.03842 0.03326 -0.0202 0.0016 1.0000
14.500 1.4257 0.04086 0.03580 -0.0195 0.0016 1.0000
14.750 1.4254 0.04344 0.03847 -0.0190 0.0015 1.0000
15.000 1.4219 0.04650 0.04165 -0.0187 0.0016 1.0000
15.250 1.4194 0.04960 0.04485 -0.0187 0.0015 1.0000
15.500 1.4157 0.05301 0.04836 -0.0190 0.0014 1.0000
15.750 1.4093 0.05694 0.05243 -0.0197 0.0015 1.0000
16.000 1.4022 0.06118 0.05679 -0.0207 0.0015 1.0000
16.250 1.3945 0.06574 0.06146 -0.0220 0.0014 1.0000
16.500 1.3824 0.07121 0.06707 -0.0240 0.0014 1.0000
16.750 1.3728 0.07655 0.07253 -0.0261 0.0014 1.0000
17.000 1.3571 0.08317 0.07929 -0.0291 0.0014 1.0000
17.250 1.3441 0.08959 0.08584 -0.0321 0.0014 1.0000
17.500 1.3241 0.09754 0.09394 -0.0360 0.0014 1.0000
17.750 1.3097 0.10464 0.10117 -0.0396 0.0014 1.0000
18.000 1.2927 0.11249 0.10915 -0.0437 0.0014 1.0000
18.250 1.2732 0.12095 0.11775 -0.0483 0.0014 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 2.5/12 AIRFOIL (hq2512-il)