HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il) Reynolds number: 200,000 Max Cl/Cd: 72.14 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq2510-il-200000-n5.txt Download as CSV file: xf-hq2510-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4136 0.08718 0.08371 -0.0379 1.0000 0.0114
-9.000 -0.4166 0.08298 0.07956 -0.0395 1.0000 0.0113
-8.750 -0.4206 0.07912 0.07577 -0.0409 1.0000 0.0109
-8.500 -0.4274 0.07488 0.07160 -0.0426 1.0000 0.0108
-8.250 -0.4407 0.07032 0.06713 -0.0442 1.0000 0.0108
-8.000 -0.4597 0.06630 0.06321 -0.0457 1.0000 0.0105
-7.750 -0.4737 0.06122 0.05812 -0.0489 0.9986 0.0103
-7.500 -0.4606 0.05187 0.04849 -0.0593 0.9881 0.0105
-7.250 -0.4455 0.04362 0.03983 -0.0658 0.9792 0.0107
-7.000 -0.4258 0.03560 0.03120 -0.0706 0.9727 0.0108
-6.750 -0.4058 0.02982 0.02473 -0.0725 0.9647 0.0109
-6.500 -0.3785 0.02548 0.01971 -0.0742 0.9599 0.0111
-6.250 -0.3520 0.02268 0.01639 -0.0747 0.9537 0.0115
-6.000 -0.3227 0.02064 0.01395 -0.0754 0.9485 0.0120
-5.750 -0.2928 0.01838 0.01141 -0.0764 0.9446 0.0132
-5.500 -0.2656 0.01779 0.01074 -0.0766 0.9373 0.0152
-5.250 -0.2347 0.01683 0.00958 -0.0773 0.9327 0.0176
-5.000 -0.2071 0.01583 0.00840 -0.0773 0.9263 0.0191
-4.750 -0.1789 0.01479 0.00731 -0.0777 0.9203 0.0226
-4.500 -0.1495 0.01428 0.00668 -0.0780 0.9147 0.0288
-4.250 -0.1219 0.01370 0.00609 -0.0781 0.9080 0.0394
-4.000 -0.0923 0.01324 0.00558 -0.0785 0.9029 0.0500
-3.750 -0.0657 0.01292 0.00522 -0.0784 0.8954 0.0629
-3.500 -0.0367 0.01250 0.00477 -0.0787 0.8901 0.0755
-3.250 -0.0104 0.01215 0.00441 -0.0784 0.8824 0.0922
-3.000 0.0175 0.01172 0.00405 -0.0786 0.8766 0.1264
-2.750 0.0429 0.01118 0.00378 -0.0784 0.8690 0.2098
-2.500 0.0685 0.01043 0.00351 -0.0784 0.8630 0.3556
-2.250 0.0922 0.00991 0.00351 -0.0778 0.8553 0.5090
-2.000 0.1182 0.00973 0.00351 -0.0771 0.8491 0.5895
-1.750 0.1441 0.00968 0.00350 -0.0765 0.8412 0.6308
-1.500 0.1708 0.00964 0.00346 -0.0759 0.8347 0.6690
-1.250 0.1956 0.00964 0.00352 -0.0750 0.8261 0.7077
-1.000 0.2217 0.00961 0.00349 -0.0743 0.8187 0.7309
-0.750 0.2481 0.00960 0.00345 -0.0737 0.8102 0.7473
-0.500 0.2749 0.00957 0.00339 -0.0733 0.8022 0.7588
-0.250 0.3022 0.00955 0.00332 -0.0730 0.7945 0.7686
0.000 0.3291 0.00955 0.00328 -0.0728 0.7858 0.7789
0.250 0.3559 0.00952 0.00323 -0.0723 0.7764 0.7890
0.500 0.3823 0.00948 0.00315 -0.0718 0.7647 0.7995
0.750 0.4085 0.00945 0.00309 -0.0712 0.7517 0.8107
1.000 0.4348 0.00943 0.00306 -0.0707 0.7395 0.8229
1.250 0.4610 0.00942 0.00307 -0.0703 0.7286 0.8363
1.500 0.4871 0.00941 0.00307 -0.0697 0.7176 0.8511
1.750 0.5133 0.00939 0.00307 -0.0692 0.7056 0.8684
2.000 0.5403 0.00936 0.00307 -0.0688 0.6921 0.8910
2.250 0.5717 0.00934 0.00310 -0.0693 0.6768 0.9231
2.500 0.6090 0.00936 0.00310 -0.0713 0.6593 1.0000
2.750 0.6356 0.00948 0.00316 -0.0711 0.6424 1.0000
3.000 0.6619 0.00962 0.00324 -0.0708 0.6227 1.0000
3.250 0.6876 0.00979 0.00336 -0.0704 0.5993 1.0000
3.500 0.7128 0.00999 0.00346 -0.0699 0.5706 1.0000
3.750 0.7374 0.01023 0.00359 -0.0693 0.5379 1.0000
4.000 0.7611 0.01055 0.00376 -0.0685 0.5019 1.0000
4.250 0.7843 0.01092 0.00398 -0.0677 0.4635 1.0000
4.500 0.8070 0.01134 0.00428 -0.0669 0.4256 1.0000
4.750 0.8294 0.01181 0.00459 -0.0660 0.3879 1.0000
5.250 0.8734 0.01284 0.00532 -0.0644 0.3147 1.0000
5.500 0.8955 0.01336 0.00573 -0.0636 0.2855 1.0000
5.750 0.9181 0.01384 0.00619 -0.0629 0.2630 1.0000
6.000 0.9404 0.01435 0.00665 -0.0621 0.2401 1.0000
6.250 0.9614 0.01496 0.00711 -0.0612 0.2046 1.0000
6.500 0.9825 0.01558 0.00760 -0.0604 0.1753 1.0000
6.750 1.0037 0.01618 0.00812 -0.0595 0.1503 1.0000
7.000 1.0238 0.01689 0.00870 -0.0586 0.1220 1.0000
7.250 1.0433 0.01766 0.00937 -0.0576 0.0944 1.0000
7.500 1.0609 0.01862 0.01013 -0.0563 0.0608 1.0000
7.750 1.0727 0.02017 0.01131 -0.0543 0.0163 1.0000
8.000 1.0902 0.02115 0.01231 -0.0529 0.0111 1.0000
8.250 1.1070 0.02217 0.01345 -0.0513 0.0091 1.0000
8.500 1.1241 0.02311 0.01456 -0.0498 0.0083 1.0000
8.750 1.1407 0.02403 0.01561 -0.0483 0.0071 1.0000
9.000 1.1558 0.02502 0.01674 -0.0467 0.0065 1.0000
9.250 1.1673 0.02615 0.01801 -0.0446 0.0059 1.0000
9.500 1.1746 0.02754 0.01956 -0.0419 0.0055 1.0000
9.750 1.1796 0.02907 0.02131 -0.0392 0.0053 1.0000
10.000 1.1845 0.03061 0.02300 -0.0367 0.0052 1.0000
10.250 1.1891 0.03225 0.02480 -0.0343 0.0051 1.0000
10.500 1.1934 0.03397 0.02668 -0.0322 0.0050 1.0000
10.750 1.1969 0.03585 0.02871 -0.0302 0.0050 1.0000
11.000 1.1992 0.03796 0.03098 -0.0284 0.0049 1.0000
11.250 1.2010 0.04021 0.03340 -0.0268 0.0048 1.0000
11.500 1.2021 0.04265 0.03601 -0.0255 0.0048 1.0000
11.750 1.2020 0.04532 0.03887 -0.0243 0.0047 1.0000
12.000 1.2014 0.04813 0.04186 -0.0234 0.0047 1.0000
12.250 1.1990 0.05128 0.04520 -0.0228 0.0047 1.0000
12.500 1.1950 0.05473 0.04886 -0.0225 0.0046 1.0000
12.750 1.1896 0.05851 0.05284 -0.0225 0.0046 1.0000
13.000 1.1825 0.06266 0.05721 -0.0231 0.0046 1.0000
13.250 1.1748 0.06707 0.06182 -0.0241 0.0046 1.0000
13.500 1.1640 0.07221 0.06718 -0.0258 0.0046 1.0000
13.750 1.1523 0.07780 0.07297 -0.0280 0.0046 1.0000
14.000 1.1397 0.08396 0.07933 -0.0309 0.0046 1.0000
14.250 1.1270 0.09050 0.08607 -0.0344 0.0046 1.0000
14.500 1.1121 0.09799 0.09376 -0.0387 0.0046 1.0000
14.750 1.0982 0.10569 0.10163 -0.0432 0.0047 1.0000
15.000 1.0842 0.11377 0.10987 -0.0481 0.0047 1.0000
15.250 1.0693 0.12251 0.11876 -0.0535 0.0048 1.0000
15.500 1.0541 0.13175 0.12814 -0.0591 0.0048 1.0000
15.750 1.0397 0.14108 0.13760 -0.0648 0.0048 1.0000
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