HQ 3.0/8 AIRFOIL (hq308-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 3.0/8 AIRFOIL (hq308-il) Reynolds number: 1,000,000 Max Cl/Cd: 135.39 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq308-il-1000000.txt Download as CSV file: xf-hq308-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.0/8 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3668 0.08788 0.08631 -0.0322 1.0000 0.0056
-8.000 -0.3682 0.08500 0.08346 -0.0323 1.0000 0.0056
-7.750 -0.3744 0.08266 0.08117 -0.0312 1.0000 0.0056
-7.500 -0.3717 0.07917 0.07770 -0.0334 0.9986 0.0056
-7.250 -0.3515 0.07357 0.07211 -0.0417 0.9939 0.0056
-7.000 -0.3288 0.06762 0.06615 -0.0513 0.9887 0.0057
-6.750 -0.2989 0.06020 0.05869 -0.0646 0.9845 0.0057
-6.500 -0.2669 0.05305 0.05142 -0.0756 0.9821 0.0057
-6.250 -0.2403 0.04707 0.04530 -0.0826 0.9755 0.0057
-6.000 -0.2144 0.03644 0.03434 -0.0920 0.9686 0.0060
-5.750 -0.1945 0.03191 0.02957 -0.0940 0.9583 0.0063
-5.500 -0.1723 0.02901 0.02646 -0.0949 0.9499 0.0066
-5.250 -0.1494 0.02633 0.02357 -0.0954 0.9422 0.0070
-5.000 -0.1264 0.02381 0.02082 -0.0953 0.9336 0.0079
-4.750 -0.1010 0.02139 0.01815 -0.0949 0.9265 0.0092
-3.500 0.0288 0.01062 0.00587 -0.0928 0.8920 0.0117
-3.250 0.0561 0.00985 0.00499 -0.0923 0.8857 0.0099
-3.000 0.0825 0.00862 0.00360 -0.0918 0.8786 0.0093
-2.750 0.1094 0.00789 0.00272 -0.0915 0.8714 0.0091
-2.500 0.1366 0.00740 0.00210 -0.0912 0.8629 0.0094
-2.250 0.1642 0.00711 0.00171 -0.0910 0.8544 0.0104
-2.000 0.1920 0.00682 0.00130 -0.0908 0.8472 0.0160
-1.750 0.2197 0.00653 0.00115 -0.0908 0.8396 0.0602
-1.500 0.2469 0.00615 0.00103 -0.0908 0.8316 0.1611
-1.250 0.2729 0.00536 0.00091 -0.0910 0.8216 0.4114
-1.000 0.2992 0.00493 0.00091 -0.0909 0.8110 0.5750
-0.750 0.3264 0.00484 0.00090 -0.0907 0.8018 0.6289
-0.500 0.3538 0.00481 0.00088 -0.0906 0.7931 0.6571
-0.250 0.3812 0.00478 0.00089 -0.0904 0.7835 0.6863
0.000 0.4082 0.00475 0.00091 -0.0901 0.7726 0.7237
0.250 0.4345 0.00472 0.00094 -0.0897 0.7588 0.7636
0.500 0.4614 0.00473 0.00094 -0.0894 0.7448 0.7816
0.750 0.4886 0.00475 0.00095 -0.0892 0.7324 0.7948
1.000 0.5156 0.00477 0.00097 -0.0890 0.7194 0.8092
1.250 0.5424 0.00479 0.00099 -0.0887 0.7060 0.8262
1.500 0.5687 0.00481 0.00102 -0.0884 0.6903 0.8473
1.750 0.5936 0.00479 0.00106 -0.0876 0.6734 0.8816
2.000 0.6248 0.00470 0.00107 -0.0883 0.6542 1.0000
2.250 0.6516 0.00484 0.00112 -0.0881 0.6324 1.0000
2.500 0.6783 0.00501 0.00119 -0.0879 0.6062 1.0000
2.750 0.7044 0.00522 0.00130 -0.0876 0.5747 1.0000
3.000 0.7300 0.00548 0.00141 -0.0873 0.5371 1.0000
3.250 0.7541 0.00590 0.00158 -0.0867 0.4760 1.0000
3.500 0.7776 0.00641 0.00179 -0.0860 0.4101 1.0000
3.750 0.8024 0.00680 0.00198 -0.0856 0.3645 1.0000
4.000 0.8264 0.00728 0.00223 -0.0851 0.3098 1.0000
4.250 0.8501 0.00779 0.00249 -0.0846 0.2566 1.0000
4.500 0.8738 0.00831 0.00276 -0.0841 0.2050 1.0000
4.750 0.8982 0.00874 0.00302 -0.0836 0.1709 1.0000
5.000 0.9227 0.00915 0.00330 -0.0832 0.1396 1.0000
5.250 0.9463 0.00967 0.00364 -0.0827 0.1022 1.0000
5.500 0.9680 0.01043 0.00410 -0.0819 0.0526 1.0000
5.750 0.9904 0.01109 0.00458 -0.0811 0.0220 1.0000
6.000 1.0134 0.01172 0.00508 -0.0804 0.0060 1.0000
6.250 1.0378 0.01217 0.00560 -0.0797 0.0039 1.0000
6.500 1.0621 0.01262 0.00614 -0.0791 0.0037 1.0000
6.750 1.0858 0.01314 0.00675 -0.0784 0.0035 1.0000
7.000 1.1089 0.01374 0.00748 -0.0776 0.0035 1.0000
7.250 1.1309 0.01446 0.00832 -0.0766 0.0035 1.0000
7.500 1.1515 0.01535 0.00935 -0.0753 0.0035 1.0000
7.750 1.1708 0.01637 0.01051 -0.0739 0.0036 1.0000
8.000 1.1878 0.01767 0.01196 -0.0721 0.0037 1.0000
8.250 1.2017 0.01947 0.01395 -0.0698 0.0038 1.0000
8.500 1.2158 0.02144 0.01607 -0.0677 0.0040 1.0000
8.750 1.2336 0.02278 0.01753 -0.0662 0.0041 1.0000
9.000 1.2481 0.02490 0.01984 -0.0643 0.0041 1.0000
9.250 1.2640 0.02660 0.02170 -0.0627 0.0041 1.0000
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Polar data table (+)
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