HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 2.5/9 AIRFOIL (hq259-il) Reynolds number: 200,000 Max Cl/Cd: 73.01 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq259-il-200000-n5.txt Download as CSV file: xf-hq259-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/9 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4026 0.08810 0.08462 -0.0338 1.0000 0.0099
-8.500 -0.4037 0.08465 0.08121 -0.0348 1.0000 0.0095
-8.250 -0.4070 0.08072 0.07734 -0.0362 1.0000 0.0094
-8.000 -0.4119 0.07720 0.07390 -0.0371 1.0000 0.0093
-7.750 -0.4214 0.07388 0.07066 -0.0375 1.0000 0.0092
-7.500 -0.4316 0.07053 0.06740 -0.0385 1.0000 0.0090
-7.250 -0.4424 0.06625 0.06317 -0.0412 1.0000 0.0090
-7.000 -0.4324 0.05989 0.05671 -0.0484 0.9949 0.0088
-6.750 -0.4109 0.05260 0.04921 -0.0569 0.9872 0.0085
-6.500 -0.3884 0.04536 0.04166 -0.0634 0.9797 0.0084
-6.250 -0.3620 0.03837 0.03423 -0.0684 0.9741 0.0083
-6.000 -0.3369 0.03238 0.02771 -0.0711 0.9674 0.0083
-5.750 -0.3081 0.02705 0.02173 -0.0732 0.9624 0.0086
-5.500 -0.2759 0.02439 0.01853 -0.0746 0.9585 0.0100
-5.250 -0.2501 0.02109 0.01470 -0.0752 0.9519 0.0113
-5.000 -0.2186 0.01901 0.01226 -0.0762 0.9478 0.0120
-4.750 -0.1892 0.01743 0.01043 -0.0765 0.9426 0.0127
-4.500 -0.1599 0.01613 0.00889 -0.0768 0.9370 0.0140
-4.250 -0.1282 0.01502 0.00762 -0.0776 0.9329 0.0158
-4.000 -0.1012 0.01423 0.00674 -0.0775 0.9258 0.0195
-3.750 -0.0709 0.01345 0.00586 -0.0780 0.9208 0.0235
-3.500 -0.0426 0.01278 0.00510 -0.0781 0.9144 0.0328
-3.250 -0.0133 0.01235 0.00465 -0.0784 0.9085 0.0494
-3.000 0.0153 0.01197 0.00434 -0.0788 0.9026 0.0772
-2.750 0.0432 0.01156 0.00394 -0.0789 0.8957 0.1075
-2.500 0.0709 0.01097 0.00361 -0.0791 0.8897 0.1912
-2.250 0.0958 0.01011 0.00338 -0.0791 0.8822 0.3774
-2.000 0.1210 0.00961 0.00335 -0.0787 0.8756 0.5322
-1.750 0.1463 0.00942 0.00337 -0.0779 0.8681 0.6190
-1.500 0.1725 0.00934 0.00332 -0.0772 0.8607 0.6657
-1.250 0.1975 0.00926 0.00333 -0.0761 0.8530 0.7175
-1.000 0.2216 0.00921 0.00335 -0.0749 0.8442 0.7582
-0.750 0.2484 0.00916 0.00325 -0.0744 0.8367 0.7755
-0.500 0.2749 0.00912 0.00316 -0.0739 0.8276 0.7875
-0.250 0.3017 0.00909 0.00310 -0.0735 0.8194 0.7996
0.000 0.3286 0.00905 0.00303 -0.0731 0.8115 0.8124
0.250 0.3548 0.00902 0.00299 -0.0726 0.8014 0.8262
0.500 0.3808 0.00896 0.00291 -0.0720 0.7895 0.8414
0.750 0.4066 0.00890 0.00283 -0.0712 0.7756 0.8588
1.000 0.4330 0.00884 0.00277 -0.0706 0.7622 0.8799
1.250 0.4620 0.00878 0.00273 -0.0706 0.7501 0.9088
1.500 0.4992 0.00872 0.00269 -0.0725 0.7372 0.9620
1.750 0.5294 0.00877 0.00269 -0.0730 0.7232 1.0000
2.000 0.5563 0.00886 0.00271 -0.0728 0.7075 1.0000
2.250 0.5831 0.00896 0.00275 -0.0726 0.6903 1.0000
2.500 0.6097 0.00908 0.00284 -0.0723 0.6720 1.0000
2.750 0.6360 0.00921 0.00291 -0.0719 0.6517 1.0000
3.000 0.6620 0.00937 0.00301 -0.0715 0.6285 1.0000
3.250 0.6875 0.00956 0.00311 -0.0710 0.5995 1.0000
3.500 0.7123 0.00980 0.00323 -0.0704 0.5660 1.0000
3.750 0.7367 0.01009 0.00343 -0.0697 0.5281 1.0000
4.000 0.7602 0.01046 0.00364 -0.0689 0.4866 1.0000
4.250 0.7831 0.01090 0.00390 -0.0681 0.4435 1.0000
4.500 0.8057 0.01138 0.00422 -0.0673 0.4002 1.0000
4.750 0.8282 0.01190 0.00457 -0.0665 0.3565 1.0000
5.000 0.8504 0.01245 0.00502 -0.0657 0.3148 1.0000
5.250 0.8731 0.01298 0.00544 -0.0650 0.2825 1.0000
5.500 0.8964 0.01347 0.00588 -0.0643 0.2578 1.0000
5.750 0.9168 0.01422 0.00638 -0.0634 0.2077 1.0000
6.000 0.9363 0.01511 0.00691 -0.0625 0.1491 1.0000
6.250 0.9581 0.01576 0.00746 -0.0617 0.1195 1.0000
6.500 0.9774 0.01673 0.00815 -0.0607 0.0734 1.0000
6.750 0.9925 0.01825 0.00929 -0.0592 0.0171 1.0000
7.000 1.0129 0.01916 0.01025 -0.0580 0.0105 1.0000
7.250 1.0336 0.02000 0.01124 -0.0569 0.0088 1.0000
7.500 1.0515 0.02118 0.01256 -0.0555 0.0066 1.0000
7.750 1.0712 0.02206 0.01360 -0.0543 0.0060 1.0000
8.000 1.0889 0.02315 0.01488 -0.0529 0.0056 1.0000
8.250 1.1051 0.02433 0.01624 -0.0513 0.0053 1.0000
8.500 1.1193 0.02568 0.01776 -0.0495 0.0051 1.0000
8.750 1.1322 0.02710 0.01935 -0.0476 0.0049 1.0000
9.000 1.1440 0.02863 0.02104 -0.0455 0.0048 1.0000
9.250 1.1540 0.03029 0.02286 -0.0433 0.0047 1.0000
9.500 1.1627 0.03200 0.02481 -0.0409 0.0046 1.0000
9.750 1.1702 0.03392 0.02692 -0.0385 0.0045 1.0000
10.000 1.1770 0.03593 0.02914 -0.0363 0.0045 1.0000
10.250 1.1821 0.03816 0.03159 -0.0340 0.0045 1.0000
10.500 1.1851 0.04056 0.03423 -0.0318 0.0045 1.0000
10.750 1.1856 0.04316 0.03709 -0.0297 0.0045 1.0000
11.000 1.1830 0.04603 0.04022 -0.0278 0.0044 1.0000
11.250 1.1789 0.04911 0.04356 -0.0262 0.0045 1.0000
11.500 1.1711 0.05263 0.04735 -0.0250 0.0045 1.0000
11.750 1.1611 0.05657 0.05155 -0.0245 0.0045 1.0000
12.000 1.1498 0.06085 0.05607 -0.0246 0.0045 1.0000
12.250 1.1357 0.06585 0.06131 -0.0256 0.0045 1.0000
12.500 1.1221 0.07113 0.06682 -0.0274 0.0046 1.0000
12.750 1.1050 0.07751 0.07342 -0.0304 0.0046 1.0000
13.000 1.0873 0.08469 0.08081 -0.0344 0.0046 1.0000
13.250 1.0704 0.09247 0.08877 -0.0392 0.0047 1.0000
13.500 1.0522 0.10138 0.09786 -0.0450 0.0047 1.0000
13.750 1.0342 0.11118 0.10782 -0.0514 0.0048 1.0000
14.000 1.0147 0.12211 0.11888 -0.0582 0.0048 1.0000
14.250 0.9942 0.13416 0.13103 -0.0653 0.0049 1.0000
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Polar data table (+)
Polar graphs
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