HQ 0/7 AIRFOIL (hq07-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 0/7 AIRFOIL (hq07-il) Reynolds number: 50,000 Max Cl/Cd: 26.16 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq07-il-50000.txt Download as CSV file: xf-hq07-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 0/7 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6421 0.11141 0.10489 0.0290 1.0000 0.2823
-8.750 -0.6432 0.10807 0.10161 0.0286 1.0000 0.2967
-8.500 -0.6473 0.10503 0.09865 0.0281 1.0000 0.3110
-8.000 -0.6213 0.09582 0.08935 0.0297 1.0000 0.3425
-7.750 -0.6160 0.09200 0.08557 0.0301 1.0000 0.3592
-7.500 -0.6255 0.08954 0.08322 0.0304 1.0000 0.3787
-7.250 -0.6047 0.08483 0.07848 0.0318 1.0000 0.3997
-6.500 -0.6683 0.05487 0.04763 -0.0125 1.0000 0.1357
-6.250 -0.6523 0.04888 0.04112 -0.0132 1.0000 0.1200
-6.000 -0.6342 0.04372 0.03510 -0.0131 1.0000 0.1093
-5.750 -0.6133 0.03926 0.03016 -0.0125 1.0000 0.1060
-5.500 -0.5902 0.03538 0.02566 -0.0116 1.0000 0.1059
-5.250 -0.5673 0.03245 0.02224 -0.0108 1.0000 0.1177
-5.000 -0.5401 0.02946 0.01866 -0.0097 1.0000 0.1225
-4.750 -0.5119 0.02665 0.01566 -0.0087 1.0000 0.1302
-4.500 -0.4838 0.02440 0.01332 -0.0076 1.0000 0.1465
-4.250 -0.4589 0.02254 0.01164 -0.0065 1.0000 0.1828
-4.000 -0.4360 0.01973 0.00948 -0.0049 1.0000 0.2645
-3.750 -0.4456 0.01703 0.00972 0.0063 1.0000 0.7139
-3.500 -0.3080 0.02020 0.01150 0.0028 1.0000 0.9528
-3.250 -0.2005 0.01872 0.00911 -0.0138 1.0000 1.0000
-3.000 -0.1846 0.01804 0.00825 -0.0136 1.0000 1.0000
-2.750 -0.1681 0.01746 0.00750 -0.0134 1.0000 1.0000
-2.500 -0.1513 0.01695 0.00677 -0.0130 1.0000 1.0000
-2.250 -0.1343 0.01651 0.00620 -0.0124 1.0000 1.0000
-2.000 -0.1173 0.01612 0.00571 -0.0117 1.0000 1.0000
-1.750 -0.1004 0.01578 0.00529 -0.0109 1.0000 1.0000
-1.500 -0.0839 0.01549 0.00494 -0.0100 1.0000 1.0000
-1.250 -0.0678 0.01523 0.00464 -0.0088 1.0000 1.0000
-1.000 -0.0523 0.01501 0.00440 -0.0075 1.0000 1.0000
-0.750 -0.0377 0.01484 0.00419 -0.0060 1.0000 1.0000
-0.500 -0.0242 0.01471 0.00406 -0.0042 1.0000 1.0000
-0.250 -0.0117 0.01463 0.00398 -0.0022 1.0000 1.0000
0.000 0.0000 0.01460 0.00395 0.0000 1.0000 1.0000
0.250 0.0117 0.01463 0.00398 0.0022 1.0000 1.0000
0.500 0.0242 0.01471 0.00406 0.0042 1.0000 1.0000
0.750 0.0377 0.01484 0.00420 0.0060 1.0000 1.0000
1.000 0.0523 0.01502 0.00440 0.0075 1.0000 1.0000
1.250 0.0677 0.01523 0.00464 0.0088 1.0000 1.0000
1.500 0.0838 0.01549 0.00494 0.0100 1.0000 1.0000
1.750 0.1004 0.01578 0.00529 0.0109 1.0000 1.0000
2.000 0.1173 0.01612 0.00571 0.0117 1.0000 1.0000
2.250 0.1343 0.01651 0.00620 0.0124 1.0000 1.0000
2.500 0.1513 0.01695 0.00676 0.0130 1.0000 1.0000
2.750 0.1682 0.01746 0.00750 0.0134 1.0000 1.0000
3.000 0.1847 0.01804 0.00824 0.0136 1.0000 1.0000
3.250 0.2007 0.01871 0.00910 0.0137 1.0000 1.0000
3.500 0.3080 0.02020 0.01150 -0.0028 0.9529 1.0000
3.750 0.4455 0.01703 0.00972 -0.0063 0.7136 1.0000
4.000 0.4360 0.01972 0.00947 0.0049 0.2647 1.0000
4.250 0.4589 0.02254 0.01164 0.0065 0.1829 1.0000
4.500 0.4838 0.02440 0.01333 0.0076 0.1465 1.0000
4.750 0.5119 0.02664 0.01565 0.0087 0.1303 1.0000
5.000 0.5401 0.02946 0.01866 0.0097 0.1225 1.0000
5.250 0.5673 0.03244 0.02223 0.0107 0.1176 1.0000
5.500 0.5902 0.03539 0.02567 0.0116 0.1058 1.0000
5.750 0.6133 0.03925 0.03015 0.0125 0.1061 1.0000
6.000 0.6343 0.04370 0.03508 0.0131 0.1093 1.0000
6.250 0.6523 0.04889 0.04113 0.0132 0.1200 1.0000
6.500 0.6684 0.05487 0.04762 0.0125 0.1357 1.0000
7.250 0.6060 0.08496 0.07861 -0.0319 0.4008 1.0000
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Polar data table (+)
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