HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.5/9 AIRFOIL (hq259-il) Reynolds number: 500,000 Max Cl/Cd: 106.8 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq259-il-500000.txt Download as CSV file: xf-hq259-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3970 0.08519 0.08301 -0.0351 1.0000 0.0152
-8.250 -0.3961 0.08187 0.07972 -0.0365 1.0000 0.0153
-8.000 -0.3978 0.07844 0.07633 -0.0379 1.0000 0.0154
-7.750 -0.4054 0.07526 0.07322 -0.0383 1.0000 0.0155
-7.500 -0.4209 0.07276 0.07079 -0.0372 1.0000 0.0154
-7.250 -0.4294 0.06844 0.06649 -0.0409 0.9990 0.0155
-6.750 -0.3352 0.03308 0.03076 -0.0687 0.9828 0.0162
-6.500 -0.3218 0.02634 0.02379 -0.0735 0.9773 0.0169
-6.250 -0.2981 0.02278 0.02010 -0.0766 0.9732 0.0178
-6.000 -0.2718 0.01916 0.01627 -0.0795 0.9698 0.0190
-5.750 -0.2491 0.01589 0.01273 -0.0808 0.9631 0.0208
-5.500 -0.2460 0.02205 0.01780 -0.0820 0.9679 0.0133
-5.250 -0.2198 0.01803 0.01322 -0.0819 0.9610 0.0127
-5.000 -0.1879 0.01636 0.01128 -0.0828 0.9571 0.0137
-4.750 -0.1591 0.01506 0.00976 -0.0829 0.9516 0.0146
-4.500 -0.1319 0.01321 0.00765 -0.0826 0.9453 0.0143
-4.250 -0.1046 0.01190 0.00616 -0.0823 0.9396 0.0145
-4.000 -0.0790 0.01101 0.00516 -0.0816 0.9321 0.0152
-3.750 -0.0531 0.01000 0.00399 -0.0811 0.9257 0.0174
-3.500 -0.0268 0.00952 0.00345 -0.0807 0.9183 0.0232
-3.250 0.0000 0.00905 0.00302 -0.0803 0.9117 0.0472
-3.000 0.0267 0.00882 0.00276 -0.0800 0.9043 0.0623
-2.750 0.0534 0.00849 0.00250 -0.0798 0.8974 0.0916
-2.500 0.0781 0.00766 0.00223 -0.0797 0.8895 0.2509
-2.250 0.1018 0.00675 0.00207 -0.0794 0.8812 0.4894
-2.000 0.1277 0.00650 0.00202 -0.0789 0.8738 0.5777
-1.750 0.1541 0.00641 0.00197 -0.0784 0.8652 0.6236
-1.500 0.1810 0.00636 0.00193 -0.0781 0.8582 0.6571
-1.250 0.2078 0.00632 0.00192 -0.0776 0.8502 0.6861
-1.000 0.2347 0.00630 0.00190 -0.0773 0.8427 0.7104
-0.750 0.2612 0.00629 0.00188 -0.0767 0.8340 0.7359
-0.500 0.2867 0.00625 0.00190 -0.0759 0.8237 0.7661
-0.250 0.3128 0.00623 0.00187 -0.0753 0.8132 0.7850
0.000 0.3395 0.00621 0.00183 -0.0749 0.8031 0.7987
0.250 0.3664 0.00618 0.00179 -0.0746 0.7932 0.8102
0.500 0.3932 0.00616 0.00178 -0.0742 0.7833 0.8226
0.750 0.4199 0.00614 0.00176 -0.0739 0.7733 0.8360
1.000 0.4462 0.00611 0.00173 -0.0734 0.7625 0.8511
1.250 0.4719 0.00608 0.00172 -0.0728 0.7509 0.8693
1.500 0.4970 0.00601 0.00172 -0.0720 0.7387 0.8943
1.750 0.5258 0.00592 0.00171 -0.0719 0.7261 0.9412
2.000 0.5638 0.00594 0.00170 -0.0742 0.7117 1.0000
2.250 0.5910 0.00604 0.00172 -0.0741 0.6957 1.0000
2.500 0.6180 0.00614 0.00179 -0.0740 0.6773 1.0000
2.750 0.6448 0.00626 0.00185 -0.0738 0.6587 1.0000
3.000 0.6713 0.00640 0.00192 -0.0735 0.6374 1.0000
3.250 0.6974 0.00657 0.00201 -0.0732 0.6121 1.0000
3.500 0.7230 0.00677 0.00214 -0.0727 0.5817 1.0000
3.750 0.7477 0.00705 0.00228 -0.0722 0.5424 1.0000
4.000 0.7715 0.00742 0.00246 -0.0715 0.4945 1.0000
4.250 0.7948 0.00786 0.00269 -0.0708 0.4439 1.0000
4.500 0.8182 0.00832 0.00295 -0.0701 0.3969 1.0000
4.750 0.8418 0.00878 0.00326 -0.0695 0.3512 1.0000
5.000 0.8639 0.00940 0.00360 -0.0687 0.2936 1.0000
5.250 0.8859 0.01004 0.00396 -0.0680 0.2412 1.0000
5.500 0.9092 0.01056 0.00430 -0.0674 0.2024 1.0000
5.750 0.9328 0.01104 0.00464 -0.0669 0.1702 1.0000
6.000 0.9555 0.01161 0.00504 -0.0662 0.1338 1.0000
6.250 0.9755 0.01250 0.00559 -0.0652 0.0789 1.0000
6.500 0.9903 0.01407 0.00670 -0.0633 0.0117 1.0000
6.750 1.0135 0.01462 0.00733 -0.0625 0.0092 1.0000
7.000 1.0359 0.01525 0.00806 -0.0617 0.0080 1.0000
7.250 1.0567 0.01610 0.00903 -0.0605 0.0070 1.0000
7.500 1.0721 0.01757 0.01071 -0.0584 0.0063 1.0000
7.750 1.0899 0.01868 0.01193 -0.0568 0.0061 1.0000
8.000 1.1064 0.01990 0.01328 -0.0551 0.0060 1.0000
8.250 1.1217 0.02132 0.01483 -0.0532 0.0060 1.0000
8.500 1.1357 0.02303 0.01669 -0.0511 0.0060 1.0000
8.750 1.1494 0.02514 0.01901 -0.0491 0.0061 1.0000
9.000 1.1634 0.02805 0.02211 -0.0473 0.0063 1.0000
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Polar data table (+)
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