HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il) Reynolds number: 100,000 Max Cl/Cd: 47.79 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1010-il-100000.txt Download as CSV file: xf-hq1010-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5673 0.09756 0.09283 -0.0204 1.0000 0.1211
-9.250 -0.5383 0.09463 0.08985 -0.0151 1.0000 0.1292
-9.000 -0.5561 0.09000 0.08533 -0.0203 1.0000 0.1339
-8.750 -0.5579 0.08538 0.08078 -0.0215 1.0000 0.1382
-8.500 -0.5502 0.08220 0.07762 -0.0211 1.0000 0.1437
-8.250 -0.5887 0.07652 0.07204 -0.0290 1.0000 0.1479
-7.750 -0.6373 0.05186 0.04636 -0.0384 1.0000 0.0643
-7.500 -0.6343 0.04718 0.04137 -0.0375 1.0000 0.0622
-7.250 -0.6293 0.04266 0.03640 -0.0361 1.0000 0.0606
-7.000 -0.6212 0.03836 0.03157 -0.0343 1.0000 0.0591
-6.750 -0.6106 0.03422 0.02672 -0.0320 1.0000 0.0575
-6.500 -0.5950 0.03125 0.02334 -0.0302 1.0000 0.0579
-6.250 -0.5773 0.02885 0.02068 -0.0285 1.0000 0.0602
-6.000 -0.5593 0.02721 0.01866 -0.0267 1.0000 0.0659
-5.750 -0.5411 0.02488 0.01616 -0.0251 1.0000 0.0712
-5.500 -0.5219 0.02339 0.01446 -0.0233 1.0000 0.0792
-5.250 -0.5040 0.02186 0.01291 -0.0217 1.0000 0.0905
-5.000 -0.4860 0.02050 0.01156 -0.0200 1.0000 0.1034
-4.750 -0.4684 0.01936 0.01044 -0.0182 1.0000 0.1188
-4.500 -0.4514 0.01830 0.00949 -0.0166 1.0000 0.1379
-4.250 -0.4345 0.01731 0.00864 -0.0148 1.0000 0.1621
-4.000 -0.4187 0.01617 0.00791 -0.0132 1.0000 0.2109
-3.750 -0.4080 0.01412 0.00719 -0.0110 1.0000 0.4131
-3.500 -0.3971 0.01372 0.00753 -0.0073 1.0000 0.5947
-3.250 -0.3831 0.01387 0.00778 -0.0041 1.0000 0.6660
-3.000 -0.3689 0.01408 0.00797 -0.0010 1.0000 0.7143
-2.750 -0.3555 0.01427 0.00817 0.0023 1.0000 0.7503
-2.500 -0.3426 0.01442 0.00829 0.0055 1.0000 0.7840
-2.250 -0.3309 0.01452 0.00836 0.0091 1.0000 0.8155
-2.000 -0.3199 0.01457 0.00842 0.0128 1.0000 0.8477
-1.750 -0.3002 0.01471 0.00853 0.0150 0.9966 0.8843
-1.500 -0.2596 0.01505 0.00879 0.0132 0.9899 0.9180
-1.250 -0.2031 0.01546 0.00903 0.0076 0.9847 0.9429
-1.000 -0.1306 0.01595 0.00934 -0.0011 0.9823 0.9637
-0.750 -0.0545 0.01626 0.00948 -0.0108 0.9799 0.9804
-0.500 0.0122 0.01635 0.00947 -0.0193 0.9749 0.9915
-0.250 0.0758 0.01643 0.00946 -0.0275 0.9700 1.0000
0.000 0.1142 0.01641 0.00940 -0.0311 0.9590 1.0000
0.250 0.1527 0.01642 0.00938 -0.0345 0.9483 1.0000
0.500 0.2047 0.01642 0.00937 -0.0400 0.9409 1.0000
0.750 0.2505 0.01630 0.00926 -0.0441 0.9287 1.0000
1.000 0.3079 0.01596 0.00897 -0.0495 0.9157 1.0000
1.250 0.3547 0.01559 0.00863 -0.0528 0.9014 1.0000
1.500 0.3863 0.01541 0.00848 -0.0536 0.8873 1.0000
1.750 0.4114 0.01532 0.00843 -0.0531 0.8730 1.0000
2.000 0.4332 0.01526 0.00843 -0.0518 0.8584 1.0000
2.250 0.4533 0.01523 0.00842 -0.0501 0.8436 1.0000
2.500 0.4736 0.01519 0.00841 -0.0482 0.8285 1.0000
2.750 0.4925 0.01520 0.00844 -0.0461 0.8120 1.0000
3.000 0.5118 0.01520 0.00850 -0.0439 0.7941 1.0000
3.250 0.5337 0.01510 0.00843 -0.0419 0.7759 1.0000
3.500 0.5573 0.01489 0.00824 -0.0399 0.7572 1.0000
3.750 0.5783 0.01477 0.00816 -0.0377 0.7331 1.0000
4.000 0.6008 0.01459 0.00804 -0.0356 0.7072 1.0000
4.250 0.6233 0.01440 0.00786 -0.0335 0.6769 1.0000
4.500 0.6455 0.01427 0.00771 -0.0314 0.6399 1.0000
4.750 0.6668 0.01423 0.00761 -0.0293 0.5920 1.0000
5.000 0.6872 0.01438 0.00757 -0.0271 0.5319 1.0000
5.250 0.7063 0.01486 0.00777 -0.0250 0.4625 1.0000
5.500 0.7247 0.01561 0.00818 -0.0232 0.3965 1.0000
5.750 0.7428 0.01650 0.00875 -0.0216 0.3386 1.0000
6.000 0.7607 0.01755 0.00949 -0.0200 0.2859 1.0000
6.250 0.7779 0.01876 0.01039 -0.0185 0.2366 1.0000
6.500 0.7959 0.01998 0.01138 -0.0171 0.1927 1.0000
6.750 0.8145 0.02112 0.01242 -0.0157 0.1570 1.0000
7.000 0.8336 0.02244 0.01360 -0.0145 0.1304 1.0000
7.250 0.8534 0.02407 0.01509 -0.0134 0.1102 1.0000
7.500 0.8736 0.02563 0.01669 -0.0122 0.0924 1.0000
7.750 0.8954 0.02765 0.01875 -0.0112 0.0795 1.0000
8.000 0.9162 0.02971 0.02088 -0.0102 0.0684 1.0000
8.250 0.9366 0.03223 0.02341 -0.0095 0.0597 1.0000
8.500 0.9555 0.03435 0.02600 -0.0080 0.0536 1.0000
8.750 0.9742 0.03732 0.02906 -0.0071 0.0497 1.0000
9.000 0.9851 0.04080 0.03306 -0.0052 0.0470 1.0000
9.250 0.9930 0.04379 0.03661 -0.0030 0.0447 1.0000
9.500 0.9969 0.04745 0.04075 -0.0009 0.0436 1.0000
9.750 0.9949 0.05158 0.04533 0.0014 0.0437 1.0000
10.000 0.9861 0.05589 0.05004 0.0037 0.0439 1.0000
10.250 0.9726 0.06001 0.05448 0.0058 0.0444 1.0000
10.500 0.9532 0.06387 0.05857 0.0081 0.0448 1.0000
10.750 0.9325 0.06804 0.06293 0.0088 0.0453 1.0000
11.000 0.9113 0.07276 0.06780 0.0080 0.0458 1.0000
11.250 0.8878 0.07851 0.07368 0.0055 0.0464 1.0000
11.500 0.8677 0.08482 0.08008 0.0020 0.0470 1.0000
11.750 0.8525 0.09154 0.08683 -0.0016 0.0478 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 1.0/10 AIRFOIL (hq1010-il)