HQ 3.0/12 AIRFOIL (hq3012-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: HQ 3.0/12 AIRFOIL (hq3012-il) Reynolds number: 500,000 Max Cl/Cd: 109.37 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq3012-il-500000.txt Download as CSV file: xf-hq3012-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 3.0/12 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5020 0.04640 0.04398 -0.0819 0.9857 0.0156
-9.000 -0.5095 0.03488 0.03177 -0.0906 0.9753 0.0138
-8.750 -0.4999 0.02573 0.02164 -0.0955 0.9698 0.0133
-8.500 -0.4802 0.02270 0.01815 -0.0960 0.9618 0.0136
-8.250 -0.4532 0.02046 0.01557 -0.0971 0.9571 0.0140
-8.000 -0.4291 0.01897 0.01382 -0.0970 0.9500 0.0144
-7.750 -0.4033 0.01777 0.01241 -0.0970 0.9436 0.0147
-7.500 -0.3805 0.01629 0.01072 -0.0965 0.9363 0.0152
-7.250 -0.3575 0.01513 0.00945 -0.0960 0.9291 0.0160
-7.000 -0.3321 0.01467 0.00894 -0.0957 0.9224 0.0171
-6.750 -0.3071 0.01415 0.00835 -0.0952 0.9152 0.0183
-6.500 -0.2817 0.01362 0.00769 -0.0948 0.9088 0.0194
-6.250 -0.2589 0.01270 0.00672 -0.0940 0.9012 0.0211
-6.000 -0.2328 0.01233 0.00629 -0.0937 0.8951 0.0235
-5.750 -0.2082 0.01178 0.00568 -0.0932 0.8877 0.0274
-5.250 -0.1551 0.01129 0.00517 -0.0927 0.8749 0.0406
-5.000 -0.1287 0.01096 0.00477 -0.0924 0.8690 0.0460
-4.750 -0.1020 0.01076 0.00454 -0.0923 0.8625 0.0507
-4.500 -0.0747 0.01065 0.00437 -0.0921 0.8563 0.0541
-4.250 -0.0484 0.01026 0.00397 -0.0919 0.8505 0.0617
-4.000 -0.0213 0.01011 0.00377 -0.0918 0.8439 0.0676
-3.750 0.0056 0.00977 0.00342 -0.0916 0.8384 0.0774
-3.500 0.0323 0.00948 0.00316 -0.0914 0.8318 0.0916
-3.250 0.0589 0.00910 0.00289 -0.0913 0.8258 0.1284
-3.000 0.0843 0.00848 0.00265 -0.0913 0.8197 0.2392
-2.750 0.1084 0.00766 0.00243 -0.0911 0.8128 0.4137
-2.500 0.1346 0.00738 0.00238 -0.0908 0.8066 0.5122
-2.250 0.1616 0.00728 0.00236 -0.0906 0.7994 0.5566
-2.000 0.1895 0.00726 0.00233 -0.0904 0.7937 0.5866
-1.750 0.2169 0.00724 0.00234 -0.0902 0.7869 0.6082
-1.500 0.2449 0.00726 0.00232 -0.0900 0.7811 0.6281
-1.250 0.2727 0.00728 0.00232 -0.0899 0.7751 0.6439
-1.000 0.3005 0.00728 0.00232 -0.0898 0.7686 0.6570
-0.750 0.3278 0.00732 0.00235 -0.0894 0.7618 0.6768
-0.500 0.3549 0.00733 0.00236 -0.0891 0.7533 0.6915
-0.250 0.3825 0.00735 0.00235 -0.0889 0.7451 0.7006
0.000 0.4100 0.00736 0.00235 -0.0887 0.7376 0.7104
0.250 0.4376 0.00739 0.00237 -0.0886 0.7307 0.7211
0.500 0.4652 0.00740 0.00237 -0.0884 0.7234 0.7291
0.750 0.4930 0.00742 0.00238 -0.0883 0.7160 0.7360
1.000 0.5205 0.00743 0.00238 -0.0882 0.7081 0.7429
1.250 0.5480 0.00745 0.00239 -0.0880 0.6990 0.7505
1.500 0.5750 0.00747 0.00240 -0.0878 0.6893 0.7578
1.750 0.6022 0.00749 0.00242 -0.0876 0.6789 0.7656
2.000 0.6292 0.00751 0.00246 -0.0873 0.6691 0.7735
2.250 0.6562 0.00756 0.00249 -0.0871 0.6595 0.7818
2.500 0.6829 0.00760 0.00254 -0.0868 0.6487 0.7908
2.750 0.7094 0.00762 0.00260 -0.0865 0.6367 0.7997
3.000 0.7357 0.00767 0.00265 -0.0861 0.6239 0.8099
3.250 0.7615 0.00772 0.00272 -0.0856 0.6093 0.8208
3.500 0.7867 0.00778 0.00279 -0.0851 0.5926 0.8327
4.000 0.8351 0.00793 0.00296 -0.0835 0.5531 0.8642
4.250 0.8569 0.00802 0.00307 -0.0822 0.5292 0.8917
4.500 0.8903 0.00814 0.00320 -0.0835 0.4976 1.0000
4.750 0.9143 0.00846 0.00338 -0.0829 0.4682 1.0000
5.000 0.9373 0.00883 0.00362 -0.0822 0.4371 1.0000
5.250 0.9600 0.00923 0.00387 -0.0814 0.4069 1.0000
5.500 0.9824 0.00962 0.00414 -0.0806 0.3783 1.0000
5.750 1.0045 0.01004 0.00443 -0.0797 0.3489 1.0000
6.000 1.0261 0.01049 0.00475 -0.0788 0.3211 1.0000
6.500 1.0689 0.01137 0.00541 -0.0769 0.2741 1.0000
6.750 1.0901 0.01182 0.00576 -0.0759 0.2538 1.0000
7.000 1.1114 0.01224 0.00613 -0.0750 0.2359 1.0000
7.250 1.1328 0.01264 0.00648 -0.0740 0.2188 1.0000
7.500 1.1536 0.01307 0.00685 -0.0730 0.2028 1.0000
7.750 1.1732 0.01355 0.00724 -0.0718 0.1850 1.0000
8.000 1.1924 0.01403 0.00766 -0.0705 0.1694 1.0000
8.250 1.2111 0.01452 0.00810 -0.0692 0.1549 1.0000
8.500 1.2294 0.01499 0.00853 -0.0678 0.1417 1.0000
8.750 1.2468 0.01543 0.00896 -0.0661 0.1312 1.0000
9.000 1.2627 0.01593 0.00943 -0.0643 0.1200 1.0000
9.250 1.2777 0.01649 0.00994 -0.0624 0.1081 1.0000
9.500 1.2920 0.01709 0.01052 -0.0605 0.0959 1.0000
9.750 1.3056 0.01776 0.01113 -0.0585 0.0843 1.0000
10.000 1.3182 0.01849 0.01182 -0.0564 0.0721 1.0000
10.500 1.3425 0.02007 0.01333 -0.0525 0.0538 1.0000
10.750 1.3548 0.02088 0.01415 -0.0507 0.0476 1.0000
11.250 1.3779 0.02264 0.01595 -0.0471 0.0370 1.0000
11.500 1.3875 0.02371 0.01700 -0.0453 0.0318 1.0000
11.750 1.3990 0.02466 0.01799 -0.0437 0.0275 1.0000
12.000 1.4068 0.02591 0.01924 -0.0419 0.0213 1.0000
12.250 1.4130 0.02733 0.02066 -0.0400 0.0160 1.0000
12.500 1.4177 0.02892 0.02227 -0.0382 0.0119 1.0000
12.750 1.4217 0.03064 0.02402 -0.0365 0.0097 1.0000
13.000 1.4268 0.03231 0.02578 -0.0350 0.0085 1.0000
13.250 1.4308 0.03414 0.02770 -0.0336 0.0078 1.0000
13.500 1.4335 0.03615 0.02981 -0.0323 0.0074 1.0000
13.750 1.4346 0.03839 0.03216 -0.0312 0.0071 1.0000
14.000 1.4363 0.04064 0.03454 -0.0302 0.0068 1.0000
14.250 1.4373 0.04304 0.03706 -0.0294 0.0067 1.0000
14.500 1.4366 0.04573 0.03987 -0.0289 0.0066 1.0000
14.750 1.4346 0.04868 0.04295 -0.0285 0.0064 1.0000
15.000 1.4320 0.05183 0.04622 -0.0284 0.0063 1.0000
15.250 1.4280 0.05531 0.04983 -0.0285 0.0062 1.0000
15.500 1.4220 0.05920 0.05385 -0.0290 0.0061 1.0000
15.750 1.4159 0.06329 0.05807 -0.0298 0.0060 1.0000
16.000 1.4082 0.06781 0.06272 -0.0309 0.0059 1.0000
16.250 1.3991 0.07276 0.06781 -0.0325 0.0058 1.0000
16.500 1.3901 0.07789 0.07307 -0.0343 0.0057 1.0000
16.750 1.3782 0.08372 0.07905 -0.0366 0.0056 1.0000
17.000 1.3653 0.08996 0.08543 -0.0392 0.0057 1.0000
17.250 1.3529 0.09627 0.09188 -0.0421 0.0056 1.0000
17.500 1.3391 0.10302 0.09876 -0.0453 0.0056 1.0000
17.750 1.3257 0.10985 0.10573 -0.0487 0.0056 1.0000
|
Polar data table (+)
Polar graphs
<< Back to HQ 3.0/12 AIRFOIL (hq3012-il)