HQ 2.0/9 AIRFOIL (hq209-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 2.0/9 AIRFOIL (hq209-il) Reynolds number: 50,000 Max Cl/Cd: 37.86 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq209-il-50000-n5.txt Download as CSV file: xf-hq209-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.0/9 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4649 0.11367 0.10654 -0.0215 1.0000 0.0567
-9.750 -0.4587 0.10974 0.10265 -0.0221 1.0000 0.0548
-9.250 -0.4692 0.09952 0.09262 -0.0312 1.0000 0.0452
-9.000 -0.4640 0.09567 0.08881 -0.0314 1.0000 0.0447
-8.750 -0.4642 0.09140 0.08462 -0.0334 1.0000 0.0446
-8.500 -0.4637 0.08736 0.08065 -0.0348 1.0000 0.0442
-8.250 -0.4647 0.08338 0.07674 -0.0364 1.0000 0.0436
-8.000 -0.4689 0.07897 0.07241 -0.0390 1.0000 0.0431
-7.750 -0.4718 0.07460 0.06808 -0.0415 1.0000 0.0425
-7.500 -0.4749 0.07023 0.06372 -0.0434 1.0000 0.0420
-7.250 -0.4766 0.06614 0.05958 -0.0447 1.0000 0.0414
-7.000 -0.4779 0.06191 0.05527 -0.0457 1.0000 0.0408
-6.750 -0.4770 0.05785 0.05106 -0.0461 1.0000 0.0403
-6.500 -0.4738 0.05380 0.04680 -0.0462 1.0000 0.0399
-6.250 -0.4678 0.04991 0.04261 -0.0460 1.0000 0.0395
-6.000 -0.4591 0.04615 0.03851 -0.0455 1.0000 0.0395
-5.750 -0.4475 0.04267 0.03463 -0.0448 1.0000 0.0396
-5.500 -0.4332 0.03933 0.03083 -0.0439 1.0000 0.0401
-5.250 -0.4165 0.03631 0.02724 -0.0429 1.0000 0.0409
-5.000 -0.3977 0.03372 0.02412 -0.0418 1.0000 0.0430
-4.750 -0.3780 0.03144 0.02139 -0.0407 1.0000 0.0467
-4.500 -0.3586 0.02969 0.01948 -0.0397 1.0000 0.0511
-4.250 -0.3369 0.02784 0.01731 -0.0383 1.0000 0.0545
-4.000 -0.3155 0.02623 0.01545 -0.0368 1.0000 0.0601
-3.750 -0.2950 0.02514 0.01416 -0.0355 1.0000 0.0717
-3.500 -0.2748 0.02385 0.01285 -0.0341 1.0000 0.0838
-3.250 -0.2540 0.02276 0.01178 -0.0333 1.0000 0.1053
-3.000 -0.2322 0.02167 0.01067 -0.0327 1.0000 0.1365
-2.750 -0.2089 0.02029 0.00968 -0.0327 1.0000 0.2067
-2.500 -0.1934 0.01844 0.00962 -0.0309 1.0000 0.5437
-2.250 -0.1823 0.01832 0.00977 -0.0266 1.0000 0.6879
-2.000 -0.1707 0.01826 0.00978 -0.0225 0.9992 0.7731
-1.750 -0.1475 0.01810 0.00974 -0.0197 0.9916 0.8785
-1.500 -0.0837 0.01799 0.00929 -0.0265 0.9871 1.0000
-1.250 -0.0471 0.01815 0.00908 -0.0295 0.9785 1.0000
-1.000 -0.0090 0.01837 0.00898 -0.0327 0.9707 1.0000
-0.750 0.0271 0.01859 0.00888 -0.0353 0.9619 1.0000
-0.500 0.0629 0.01882 0.00888 -0.0377 0.9532 1.0000
-0.250 0.1021 0.01910 0.00895 -0.0407 0.9454 1.0000
0.000 0.1354 0.01933 0.00902 -0.0425 0.9355 1.0000
0.250 0.1710 0.01958 0.00911 -0.0446 0.9264 1.0000
0.500 0.2106 0.01983 0.00925 -0.0473 0.9182 1.0000
0.750 0.2435 0.02007 0.00942 -0.0488 0.9076 1.0000
1.000 0.2785 0.02030 0.00960 -0.0506 0.8975 1.0000
1.250 0.3184 0.02051 0.00978 -0.0531 0.8889 1.0000
1.500 0.3520 0.02071 0.00998 -0.0544 0.8780 1.0000
1.750 0.3837 0.02092 0.01023 -0.0552 0.8662 1.0000
2.000 0.4178 0.02103 0.01036 -0.0562 0.8531 1.0000
2.250 0.4542 0.02096 0.01035 -0.0572 0.8374 1.0000
2.500 0.4901 0.02080 0.01025 -0.0578 0.8209 1.0000
2.750 0.5171 0.02081 0.01038 -0.0571 0.8017 1.0000
3.000 0.5465 0.02081 0.01047 -0.0568 0.7852 1.0000
3.250 0.5763 0.02078 0.01055 -0.0565 0.7687 1.0000
3.500 0.6052 0.02074 0.01063 -0.0560 0.7511 1.0000
3.750 0.6308 0.02077 0.01086 -0.0550 0.7304 1.0000
4.000 0.6589 0.02070 0.01092 -0.0541 0.7095 1.0000
4.250 0.6852 0.02068 0.01104 -0.0530 0.6850 1.0000
4.500 0.7111 0.02067 0.01116 -0.0518 0.6574 1.0000
4.750 0.7360 0.02071 0.01131 -0.0504 0.6248 1.0000
5.000 0.7611 0.02076 0.01150 -0.0489 0.5872 1.0000
5.250 0.7844 0.02096 0.01171 -0.0472 0.5423 1.0000
5.500 0.8064 0.02130 0.01195 -0.0454 0.4911 1.0000
5.750 0.8260 0.02190 0.01234 -0.0435 0.4357 1.0000
6.000 0.8436 0.02274 0.01294 -0.0416 0.3805 1.0000
6.250 0.8601 0.02371 0.01380 -0.0399 0.3282 1.0000
6.500 0.8758 0.02483 0.01470 -0.0383 0.2807 1.0000
6.750 0.8915 0.02604 0.01580 -0.0368 0.2357 1.0000
7.000 0.9073 0.02734 0.01703 -0.0354 0.1929 1.0000
7.250 0.9229 0.02883 0.01841 -0.0339 0.1580 1.0000
7.500 0.9387 0.03040 0.01986 -0.0326 0.1301 1.0000
7.750 0.9560 0.03202 0.02148 -0.0314 0.1112 1.0000
8.000 0.9758 0.03379 0.02336 -0.0303 0.0997 1.0000
8.250 0.9949 0.03560 0.02528 -0.0292 0.0883 1.0000
8.500 1.0129 0.03752 0.02741 -0.0282 0.0780 1.0000
8.750 1.0346 0.03995 0.03005 -0.0273 0.0703 1.0000
9.000 1.0511 0.04262 0.03307 -0.0260 0.0611 1.0000
9.250 1.0594 0.04493 0.03558 -0.0245 0.0516 1.0000
9.500 1.0649 0.04696 0.03760 -0.0230 0.0441 1.0000
9.750 1.0693 0.04990 0.04110 -0.0211 0.0378 1.0000
10.000 1.0731 0.05297 0.04443 -0.0194 0.0348 1.0000
10.250 1.0737 0.05615 0.04776 -0.0176 0.0330 1.0000
10.500 1.0693 0.06002 0.05188 -0.0159 0.0320 1.0000
10.750 1.0592 0.06395 0.05619 -0.0143 0.0315 1.0000
11.000 1.0447 0.06831 0.06090 -0.0136 0.0312 1.0000
11.250 1.0291 0.07299 0.06584 -0.0139 0.0311 1.0000
11.500 1.0113 0.07830 0.07139 -0.0154 0.0311 1.0000
11.750 0.9933 0.08414 0.07742 -0.0179 0.0312 1.0000
12.000 0.9748 0.09074 0.08418 -0.0215 0.0314 1.0000
12.250 0.9574 0.09794 0.09150 -0.0260 0.0317 1.0000
12.500 0.9407 0.10588 0.09952 -0.0310 0.0320 1.0000
12.750 0.9260 0.11416 0.10785 -0.0362 0.0323 1.0000
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Polar data table (+)
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