HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/9 AIRFOIL (hq159-il) Reynolds number: 50,000 Max Cl/Cd: 35.66 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq159-il-50000-n5.txt Download as CSV file: xf-hq159-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/9 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5057 0.09619 0.08928 -0.0222 1.0000 0.0430
-9.000 -0.5072 0.09156 0.08473 -0.0247 1.0000 0.0428
-8.750 -0.5100 0.08674 0.07999 -0.0277 1.0000 0.0428
-8.500 -0.5136 0.08179 0.07511 -0.0310 1.0000 0.0420
-8.250 -0.5212 0.07655 0.06994 -0.0348 1.0000 0.0410
-8.000 -0.5286 0.07169 0.06509 -0.0377 1.0000 0.0409
-7.750 -0.5340 0.06692 0.06025 -0.0399 1.0000 0.0400
-7.500 -0.5373 0.06232 0.05553 -0.0414 1.0000 0.0400
-7.250 -0.5377 0.05790 0.05090 -0.0422 1.0000 0.0400
-7.000 -0.5346 0.05364 0.04638 -0.0425 1.0000 0.0399
-6.750 -0.5285 0.04955 0.04196 -0.0423 1.0000 0.0398
-6.500 -0.5196 0.04566 0.03767 -0.0417 1.0000 0.0400
-6.250 -0.5079 0.04206 0.03365 -0.0408 1.0000 0.0404
-6.000 -0.4936 0.03876 0.02989 -0.0396 1.0000 0.0412
-5.750 -0.4768 0.03574 0.02638 -0.0384 1.0000 0.0425
-5.500 -0.4578 0.03310 0.02321 -0.0370 1.0000 0.0452
-5.250 -0.4392 0.03096 0.02081 -0.0358 1.0000 0.0498
-5.000 -0.4194 0.02919 0.01874 -0.0345 1.0000 0.0549
-4.750 -0.3976 0.02732 0.01648 -0.0328 1.0000 0.0596
-4.500 -0.3786 0.02591 0.01507 -0.0314 1.0000 0.0703
-4.250 -0.3593 0.02451 0.01361 -0.0297 1.0000 0.0813
-4.000 -0.3401 0.02338 0.01235 -0.0282 1.0000 0.0999
-3.750 -0.3210 0.02211 0.01115 -0.0269 1.0000 0.1196
-3.500 -0.3017 0.02094 0.01014 -0.0260 1.0000 0.1588
-3.250 -0.2827 0.01939 0.00913 -0.0254 1.0000 0.2499
-3.000 -0.2710 0.01786 0.00902 -0.0226 1.0000 0.5180
-2.750 -0.2591 0.01770 0.00914 -0.0185 1.0000 0.6524
-2.500 -0.2475 0.01768 0.00918 -0.0144 1.0000 0.7304
-2.250 -0.2371 0.01764 0.00917 -0.0098 1.0000 0.7983
-2.000 -0.2227 0.01752 0.00907 -0.0058 1.0000 0.8661
-1.750 -0.1717 0.01751 0.00881 -0.0091 1.0000 0.9422
-1.500 -0.1120 0.01749 0.00844 -0.0161 1.0000 1.0000
-1.250 -0.0965 0.01744 0.00817 -0.0158 0.9962 1.0000
-1.000 -0.0591 0.01760 0.00806 -0.0191 0.9865 1.0000
-0.750 -0.0224 0.01779 0.00797 -0.0221 0.9768 1.0000
-0.500 0.0146 0.01801 0.00798 -0.0249 0.9675 1.0000
-0.250 0.0538 0.01826 0.00806 -0.0281 0.9588 1.0000
0.000 0.0886 0.01848 0.00811 -0.0302 0.9483 1.0000
0.250 0.1245 0.01871 0.00823 -0.0325 0.9383 1.0000
0.500 0.1632 0.01895 0.00839 -0.0352 0.9290 1.0000
0.750 0.2006 0.01916 0.00856 -0.0376 0.9189 1.0000
1.000 0.2355 0.01939 0.00875 -0.0394 0.9080 1.0000
1.250 0.2718 0.01960 0.00896 -0.0414 0.8975 1.0000
1.500 0.3086 0.01980 0.00917 -0.0434 0.8872 1.0000
1.750 0.3458 0.01995 0.00939 -0.0452 0.8763 1.0000
2.000 0.3800 0.02002 0.00951 -0.0461 0.8613 1.0000
2.250 0.4119 0.02002 0.00957 -0.0464 0.8425 1.0000
2.500 0.4457 0.01992 0.00956 -0.0466 0.8240 1.0000
2.750 0.4782 0.01986 0.00964 -0.0467 0.8071 1.0000
3.000 0.5042 0.01996 0.00984 -0.0460 0.7883 1.0000
3.250 0.5323 0.01998 0.00998 -0.0454 0.7699 1.0000
3.500 0.5608 0.01996 0.01008 -0.0447 0.7508 1.0000
3.750 0.5861 0.02000 0.01033 -0.0435 0.7282 1.0000
4.000 0.6117 0.02000 0.01048 -0.0423 0.7035 1.0000
4.250 0.6370 0.02000 0.01061 -0.0409 0.6756 1.0000
4.500 0.6622 0.01998 0.01071 -0.0394 0.6431 1.0000
4.750 0.6857 0.02003 0.01084 -0.0376 0.6025 1.0000
5.000 0.7082 0.02015 0.01103 -0.0357 0.5520 1.0000
5.250 0.7293 0.02045 0.01118 -0.0335 0.4912 1.0000
5.500 0.7484 0.02107 0.01153 -0.0315 0.4252 1.0000
5.750 0.7658 0.02196 0.01213 -0.0296 0.3613 1.0000
6.000 0.7825 0.02303 0.01294 -0.0280 0.3040 1.0000
6.250 0.7988 0.02426 0.01393 -0.0265 0.2527 1.0000
6.500 0.8152 0.02556 0.01518 -0.0252 0.2032 1.0000
6.750 0.8324 0.02697 0.01647 -0.0239 0.1639 1.0000
7.000 0.8499 0.02850 0.01789 -0.0227 0.1354 1.0000
7.250 0.8682 0.03003 0.01939 -0.0216 0.1128 1.0000
7.500 0.8887 0.03177 0.02121 -0.0205 0.0996 1.0000
7.750 0.9088 0.03353 0.02307 -0.0196 0.0868 1.0000
8.000 0.9288 0.03544 0.02513 -0.0186 0.0758 1.0000
8.250 0.9508 0.03787 0.02762 -0.0178 0.0683 1.0000
8.500 0.9682 0.04027 0.03042 -0.0165 0.0584 1.0000
8.750 0.9842 0.04332 0.03386 -0.0153 0.0517 1.0000
9.000 0.9945 0.04556 0.03629 -0.0140 0.0444 1.0000
9.250 1.0016 0.04882 0.04003 -0.0124 0.0394 1.0000
9.500 1.0064 0.05187 0.04346 -0.0108 0.0356 1.0000
9.750 1.0095 0.05458 0.04635 -0.0094 0.0331 1.0000
10.000 1.0084 0.05812 0.04999 -0.0081 0.0314 1.0000
10.250 0.9990 0.06188 0.05415 -0.0062 0.0311 1.0000
10.500 0.9852 0.06585 0.05844 -0.0048 0.0308 1.0000
10.750 0.9699 0.07013 0.06296 -0.0046 0.0308 1.0000
11.000 0.9519 0.07504 0.06810 -0.0056 0.0308 1.0000
11.250 0.9341 0.08050 0.07374 -0.0079 0.0309 1.0000
11.500 0.9168 0.08659 0.07996 -0.0112 0.0311 1.0000
11.750 0.8993 0.09358 0.08704 -0.0156 0.0313 1.0000
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Polar data table (+)
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