HQ 1.5/8.5 AIRFOIL (hq1585-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/8.5 AIRFOIL (hq1585-il) Reynolds number: 200,000 Max Cl/Cd: 58.3 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1585-il-200000-n5.txt Download as CSV file: xf-hq1585-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/8.5 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5447 0.08732 0.08381 -0.0180 1.0000 0.0126
-9.250 -0.5500 0.08165 0.07822 -0.0213 1.0000 0.0126
-9.000 -0.5581 0.07535 0.07198 -0.0254 1.0000 0.0123
-8.750 -0.5717 0.06721 0.06392 -0.0324 1.0000 0.0123
-8.500 -0.5923 0.05815 0.05477 -0.0401 1.0000 0.0118
-8.250 -0.6064 0.05028 0.04665 -0.0436 1.0000 0.0115
-8.000 -0.6102 0.04421 0.04025 -0.0448 1.0000 0.0116
-7.750 -0.6069 0.03914 0.03479 -0.0450 1.0000 0.0120
-7.500 -0.5988 0.03474 0.02996 -0.0446 1.0000 0.0124
-7.250 -0.5854 0.03161 0.02643 -0.0439 1.0000 0.0137
-7.000 -0.5707 0.02848 0.02279 -0.0427 1.0000 0.0149
-6.750 -0.5541 0.02604 0.01988 -0.0415 1.0000 0.0156
-6.500 -0.5386 0.02314 0.01654 -0.0401 1.0000 0.0164
-6.250 -0.5211 0.02160 0.01485 -0.0388 1.0000 0.0174
-6.000 -0.4922 0.02049 0.01358 -0.0397 0.9957 0.0190
-5.750 -0.4587 0.01959 0.01248 -0.0412 0.9893 0.0219
-5.500 -0.4250 0.01836 0.01100 -0.0426 0.9837 0.0234
-5.250 -0.3934 0.01671 0.00921 -0.0438 0.9776 0.0256
-5.000 -0.3611 0.01594 0.00839 -0.0451 0.9709 0.0286
-4.750 -0.3282 0.01506 0.00739 -0.0464 0.9649 0.0308
-4.500 -0.2969 0.01432 0.00654 -0.0473 0.9574 0.0331
-4.250 -0.2637 0.01373 0.00584 -0.0485 0.9515 0.0352
-4.000 -0.2335 0.01305 0.00514 -0.0493 0.9435 0.0410
-3.750 -0.2010 0.01262 0.00461 -0.0503 0.9374 0.0482
-3.500 -0.1717 0.01212 0.00407 -0.0507 0.9293 0.0612
-3.250 -0.1416 0.01154 0.00364 -0.0515 0.9229 0.0984
-3.000 -0.1152 0.01083 0.00335 -0.0517 0.9145 0.2031
-2.750 -0.0885 0.01006 0.00312 -0.0520 0.9076 0.3398
-2.500 -0.0645 0.00940 0.00304 -0.0515 0.8994 0.4908
-2.250 -0.0390 0.00911 0.00305 -0.0508 0.8926 0.5870
-2.000 -0.0135 0.00899 0.00306 -0.0500 0.8849 0.6435
-1.750 0.0127 0.00895 0.00308 -0.0493 0.8782 0.6860
-1.500 0.0391 0.00894 0.00305 -0.0488 0.8707 0.7129
-1.250 0.0661 0.00893 0.00301 -0.0483 0.8639 0.7307
-1.000 0.0927 0.00891 0.00296 -0.0478 0.8556 0.7469
-0.750 0.1190 0.00889 0.00290 -0.0472 0.8458 0.7629
-0.500 0.1450 0.00887 0.00284 -0.0465 0.8362 0.7809
-0.250 0.1707 0.00883 0.00280 -0.0457 0.8264 0.7971
0.000 0.1962 0.00879 0.00276 -0.0449 0.8154 0.8122
0.250 0.2211 0.00873 0.00272 -0.0438 0.8035 0.8299
0.500 0.2460 0.00868 0.00269 -0.0428 0.7904 0.8477
0.750 0.2718 0.00864 0.00265 -0.0420 0.7773 0.8637
1.000 0.2985 0.00862 0.00265 -0.0415 0.7655 0.8799
1.250 0.3264 0.00859 0.00264 -0.0412 0.7549 0.8986
1.500 0.3567 0.00856 0.00265 -0.0415 0.7425 0.9217
1.750 0.3915 0.00854 0.00267 -0.0429 0.7293 0.9499
2.000 0.4286 0.00853 0.00268 -0.0448 0.7138 0.9962
2.250 0.4556 0.00859 0.00272 -0.0446 0.6964 1.0000
2.500 0.4820 0.00868 0.00277 -0.0443 0.6758 1.0000
2.750 0.5071 0.00881 0.00277 -0.0435 0.6365 1.0000
3.000 0.5305 0.00910 0.00277 -0.0425 0.5734 1.0000
3.250 0.5528 0.00958 0.00286 -0.0414 0.4887 1.0000
3.500 0.5756 0.01011 0.00306 -0.0406 0.4204 1.0000
3.750 0.5987 0.01068 0.00337 -0.0399 0.3565 1.0000
4.000 0.6226 0.01121 0.00367 -0.0395 0.3078 1.0000
4.250 0.6472 0.01166 0.00399 -0.0391 0.2701 1.0000
4.500 0.6713 0.01219 0.00434 -0.0387 0.2263 1.0000
4.750 0.6957 0.01269 0.00471 -0.0383 0.1917 1.0000
5.000 0.7199 0.01324 0.00511 -0.0379 0.1496 1.0000
5.250 0.7427 0.01400 0.00566 -0.0374 0.1006 1.0000
5.500 0.7656 0.01477 0.00625 -0.0368 0.0670 1.0000
5.750 0.7895 0.01538 0.00684 -0.0363 0.0562 1.0000
6.000 0.8136 0.01598 0.00747 -0.0358 0.0513 1.0000
6.250 0.8373 0.01659 0.00815 -0.0352 0.0478 1.0000
6.500 0.8613 0.01714 0.00878 -0.0348 0.0422 1.0000
6.750 0.8835 0.01793 0.00963 -0.0342 0.0362 1.0000
7.000 0.9075 0.01846 0.01031 -0.0336 0.0333 1.0000
7.250 0.9307 0.01909 0.01105 -0.0331 0.0286 1.0000
7.500 0.9526 0.01991 0.01194 -0.0324 0.0228 1.0000
7.750 0.9756 0.02055 0.01264 -0.0319 0.0176 1.0000
8.000 0.9960 0.02159 0.01379 -0.0309 0.0148 1.0000
8.250 1.0164 0.02259 0.01490 -0.0300 0.0131 1.0000
8.500 1.0358 0.02367 0.01610 -0.0290 0.0119 1.0000
8.750 1.0525 0.02519 0.01775 -0.0277 0.0111 1.0000
9.000 1.0700 0.02661 0.01938 -0.0265 0.0103 1.0000
9.250 1.0873 0.02800 0.02095 -0.0253 0.0094 1.0000
9.500 1.1038 0.02936 0.02248 -0.0242 0.0087 1.0000
9.750 1.1176 0.03110 0.02443 -0.0227 0.0082 1.0000
10.000 1.1293 0.03304 0.02659 -0.0212 0.0079 1.0000
10.250 1.1378 0.03522 0.02902 -0.0194 0.0076 1.0000
10.500 1.1421 0.03771 0.03180 -0.0173 0.0075 1.0000
10.750 1.1391 0.04044 0.03482 -0.0146 0.0073 1.0000
11.000 1.1326 0.04335 0.03800 -0.0121 0.0072 1.0000
11.250 1.1220 0.04671 0.04165 -0.0101 0.0072 1.0000
11.500 1.1071 0.05071 0.04594 -0.0091 0.0071 1.0000
11.750 1.0922 0.05503 0.05051 -0.0091 0.0071 1.0000
12.000 1.0797 0.05945 0.05515 -0.0101 0.0071 1.0000
12.250 1.0611 0.06531 0.06124 -0.0126 0.0071 1.0000
12.500 1.0444 0.07170 0.06784 -0.0164 0.0071 1.0000
12.750 1.0281 0.07903 0.07536 -0.0213 0.0072 1.0000
13.000 1.0068 0.08872 0.08523 -0.0282 0.0073 1.0000
13.250 0.9755 0.10266 0.09940 -0.0375 0.0076 1.0000
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Polar data table (+)
Polar graphs
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