HQ 1.5/9 B AIRFOIL (hq159b-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/9 B AIRFOIL (hq159b-il) Reynolds number: 200,000 Max Cl/Cd: 60.59 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq159b-il-200000-n5.txt Download as CSV file: xf-hq159b-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/9 B 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 -0.5309 0.08569 0.08215 -0.0229 1.0000 0.0117
-9.250 -0.5363 0.08026 0.07677 -0.0260 1.0000 0.0110
-9.000 -0.5447 0.07404 0.07067 -0.0303 1.0000 0.0112
-8.750 -0.5595 0.06616 0.06285 -0.0375 1.0000 0.0109
-8.500 -0.5767 0.05974 0.05636 -0.0420 1.0000 0.0108
-8.250 -0.5901 0.05294 0.04938 -0.0445 1.0000 0.0103
-8.000 -0.5961 0.04708 0.04326 -0.0452 1.0000 0.0104
-7.750 -0.5964 0.04192 0.03778 -0.0449 1.0000 0.0110
-7.500 -0.5957 0.03636 0.03178 -0.0438 1.0000 0.0111
-7.250 -0.5913 0.03124 0.02611 -0.0420 1.0000 0.0111
-7.000 -0.5815 0.02727 0.02158 -0.0401 1.0000 0.0114
-6.750 -0.5679 0.02428 0.01811 -0.0382 1.0000 0.0118
-6.500 -0.5522 0.02200 0.01543 -0.0364 1.0000 0.0123
-6.250 -0.5246 0.01995 0.01297 -0.0368 0.9965 0.0131
-6.000 -0.4929 0.01839 0.01112 -0.0379 0.9915 0.0142
-5.750 -0.4607 0.01742 0.01011 -0.0394 0.9862 0.0171
-5.500 -0.4266 0.01643 0.00894 -0.0410 0.9820 0.0208
-5.250 -0.3956 0.01552 0.00799 -0.0420 0.9758 0.0250
-5.000 -0.3618 0.01486 0.00726 -0.0436 0.9711 0.0311
-4.750 -0.3284 0.01464 0.00697 -0.0449 0.9656 0.0390
-4.500 -0.2970 0.01389 0.00615 -0.0460 0.9592 0.0449
-4.250 -0.2636 0.01340 0.00554 -0.0473 0.9539 0.0502
-4.000 -0.2337 0.01290 0.00504 -0.0479 0.9461 0.0589
-3.750 -0.2013 0.01246 0.00455 -0.0490 0.9402 0.0692
-3.500 -0.1721 0.01205 0.00415 -0.0494 0.9318 0.0843
-3.250 -0.1421 0.01157 0.00374 -0.0501 0.9249 0.1167
-3.000 -0.1150 0.01083 0.00343 -0.0505 0.9165 0.2243
-2.750 -0.0900 0.00994 0.00319 -0.0505 0.9080 0.3884
-2.500 -0.0639 0.00941 0.00311 -0.0502 0.9002 0.5154
-2.250 -0.0385 0.00919 0.00311 -0.0495 0.8906 0.5967
-2.000 -0.0118 0.00910 0.00307 -0.0489 0.8816 0.6479
-1.750 0.0148 0.00903 0.00303 -0.0483 0.8728 0.6830
-1.500 0.0406 0.00899 0.00301 -0.0475 0.8628 0.7119
-1.250 0.0670 0.00896 0.00295 -0.0469 0.8535 0.7335
-1.000 0.0933 0.00893 0.00291 -0.0462 0.8450 0.7526
-0.750 0.1195 0.00891 0.00289 -0.0456 0.8360 0.7692
-0.500 0.1460 0.00889 0.00285 -0.0451 0.8272 0.7828
-0.250 0.1722 0.00886 0.00279 -0.0444 0.8168 0.7967
0.000 0.1980 0.00883 0.00275 -0.0437 0.8042 0.8101
0.250 0.2241 0.00880 0.00271 -0.0431 0.7919 0.8222
0.500 0.2503 0.00878 0.00268 -0.0426 0.7798 0.8341
0.750 0.2764 0.00876 0.00265 -0.0419 0.7667 0.8470
1.000 0.3025 0.00873 0.00262 -0.0413 0.7521 0.8611
1.250 0.3292 0.00872 0.00262 -0.0408 0.7375 0.8768
1.500 0.3568 0.00871 0.00263 -0.0405 0.7255 0.8956
1.750 0.3869 0.00870 0.00266 -0.0407 0.7133 0.9193
2.000 0.4217 0.00871 0.00269 -0.0420 0.6990 0.9480
2.250 0.4596 0.00873 0.00273 -0.0441 0.6823 0.9843
2.500 0.4887 0.00882 0.00279 -0.0445 0.6655 1.0000
2.750 0.5147 0.00894 0.00289 -0.0441 0.6474 1.0000
3.000 0.5403 0.00908 0.00299 -0.0436 0.6216 1.0000
3.250 0.5641 0.00931 0.00304 -0.0427 0.5685 1.0000
3.500 0.5850 0.00981 0.00311 -0.0414 0.4827 1.0000
3.750 0.6061 0.01044 0.00335 -0.0403 0.4049 1.0000
4.000 0.6280 0.01107 0.00364 -0.0395 0.3315 1.0000
4.250 0.6508 0.01165 0.00401 -0.0389 0.2813 1.0000
4.500 0.6748 0.01212 0.00436 -0.0384 0.2496 1.0000
4.750 0.6990 0.01258 0.00473 -0.0380 0.2203 1.0000
5.000 0.7231 0.01306 0.00511 -0.0375 0.1852 1.0000
5.250 0.7461 0.01369 0.00552 -0.0370 0.1331 1.0000
5.500 0.7678 0.01451 0.00610 -0.0363 0.0888 1.0000
5.750 0.7900 0.01530 0.00680 -0.0356 0.0593 1.0000
6.000 0.8130 0.01599 0.00745 -0.0349 0.0458 1.0000
6.250 0.8364 0.01660 0.00810 -0.0343 0.0400 1.0000
6.500 0.8587 0.01734 0.00891 -0.0336 0.0365 1.0000
6.750 0.8808 0.01809 0.00977 -0.0327 0.0345 1.0000
7.000 0.9030 0.01880 0.01061 -0.0320 0.0318 1.0000
7.250 0.9237 0.01967 0.01151 -0.0312 0.0280 1.0000
7.500 0.9438 0.02066 0.01260 -0.0302 0.0251 1.0000
7.750 0.9649 0.02147 0.01351 -0.0294 0.0219 1.0000
8.000 0.9820 0.02284 0.01499 -0.0281 0.0196 1.0000
8.250 1.0011 0.02406 0.01637 -0.0270 0.0180 1.0000
8.500 1.0210 0.02514 0.01764 -0.0260 0.0157 1.0000
8.750 1.0401 0.02619 0.01883 -0.0250 0.0136 1.0000
9.000 1.0573 0.02748 0.02024 -0.0238 0.0124 1.0000
9.250 1.0712 0.02940 0.02231 -0.0224 0.0115 1.0000
9.500 1.0834 0.03198 0.02513 -0.0207 0.0109 1.0000
9.750 1.0964 0.03411 0.02756 -0.0191 0.0105 1.0000
10.000 1.1061 0.03652 0.03030 -0.0173 0.0101 1.0000
10.500 1.1155 0.04100 0.03543 -0.0129 0.0089 1.0000
10.750 1.1130 0.04333 0.03803 -0.0103 0.0085 1.0000
11.000 1.1062 0.04607 0.04104 -0.0079 0.0082 1.0000
11.250 1.0981 0.04893 0.04414 -0.0062 0.0079 1.0000
11.500 1.0859 0.05253 0.04799 -0.0052 0.0078 1.0000
11.750 1.0706 0.05679 0.05249 -0.0053 0.0077 1.0000
12.000 1.0517 0.06207 0.05802 -0.0066 0.0078 1.0000
12.250 1.0313 0.06823 0.06440 -0.0094 0.0078 1.0000
12.500 1.0092 0.07566 0.07204 -0.0139 0.0079 1.0000
12.750 0.9849 0.08491 0.08147 -0.0203 0.0080 1.0000
13.000 0.9454 0.10008 0.09681 -0.0308 0.0088 1.0000
13.250 0.9079 0.11654 0.11333 -0.0406 0.0097 1.0000
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Polar data table (+)
Polar graphs
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