HQ 1.0/10 AIRFOIL (hq1010-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 1.0/10 AIRFOIL (hq1010-il) Reynolds number: 200,000 Max Cl/Cd: 63.07 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq1010-il-200000.txt Download as CSV file: xf-hq1010-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.0/10 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4657 0.10907 0.10576 -0.0149 1.0000 0.0541
-10.750 -0.4760 0.10350 0.10023 -0.0190 1.0000 0.0545
-10.500 -0.4838 0.09769 0.09444 -0.0225 1.0000 0.0547
-10.250 -0.4931 0.09120 0.08797 -0.0265 1.0000 0.0548
-10.000 -0.4809 0.08725 0.08405 -0.0224 1.0000 0.0566
-9.750 -0.4777 0.08339 0.08020 -0.0227 1.0000 0.0576
-9.500 -0.4793 0.07879 0.07561 -0.0240 1.0000 0.0587
-9.250 -0.4860 0.07337 0.07022 -0.0263 1.0000 0.0595
-9.000 -0.4988 0.06681 0.06370 -0.0300 1.0000 0.0601
-8.000 -0.6683 0.04381 0.03892 -0.0378 1.0000 0.0319
-7.750 -0.6617 0.03862 0.03359 -0.0367 1.0000 0.0299
-7.500 -0.6558 0.03422 0.02878 -0.0348 1.0000 0.0288
-7.250 -0.6460 0.03040 0.02447 -0.0327 1.0000 0.0281
-7.000 -0.6327 0.02732 0.02097 -0.0306 1.0000 0.0281
-6.750 -0.6168 0.02484 0.01811 -0.0286 1.0000 0.0288
-6.500 -0.5995 0.02296 0.01593 -0.0266 1.0000 0.0299
-6.250 -0.5828 0.02142 0.01412 -0.0247 1.0000 0.0322
-6.000 -0.5676 0.01990 0.01260 -0.0228 1.0000 0.0356
-5.750 -0.5515 0.01890 0.01150 -0.0208 1.0000 0.0390
-5.500 -0.5351 0.01797 0.01041 -0.0186 1.0000 0.0429
-5.250 -0.5205 0.01684 0.00934 -0.0168 1.0000 0.0503
-5.000 -0.5047 0.01593 0.00839 -0.0149 1.0000 0.0592
-4.750 -0.4860 0.01528 0.00771 -0.0137 0.9996 0.0705
-4.500 -0.4478 0.01452 0.00699 -0.0164 0.9947 0.0882
-4.250 -0.4113 0.01375 0.00629 -0.0187 0.9890 0.1111
-4.000 -0.3745 0.01271 0.00560 -0.0213 0.9838 0.1694
-3.750 -0.3440 0.01130 0.00509 -0.0231 0.9773 0.3536
-3.500 -0.3108 0.01056 0.00504 -0.0246 0.9714 0.5290
-3.250 -0.2757 0.01042 0.00508 -0.0260 0.9653 0.6048
-3.000 -0.2398 0.01040 0.00509 -0.0274 0.9591 0.6536
-2.750 -0.1990 0.01041 0.00512 -0.0297 0.9554 0.6907
-2.500 -0.1672 0.01043 0.00512 -0.0303 0.9474 0.7168
-2.250 -0.1298 0.01046 0.00513 -0.0318 0.9424 0.7432
-2.000 -0.0996 0.01049 0.00518 -0.0317 0.9346 0.7650
-1.750 -0.0654 0.01047 0.00514 -0.0325 0.9286 0.7839
-1.500 -0.0361 0.01044 0.00509 -0.0325 0.9202 0.7979
-1.250 -0.0041 0.01037 0.00499 -0.0328 0.9137 0.8110
-1.000 0.0218 0.01035 0.00496 -0.0320 0.9041 0.8254
-0.750 0.0492 0.01030 0.00491 -0.0314 0.8967 0.8409
-0.500 0.0737 0.01027 0.00488 -0.0302 0.8874 0.8560
-0.250 0.0983 0.01023 0.00485 -0.0291 0.8780 0.8697
0.000 0.1242 0.01014 0.00475 -0.0281 0.8687 0.8813
0.250 0.1493 0.01004 0.00464 -0.0269 0.8573 0.8930
0.500 0.1746 0.00996 0.00457 -0.0259 0.8453 0.9053
0.750 0.2017 0.00990 0.00449 -0.0252 0.8334 0.9177
1.000 0.2315 0.00984 0.00442 -0.0252 0.8218 0.9302
1.250 0.2644 0.00980 0.00437 -0.0258 0.8111 0.9426
1.500 0.2999 0.00977 0.00434 -0.0272 0.8007 0.9547
1.750 0.3404 0.00975 0.00433 -0.0296 0.7887 0.9639
2.000 0.3811 0.00971 0.00430 -0.0322 0.7763 0.9733
2.250 0.4222 0.00966 0.00428 -0.0349 0.7627 0.9830
2.500 0.4646 0.00959 0.00421 -0.0379 0.7471 0.9923
2.750 0.5000 0.00951 0.00412 -0.0396 0.7297 1.0000
3.000 0.5152 0.00948 0.00410 -0.0374 0.7117 1.0000
3.250 0.5330 0.00948 0.00408 -0.0355 0.6914 1.0000
3.500 0.5536 0.00953 0.00408 -0.0340 0.6691 1.0000
3.750 0.5755 0.00960 0.00409 -0.0327 0.6418 1.0000
4.000 0.5979 0.00971 0.00416 -0.0314 0.6077 1.0000
4.250 0.6203 0.00989 0.00423 -0.0302 0.5659 1.0000
4.500 0.6421 0.01018 0.00434 -0.0289 0.5128 1.0000
4.750 0.6625 0.01067 0.00454 -0.0275 0.4470 1.0000
5.000 0.6824 0.01130 0.00486 -0.0262 0.3794 1.0000
5.250 0.7029 0.01196 0.00529 -0.0251 0.3241 1.0000
5.500 0.7240 0.01262 0.00574 -0.0242 0.2815 1.0000
5.750 0.7457 0.01325 0.00624 -0.0233 0.2447 1.0000
6.000 0.7673 0.01389 0.00675 -0.0225 0.2028 1.0000
6.250 0.7882 0.01465 0.00729 -0.0216 0.1568 1.0000
6.500 0.8080 0.01557 0.00803 -0.0206 0.1223 1.0000
6.750 0.8271 0.01659 0.00889 -0.0194 0.0988 1.0000
7.000 0.8477 0.01747 0.00978 -0.0184 0.0804 1.0000
7.250 0.8671 0.01851 0.01080 -0.0172 0.0663 1.0000
7.500 0.8865 0.01955 0.01183 -0.0161 0.0537 1.0000
7.750 0.9050 0.02077 0.01308 -0.0148 0.0440 1.0000
8.000 0.9219 0.02232 0.01461 -0.0134 0.0364 1.0000
8.250 0.9423 0.02319 0.01563 -0.0123 0.0304 1.0000
8.500 0.9579 0.02582 0.01831 -0.0108 0.0270 1.0000
8.750 0.9782 0.02748 0.02019 -0.0096 0.0251 1.0000
9.000 0.9972 0.02957 0.02252 -0.0083 0.0236 1.0000
9.250 1.0144 0.03168 0.02487 -0.0071 0.0225 1.0000
9.500 1.0296 0.03323 0.02654 -0.0059 0.0209 1.0000
9.750 1.0395 0.03654 0.03005 -0.0045 0.0196 1.0000
10.000 1.0408 0.04072 0.03467 -0.0022 0.0190 1.0000
10.250 1.0427 0.04358 0.03789 0.0001 0.0189 1.0000
10.500 1.0381 0.04692 0.04157 0.0027 0.0188 1.0000
10.750 1.0266 0.05008 0.04500 0.0058 0.0188 1.0000
11.000 1.0120 0.05358 0.04874 0.0082 0.0189 1.0000
11.250 0.9944 0.05704 0.05245 0.0097 0.0188 1.0000
11.500 0.9110 0.05420 0.05000 0.0116 0.0193 1.0000
11.750 0.8946 0.05872 0.05469 0.0108 0.0194 1.0000
12.000 0.8752 0.06422 0.06037 0.0090 0.0195 1.0000
12.250 0.8555 0.07030 0.06663 0.0063 0.0197 1.0000
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Polar data table (+)
Polar graphs
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