HQ 1.5/10 AIRFOIL (hq1510-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/10 AIRFOIL (hq1510-il) Reynolds number: 50,000 Max Cl/Cd: 35.14 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq1510-il-50000-n5.txt Download as CSV file: xf-hq1510-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 1.5/10 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.5129 0.09503 0.08806 -0.0271 1.0000 0.0424
-9.500 -0.5161 0.08987 0.08297 -0.0299 1.0000 0.0420
-9.250 -0.5211 0.08423 0.07740 -0.0334 1.0000 0.0416
-9.000 -0.5294 0.07816 0.07139 -0.0378 1.0000 0.0411
-8.750 -0.5437 0.07273 0.06601 -0.0414 1.0000 0.0406
-8.500 -0.5586 0.06801 0.06126 -0.0434 1.0000 0.0402
-8.250 -0.5687 0.06352 0.05666 -0.0446 1.0000 0.0398
-8.000 -0.5756 0.05919 0.05215 -0.0451 1.0000 0.0401
-7.750 -0.5786 0.05511 0.04782 -0.0450 1.0000 0.0407
-7.500 -0.5784 0.05117 0.04354 -0.0444 1.0000 0.0419
-7.250 -0.5751 0.04736 0.03927 -0.0434 1.0000 0.0434
-7.000 -0.5681 0.04382 0.03516 -0.0421 1.0000 0.0448
-6.750 -0.5576 0.04051 0.03149 -0.0407 1.0000 0.0462
-6.500 -0.5440 0.03784 0.02860 -0.0394 1.0000 0.0479
-6.250 -0.5284 0.03544 0.02592 -0.0379 1.0000 0.0503
-6.000 -0.5116 0.03338 0.02347 -0.0364 1.0000 0.0555
-5.750 -0.4942 0.03136 0.02112 -0.0349 1.0000 0.0607
-5.500 -0.4756 0.02962 0.01922 -0.0334 1.0000 0.0664
-5.250 -0.4574 0.02814 0.01759 -0.0318 1.0000 0.0760
-5.000 -0.4381 0.02679 0.01601 -0.0301 1.0000 0.0857
-4.750 -0.4204 0.02569 0.01486 -0.0285 1.0000 0.0991
-4.500 -0.4027 0.02455 0.01368 -0.0268 1.0000 0.1108
-4.250 -0.3853 0.02354 0.01268 -0.0253 1.0000 0.1275
-4.000 -0.3674 0.02254 0.01175 -0.0240 1.0000 0.1509
-3.750 -0.3490 0.02145 0.01080 -0.0231 1.0000 0.1866
-3.500 -0.3319 0.02008 0.01000 -0.0223 1.0000 0.2791
-3.250 -0.3192 0.01891 0.00994 -0.0198 1.0000 0.4772
-3.000 -0.3063 0.01879 0.01015 -0.0163 1.0000 0.6018
-2.750 -0.2929 0.01886 0.01025 -0.0130 1.0000 0.6749
-2.500 -0.2754 0.01901 0.01041 -0.0102 0.9965 0.7365
-2.250 -0.2516 0.01919 0.01055 -0.0081 0.9880 0.7972
-2.000 -0.2247 0.01929 0.01055 -0.0069 0.9802 0.8457
-1.750 -0.1886 0.01937 0.01049 -0.0076 0.9736 0.8870
-1.500 -0.1372 0.01951 0.01040 -0.0118 0.9694 0.9238
-1.250 -0.0832 0.01958 0.01023 -0.0173 0.9640 0.9483
-1.000 -0.0272 0.01966 0.01009 -0.0237 0.9589 0.9676
-0.750 0.0255 0.01968 0.00991 -0.0295 0.9523 0.9873
-0.500 0.0687 0.01975 0.00983 -0.0338 0.9445 1.0000
-0.250 0.0967 0.01980 0.00976 -0.0352 0.9330 1.0000
0.000 0.1264 0.01989 0.00973 -0.0366 0.9221 1.0000
0.250 0.1610 0.02001 0.00976 -0.0388 0.9127 1.0000
0.500 0.1965 0.02014 0.00982 -0.0410 0.9033 1.0000
0.750 0.2264 0.02030 0.00992 -0.0420 0.8922 1.0000
1.000 0.2576 0.02047 0.01005 -0.0431 0.8816 1.0000
1.250 0.2923 0.02061 0.01017 -0.0447 0.8720 1.0000
1.500 0.3243 0.02077 0.01034 -0.0457 0.8611 1.0000
1.750 0.3537 0.02093 0.01051 -0.0461 0.8486 1.0000
2.000 0.3864 0.02094 0.01055 -0.0467 0.8342 1.0000
2.250 0.4204 0.02083 0.01047 -0.0471 0.8182 1.0000
2.500 0.4472 0.02081 0.01052 -0.0463 0.7994 1.0000
2.750 0.4750 0.02080 0.01057 -0.0458 0.7823 1.0000
3.000 0.5027 0.02083 0.01067 -0.0452 0.7665 1.0000
3.250 0.5304 0.02083 0.01080 -0.0446 0.7501 1.0000
3.500 0.5585 0.02079 0.01086 -0.0439 0.7334 1.0000
3.750 0.5819 0.02090 0.01108 -0.0427 0.7128 1.0000
4.000 0.6087 0.02085 0.01114 -0.0417 0.6926 1.0000
4.250 0.6329 0.02089 0.01130 -0.0404 0.6686 1.0000
4.500 0.6576 0.02090 0.01147 -0.0391 0.6424 1.0000
4.750 0.6819 0.02093 0.01159 -0.0376 0.6118 1.0000
5.000 0.7058 0.02097 0.01169 -0.0361 0.5763 1.0000
5.250 0.7286 0.02110 0.01181 -0.0343 0.5341 1.0000
5.500 0.7500 0.02139 0.01204 -0.0325 0.4853 1.0000
5.750 0.7695 0.02190 0.01235 -0.0306 0.4333 1.0000
6.000 0.7874 0.02265 0.01285 -0.0287 0.3822 1.0000
6.250 0.8042 0.02354 0.01356 -0.0270 0.3339 1.0000
6.500 0.8203 0.02455 0.01439 -0.0255 0.2907 1.0000
6.750 0.8360 0.02569 0.01538 -0.0239 0.2504 1.0000
7.000 0.8513 0.02691 0.01656 -0.0225 0.2119 1.0000
7.250 0.8665 0.02819 0.01776 -0.0211 0.1753 1.0000
7.500 0.8828 0.02962 0.01912 -0.0197 0.1496 1.0000
7.750 0.8991 0.03111 0.02052 -0.0185 0.1301 1.0000
8.000 0.9166 0.03260 0.02204 -0.0174 0.1136 1.0000
8.250 0.9359 0.03424 0.02370 -0.0164 0.1034 1.0000
8.500 0.9573 0.03596 0.02557 -0.0155 0.0948 1.0000
8.750 0.9771 0.03778 0.02760 -0.0146 0.0859 1.0000
9.000 0.9956 0.03972 0.02959 -0.0138 0.0788 1.0000
9.250 1.0171 0.04225 0.03260 -0.0130 0.0735 1.0000
9.500 1.0337 0.04455 0.03500 -0.0121 0.0679 1.0000
9.750 1.0427 0.04730 0.03820 -0.0105 0.0618 1.0000
10.000 1.0521 0.05034 0.04157 -0.0090 0.0576 1.0000
10.250 1.0588 0.05337 0.04468 -0.0078 0.0535 1.0000
10.500 1.0502 0.05663 0.04851 -0.0054 0.0501 1.0000
10.750 1.0423 0.05932 0.05145 -0.0033 0.0468 1.0000
11.000 1.0396 0.06160 0.05374 -0.0020 0.0436 1.0000
11.250 1.0300 0.06522 0.05748 -0.0011 0.0419 1.0000
11.500 1.0119 0.06966 0.06225 -0.0009 0.0414 1.0000
11.750 0.9928 0.07467 0.06753 -0.0019 0.0412 1.0000
12.000 0.9732 0.08036 0.07343 -0.0040 0.0413 1.0000
12.250 0.9535 0.08670 0.07994 -0.0072 0.0415 1.0000
12.500 0.9356 0.09352 0.08687 -0.0110 0.0420 1.0000
12.750 0.9157 0.10145 0.09488 -0.0159 0.0423 1.0000
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Polar data table (+)
Polar graphs
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