HQ 1.5/9 AIRFOIL (hq159-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 1.5/9 AIRFOIL (hq159-il) Reynolds number: 200,000 Max Cl/Cd: 61.74 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq159-il-200000-n5.txt Download as CSV file: xf-hq159-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5302 0.08484 0.08137 -0.0232 1.0000 0.0094
-9.250 -0.5334 0.07906 0.07564 -0.0264 1.0000 0.0091
-8.750 -0.5477 0.06556 0.06220 -0.0367 1.0000 0.0088
-8.500 -0.5625 0.05968 0.05628 -0.0411 1.0000 0.0087
-8.250 -0.5741 0.05374 0.05020 -0.0435 1.0000 0.0087
-8.000 -0.5824 0.04793 0.04416 -0.0445 1.0000 0.0087
-7.750 -0.5850 0.04301 0.03897 -0.0444 1.0000 0.0086
-7.500 -0.5853 0.03788 0.03342 -0.0435 1.0000 0.0086
-7.250 -0.5808 0.03352 0.02859 -0.0420 1.0000 0.0087
-7.000 -0.5712 0.03014 0.02472 -0.0403 1.0000 0.0089
-6.750 -0.5619 0.02656 0.02066 -0.0385 1.0000 0.0092
-6.500 -0.5476 0.02508 0.01903 -0.0371 1.0000 0.0102
-6.250 -0.5261 0.02400 0.01776 -0.0367 0.9984 0.0114
-6.000 -0.4942 0.02190 0.01529 -0.0381 0.9935 0.0125
-5.750 -0.4630 0.01977 0.01279 -0.0390 0.9884 0.0132
-5.500 -0.4300 0.01810 0.01086 -0.0403 0.9843 0.0142
-5.250 -0.3996 0.01664 0.00921 -0.0410 0.9783 0.0153
-5.000 -0.3664 0.01572 0.00821 -0.0426 0.9737 0.0186
-4.750 -0.3347 0.01490 0.00727 -0.0436 0.9677 0.0217
-4.500 -0.3030 0.01393 0.00619 -0.0446 0.9616 0.0264
-4.250 -0.2707 0.01331 0.00551 -0.0457 0.9555 0.0364
-4.000 -0.2396 0.01280 0.00504 -0.0466 0.9484 0.0535
-3.750 -0.2079 0.01233 0.00461 -0.0477 0.9419 0.0766
-3.500 -0.1778 0.01183 0.00416 -0.0484 0.9343 0.1061
-3.250 -0.1484 0.01123 0.00380 -0.0490 0.9268 0.1711
-3.000 -0.1202 0.01047 0.00346 -0.0496 0.9193 0.2871
-2.500 -0.0671 0.00936 0.00317 -0.0494 0.9038 0.5384
-2.250 -0.0412 0.00919 0.00315 -0.0488 0.8947 0.6023
-2.000 -0.0140 0.00909 0.00311 -0.0484 0.8868 0.6465
-1.750 0.0122 0.00903 0.00309 -0.0477 0.8783 0.6858
-1.500 0.0382 0.00898 0.00306 -0.0469 0.8692 0.7159
-1.250 0.0651 0.00896 0.00300 -0.0464 0.8607 0.7370
-1.000 0.0910 0.00892 0.00296 -0.0457 0.8511 0.7563
-0.750 0.1169 0.00890 0.00293 -0.0450 0.8414 0.7751
-0.500 0.1434 0.00887 0.00288 -0.0445 0.8328 0.7903
-0.250 0.1699 0.00884 0.00283 -0.0440 0.8231 0.8018
0.000 0.1962 0.00881 0.00278 -0.0434 0.8116 0.8136
0.250 0.2222 0.00878 0.00273 -0.0428 0.7986 0.8260
0.500 0.2482 0.00875 0.00268 -0.0421 0.7850 0.8391
0.750 0.2743 0.00871 0.00264 -0.0414 0.7712 0.8533
1.000 0.3008 0.00868 0.00260 -0.0409 0.7568 0.8691
1.250 0.3282 0.00866 0.00259 -0.0405 0.7427 0.8872
1.500 0.3578 0.00864 0.00260 -0.0406 0.7294 0.9076
1.750 0.3916 0.00864 0.00262 -0.0417 0.7153 0.9320
2.000 0.4285 0.00866 0.00263 -0.0436 0.6996 0.9628
2.250 0.4643 0.00871 0.00268 -0.0453 0.6821 1.0000
2.500 0.4898 0.00882 0.00275 -0.0448 0.6632 1.0000
2.750 0.5153 0.00895 0.00282 -0.0443 0.6423 1.0000
3.000 0.5407 0.00909 0.00291 -0.0438 0.6171 1.0000
3.250 0.5656 0.00928 0.00301 -0.0432 0.5843 1.0000
3.500 0.5896 0.00955 0.00315 -0.0423 0.5397 1.0000
3.750 0.6125 0.00993 0.00330 -0.0414 0.4834 1.0000
4.000 0.6344 0.01045 0.00353 -0.0404 0.4194 1.0000
4.250 0.6561 0.01105 0.00383 -0.0395 0.3535 1.0000
4.750 0.7014 0.01218 0.00458 -0.0381 0.2605 1.0000
5.000 0.7255 0.01263 0.00497 -0.0377 0.2336 1.0000
5.250 0.7493 0.01310 0.00537 -0.0372 0.2037 1.0000
5.500 0.7726 0.01365 0.00579 -0.0366 0.1647 1.0000
5.750 0.7949 0.01433 0.00628 -0.0360 0.1231 1.0000
6.000 0.8169 0.01506 0.00686 -0.0353 0.0918 1.0000
6.250 0.8391 0.01579 0.00753 -0.0346 0.0694 1.0000
6.500 0.8614 0.01648 0.00816 -0.0340 0.0544 1.0000
6.750 0.8844 0.01708 0.00881 -0.0333 0.0454 1.0000
7.000 0.9072 0.01771 0.00953 -0.0326 0.0401 1.0000
7.250 0.9293 0.01840 0.01026 -0.0319 0.0330 1.0000
7.500 0.9518 0.01901 0.01090 -0.0313 0.0234 1.0000
7.750 0.9729 0.01982 0.01173 -0.0305 0.0158 1.0000
8.000 0.9922 0.02089 0.01294 -0.0294 0.0116 1.0000
8.250 1.0115 0.02194 0.01412 -0.0282 0.0095 1.0000
8.500 1.0286 0.02324 0.01554 -0.0269 0.0083 1.0000
8.750 1.0441 0.02471 0.01718 -0.0253 0.0077 1.0000
9.000 1.0596 0.02615 0.01883 -0.0238 0.0074 1.0000
9.250 1.0737 0.02778 0.02067 -0.0221 0.0071 1.0000
9.500 1.0866 0.02955 0.02266 -0.0204 0.0068 1.0000
9.750 1.0980 0.03151 0.02486 -0.0186 0.0067 1.0000
10.000 1.1074 0.03364 0.02725 -0.0166 0.0065 1.0000
10.250 1.1139 0.03586 0.02974 -0.0145 0.0063 1.0000
10.500 1.1166 0.03797 0.03209 -0.0120 0.0061 1.0000
10.750 1.1189 0.03972 0.03404 -0.0098 0.0057 1.0000
11.000 1.1164 0.04214 0.03669 -0.0076 0.0056 1.0000
11.250 1.1152 0.04418 0.03895 -0.0062 0.0053 1.0000
11.500 1.1130 0.04630 0.04118 -0.0052 0.0050 1.0000
11.750 1.1032 0.04994 0.04508 -0.0046 0.0050 1.0000
12.000 1.0924 0.05377 0.04910 -0.0048 0.0049 1.0000
12.250 1.0789 0.05844 0.05400 -0.0060 0.0049 1.0000
12.500 1.0601 0.06436 0.06013 -0.0086 0.0047 1.0000
12.750 1.0472 0.07027 0.06627 -0.0117 0.0049 1.0000
13.000 1.0299 0.07758 0.07377 -0.0163 0.0049 1.0000
13.250 1.0110 0.08624 0.08262 -0.0222 0.0049 1.0000
13.500 0.9209 0.11653 0.11331 -0.0402 0.0067 1.0000
13.750 0.8898 0.13150 0.12829 -0.0477 0.0070 1.0000
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Polar data table (+)
Polar graphs
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