HQ 2.5/10 AIRFOIL (hq2510-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: HQ 2.5/10 AIRFOIL (hq2510-il) Reynolds number: 200,000 Max Cl/Cd: 79.06 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-hq2510-il-200000.txt Download as CSV file: xf-hq2510-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/10 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3948 0.08891 0.08552 -0.0372 1.0000 0.0436
-8.500 -0.3984 0.08520 0.08187 -0.0389 1.0000 0.0450
-8.250 -0.4049 0.08156 0.07831 -0.0409 1.0000 0.0461
-8.000 -0.4193 0.07789 0.07474 -0.0431 1.0000 0.0468
-7.750 -0.4384 0.07428 0.07124 -0.0448 1.0000 0.0470
-7.500 -0.4562 0.07136 0.06832 -0.0447 1.0000 0.0471
-6.750 -0.4338 0.04646 0.04341 -0.0498 0.9905 0.0500
-6.500 -0.4136 0.04196 0.03888 -0.0525 0.9861 0.0516
-6.250 -0.3902 0.03682 0.03358 -0.0570 0.9823 0.0540
-6.000 -0.3715 0.03209 0.02852 -0.0604 0.9754 0.0580
-5.750 -0.3861 0.04533 0.04067 -0.0599 0.9793 0.0614
-5.500 -0.3646 0.03156 0.02637 -0.0615 0.9751 0.0322
-5.250 -0.3316 0.02530 0.01914 -0.0630 0.9716 0.0280
-5.000 -0.2943 0.02272 0.01614 -0.0651 0.9690 0.0290
-4.750 -0.2647 0.02159 0.01477 -0.0657 0.9628 0.0325
-4.500 -0.2286 0.01986 0.01269 -0.0669 0.9591 0.0350
-4.250 -0.1923 0.01766 0.01032 -0.0685 0.9565 0.0387
-4.000 -0.1594 0.01695 0.00953 -0.0695 0.9515 0.0489
-3.750 -0.1274 0.01563 0.00827 -0.0706 0.9465 0.0653
-3.500 -0.0885 0.01505 0.00765 -0.0731 0.9434 0.0866
-3.250 -0.0480 0.01418 0.00687 -0.0759 0.9412 0.1106
-3.000 -0.0223 0.01338 0.00631 -0.0758 0.9333 0.1587
-2.750 0.0064 0.01156 0.00611 -0.0770 0.9295 0.5521
-2.500 0.0354 0.01148 0.00619 -0.0768 0.9237 0.6422
-2.250 0.0660 0.01145 0.00615 -0.0768 0.9179 0.6855
-2.000 0.1010 0.01139 0.00607 -0.0777 0.9143 0.7194
-1.750 0.1257 0.01143 0.00609 -0.0766 0.9061 0.7452
-1.500 0.1569 0.01134 0.00599 -0.0766 0.9013 0.7709
-1.250 0.1796 0.01135 0.00604 -0.0748 0.8933 0.7998
-1.000 0.2064 0.01122 0.00593 -0.0737 0.8876 0.8244
-0.750 0.2299 0.01116 0.00586 -0.0724 0.8794 0.8437
-0.500 0.2577 0.01098 0.00565 -0.0718 0.8735 0.8590
-0.250 0.2819 0.01090 0.00557 -0.0708 0.8652 0.8730
0.000 0.3101 0.01073 0.00537 -0.0704 0.8592 0.8871
0.250 0.3344 0.01061 0.00527 -0.0693 0.8497 0.9033
0.500 0.3636 0.01040 0.00504 -0.0690 0.8406 0.9206
0.750 0.3978 0.01018 0.00478 -0.0697 0.8312 0.9404
1.000 0.4377 0.01007 0.00467 -0.0721 0.8212 0.9615
1.250 0.4819 0.00995 0.00453 -0.0755 0.8127 0.9850
1.500 0.5152 0.00988 0.00440 -0.0769 0.8033 1.0000
1.750 0.5393 0.00990 0.00438 -0.0764 0.7917 1.0000
2.000 0.5655 0.00992 0.00434 -0.0761 0.7801 1.0000
2.250 0.5923 0.00992 0.00433 -0.0758 0.7681 1.0000
2.500 0.6191 0.00994 0.00430 -0.0754 0.7557 1.0000
2.750 0.6458 0.00996 0.00429 -0.0751 0.7425 1.0000
3.000 0.6722 0.00999 0.00430 -0.0746 0.7281 1.0000
3.250 0.6983 0.01003 0.00434 -0.0741 0.7120 1.0000
3.500 0.7244 0.01007 0.00435 -0.0735 0.6944 1.0000
3.750 0.7501 0.01014 0.00439 -0.0728 0.6750 1.0000
4.000 0.7755 0.01023 0.00446 -0.0721 0.6531 1.0000
4.250 0.8004 0.01035 0.00458 -0.0714 0.6283 1.0000
4.500 0.8248 0.01051 0.00469 -0.0705 0.5992 1.0000
4.750 0.8483 0.01073 0.00482 -0.0695 0.5640 1.0000
5.000 0.8704 0.01106 0.00500 -0.0683 0.5195 1.0000
5.250 0.8910 0.01155 0.00526 -0.0670 0.4699 1.0000
5.500 0.9109 0.01213 0.00565 -0.0656 0.4194 1.0000
5.750 0.9304 0.01278 0.00609 -0.0642 0.3740 1.0000
6.000 0.9506 0.01342 0.00657 -0.0630 0.3352 1.0000
6.250 0.9704 0.01411 0.00711 -0.0618 0.3032 1.0000
6.500 0.9906 0.01479 0.00768 -0.0607 0.2741 1.0000
6.750 1.0107 0.01544 0.00824 -0.0596 0.2445 1.0000
7.000 1.0308 0.01605 0.00880 -0.0586 0.2139 1.0000
7.250 1.0508 0.01667 0.00934 -0.0576 0.1867 1.0000
7.500 1.0705 0.01731 0.00990 -0.0565 0.1592 1.0000
7.750 1.0898 0.01800 0.01047 -0.0555 0.1299 1.0000
8.000 1.1080 0.01882 0.01114 -0.0543 0.0981 1.0000
8.250 1.1141 0.02098 0.01269 -0.0514 0.0304 1.0000
8.500 1.1271 0.02237 0.01415 -0.0491 0.0230 1.0000
8.750 1.1354 0.02401 0.01594 -0.0463 0.0201 1.0000
9.000 1.1457 0.02530 0.01742 -0.0438 0.0188 1.0000
9.250 1.1555 0.02652 0.01876 -0.0413 0.0174 1.0000
9.500 1.1646 0.02779 0.02013 -0.0389 0.0161 1.0000
9.750 1.1719 0.02925 0.02168 -0.0364 0.0149 1.0000
10.000 1.1770 0.03109 0.02363 -0.0339 0.0144 1.0000
10.250 1.1834 0.03311 0.02577 -0.0316 0.0140 1.0000
10.500 1.1903 0.03571 0.02850 -0.0296 0.0136 1.0000
10.750 1.2004 0.03834 0.03132 -0.0280 0.0134 1.0000
11.000 1.2094 0.04124 0.03445 -0.0264 0.0133 1.0000
11.250 1.2147 0.04438 0.03784 -0.0247 0.0133 1.0000
11.500 1.2180 0.04691 0.04061 -0.0230 0.0133 1.0000
11.750 1.2154 0.05022 0.04418 -0.0212 0.0134 1.0000
12.000 1.2121 0.05316 0.04737 -0.0197 0.0135 1.0000
12.250 1.2061 0.05639 0.05084 -0.0186 0.0135 1.0000
12.500 1.1988 0.05972 0.05442 -0.0178 0.0137 1.0000
12.750 1.1877 0.06372 0.05869 -0.0177 0.0139 1.0000
13.000 1.1727 0.06849 0.06374 -0.0183 0.0141 1.0000
13.250 1.1545 0.07409 0.06963 -0.0198 0.0144 1.0000
13.500 1.1345 0.08041 0.07621 -0.0223 0.0145 1.0000
13.750 1.1127 0.08771 0.08376 -0.0260 0.0149 1.0000
14.000 1.0905 0.09580 0.09207 -0.0309 0.0151 1.0000
14.250 1.0663 0.10527 0.10175 -0.0372 0.0152 1.0000
14.500 1.0396 0.11657 0.11325 -0.0450 0.0155 1.0000
14.750 1.0088 0.13060 0.12743 -0.0544 0.0158 1.0000
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Polar data table (+)
Polar graphs
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