HQ 2.1/9.5 AIRFOIL (hq2195-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 2.1/9.5 AIRFOIL (hq2195-il) Reynolds number: 100,000 Max Cl/Cd: 52.65 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq2195-il-100000-n5.txt Download as CSV file: xf-hq2195-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.1/9.5 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4878 0.07648 0.07142 -0.0455 1.0000 0.0330
-8.750 -0.4970 0.07182 0.06680 -0.0479 1.0000 0.0327
-8.500 -0.5077 0.06769 0.06269 -0.0498 1.0000 0.0324
-8.250 -0.5164 0.06350 0.05845 -0.0509 1.0000 0.0321
-8.000 -0.5242 0.05936 0.05423 -0.0514 1.0000 0.0319
-7.750 -0.5296 0.05537 0.05011 -0.0513 1.0000 0.0317
-7.500 -0.5329 0.05153 0.04608 -0.0506 1.0000 0.0315
-7.250 -0.5339 0.04784 0.04216 -0.0495 1.0000 0.0315
-7.000 -0.5323 0.04431 0.03833 -0.0482 1.0000 0.0315
-6.750 -0.5275 0.04090 0.03451 -0.0468 1.0000 0.0321
-6.500 -0.5186 0.03807 0.03108 -0.0453 1.0000 0.0332
-6.250 -0.4926 0.03436 0.02708 -0.0476 0.9955 0.0343
-6.000 -0.4632 0.03148 0.02378 -0.0494 0.9906 0.0348
-5.750 -0.4322 0.02908 0.02099 -0.0510 0.9858 0.0356
-5.500 -0.4002 0.02703 0.01858 -0.0526 0.9816 0.0366
-5.250 -0.3698 0.02560 0.01689 -0.0536 0.9763 0.0390
-5.000 -0.3372 0.02413 0.01509 -0.0548 0.9722 0.0414
-4.750 -0.3041 0.02270 0.01339 -0.0559 0.9688 0.0427
-4.500 -0.2755 0.02142 0.01197 -0.0562 0.9634 0.0441
-4.250 -0.2443 0.02028 0.01085 -0.0573 0.9592 0.0465
-4.000 -0.2103 0.01951 0.01005 -0.0588 0.9559 0.0513
-3.750 -0.1824 0.01885 0.00930 -0.0590 0.9502 0.0558
-3.500 -0.1520 0.01805 0.00850 -0.0599 0.9457 0.0616
-3.250 -0.1181 0.01741 0.00784 -0.0614 0.9423 0.0733
-3.000 -0.0881 0.01683 0.00729 -0.0621 0.9373 0.0950
-2.750 -0.0596 0.01599 0.00680 -0.0628 0.9322 0.1616
-2.500 -0.0335 0.01439 0.00668 -0.0636 0.9284 0.4958
-2.250 -0.0068 0.01421 0.00686 -0.0629 0.9239 0.6174
-2.000 0.0173 0.01424 0.00698 -0.0617 0.9178 0.6799
-1.750 0.0436 0.01428 0.00711 -0.0605 0.9135 0.7373
-1.500 0.0664 0.01434 0.00724 -0.0586 0.9083 0.7814
-1.250 0.0899 0.01437 0.00726 -0.0573 0.9022 0.8077
-1.000 0.1216 0.01432 0.00712 -0.0578 0.8979 0.8231
-0.750 0.1499 0.01423 0.00696 -0.0575 0.8888 0.8374
-0.500 0.1851 0.01399 0.00665 -0.0582 0.8800 0.8511
-0.250 0.2145 0.01380 0.00640 -0.0579 0.8685 0.8657
0.000 0.2420 0.01368 0.00626 -0.0574 0.8588 0.8816
0.250 0.2742 0.01354 0.00609 -0.0579 0.8522 0.8979
0.500 0.3040 0.01346 0.00603 -0.0581 0.8433 0.9164
0.750 0.3423 0.01332 0.00587 -0.0598 0.8368 0.9352
1.000 0.3804 0.01326 0.00583 -0.0619 0.8273 0.9596
1.250 0.4205 0.01316 0.00572 -0.0644 0.8187 1.0000
1.500 0.4475 0.01317 0.00570 -0.0642 0.8085 1.0000
1.750 0.4736 0.01322 0.00575 -0.0640 0.7968 1.0000
2.000 0.5008 0.01324 0.00576 -0.0637 0.7852 1.0000
2.250 0.5283 0.01326 0.00578 -0.0635 0.7727 1.0000
2.500 0.5554 0.01329 0.00582 -0.0631 0.7588 1.0000
2.750 0.5820 0.01332 0.00587 -0.0626 0.7430 1.0000
3.000 0.6084 0.01336 0.00593 -0.0621 0.7250 1.0000
3.250 0.6350 0.01339 0.00597 -0.0615 0.7046 1.0000
3.500 0.6602 0.01345 0.00607 -0.0606 0.6786 1.0000
3.750 0.6853 0.01351 0.00609 -0.0597 0.6441 1.0000
4.000 0.7094 0.01361 0.00606 -0.0585 0.5906 1.0000
4.250 0.7323 0.01391 0.00602 -0.0570 0.5255 1.0000
4.500 0.7538 0.01447 0.00623 -0.0556 0.4699 1.0000
4.750 0.7758 0.01504 0.00659 -0.0546 0.4266 1.0000
5.000 0.7983 0.01560 0.00700 -0.0537 0.3913 1.0000
5.250 0.8210 0.01614 0.00748 -0.0529 0.3587 1.0000
5.500 0.8434 0.01669 0.00795 -0.0521 0.3253 1.0000
5.750 0.8654 0.01729 0.00845 -0.0512 0.2909 1.0000
6.000 0.8868 0.01793 0.00899 -0.0503 0.2483 1.0000
6.250 0.9066 0.01877 0.00962 -0.0493 0.1963 1.0000
6.500 0.9252 0.01982 0.01037 -0.0482 0.1505 1.0000
6.750 0.9440 0.02091 0.01124 -0.0472 0.1195 1.0000
7.000 0.9636 0.02189 0.01216 -0.0462 0.0994 1.0000
7.250 0.9828 0.02289 0.01315 -0.0452 0.0856 1.0000
7.500 1.0012 0.02396 0.01427 -0.0440 0.0754 1.0000
7.750 1.0182 0.02513 0.01541 -0.0427 0.0668 1.0000
8.000 1.0367 0.02617 0.01658 -0.0415 0.0595 1.0000
8.250 1.0517 0.02755 0.01794 -0.0400 0.0549 1.0000
8.500 1.0696 0.02880 0.01938 -0.0386 0.0513 1.0000
8.750 1.0870 0.03012 0.02082 -0.0374 0.0479 1.0000
9.000 1.1029 0.03151 0.02223 -0.0362 0.0449 1.0000
9.250 1.1203 0.03306 0.02393 -0.0350 0.0423 1.0000
9.500 1.1385 0.03472 0.02581 -0.0339 0.0401 1.0000
9.750 1.1557 0.03648 0.02776 -0.0328 0.0384 1.0000
10.000 1.1707 0.03812 0.02956 -0.0316 0.0366 1.0000
10.250 1.1838 0.03991 0.03139 -0.0304 0.0348 1.0000
10.500 1.1954 0.04214 0.03390 -0.0289 0.0333 1.0000
10.750 1.2041 0.04458 0.03669 -0.0271 0.0323 1.0000
11.000 1.2084 0.04714 0.03958 -0.0249 0.0313 1.0000
11.250 1.2086 0.04963 0.04236 -0.0227 0.0303 1.0000
11.500 1.2062 0.05211 0.04509 -0.0205 0.0294 1.0000
11.750 1.2024 0.05455 0.04774 -0.0188 0.0285 1.0000
12.000 1.1979 0.05715 0.05051 -0.0175 0.0278 1.0000
12.250 1.1927 0.06004 0.05354 -0.0166 0.0273 1.0000
12.500 1.1849 0.06351 0.05717 -0.0161 0.0268 1.0000
12.750 1.1725 0.06780 0.06166 -0.0162 0.0265 1.0000
13.000 1.1556 0.07283 0.06691 -0.0172 0.0264 1.0000
13.250 1.1352 0.07865 0.07299 -0.0192 0.0263 1.0000
13.500 1.1140 0.08520 0.07977 -0.0224 0.0263 1.0000
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Polar data table (+)
Polar graphs
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