NACA 11-H-09 AIRFOIL (n11h9-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 11-H-09 AIRFOIL (n11h9-il) Reynolds number: 50,000 Max Cl/Cd: 24.57 at α=10° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n11h9-il-50000-n5.txt Download as CSV file: xf-n11h9-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 11-H-09 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3311 0.10641 0.10046 -0.0194 0.8770 0.0656
-8.500 -0.3281 0.10417 0.09826 -0.0223 0.8678 0.0674
-8.250 -0.3284 0.10240 0.09654 -0.0262 0.8585 0.0683
-8.000 -0.3306 0.10093 0.09509 -0.0296 0.8498 0.0688
-7.750 -0.3179 0.09587 0.09002 -0.0286 0.8434 0.0705
-7.500 -0.3080 0.09234 0.08650 -0.0288 0.8364 0.0735
-7.250 -0.3023 0.08971 0.08387 -0.0304 0.8295 0.0765
-7.000 -0.2981 0.08793 0.08206 -0.0346 0.8217 0.0801
-6.750 -0.2932 0.08674 0.08075 -0.0384 0.8150 0.0812
-6.500 -0.2791 0.08115 0.07527 -0.0362 0.8092 0.0849
-6.250 -0.2694 0.07855 0.07261 -0.0375 0.8031 0.0899
-6.000 -0.2589 0.07736 0.07122 -0.0416 0.7965 0.0939
-5.750 -0.2469 0.07273 0.06668 -0.0402 0.7911 0.0981
-5.500 -0.2340 0.07253 0.06609 -0.0430 0.7856 0.1064
-5.250 -0.2214 0.06735 0.06113 -0.0419 0.7798 0.1118
-5.000 -0.2082 0.06497 0.05859 -0.0427 0.7745 0.1225
-4.750 -0.1947 0.06269 0.05612 -0.0429 0.7698 0.1357
-4.500 -0.1807 0.05990 0.05328 -0.0430 0.7641 0.1513
-4.250 -0.1679 0.05715 0.05050 -0.0420 0.7594 0.1702
-4.000 -0.1565 0.05472 0.04801 -0.0410 0.7549 0.1995
-3.750 -0.1467 0.05236 0.04563 -0.0398 0.7492 0.2409
-3.500 -0.1346 0.04980 0.04306 -0.0381 0.7449 0.2732
-3.250 -0.1175 0.04752 0.04064 -0.0372 0.7411 0.2903
-3.000 -0.0428 0.04571 0.03733 -0.0422 0.7358 0.0757
-2.750 -0.0181 0.04358 0.03479 -0.0410 0.7314 0.0617
-2.500 0.0067 0.04180 0.03256 -0.0395 0.7280 0.0542
-2.250 0.0297 0.04021 0.03065 -0.0390 0.7228 0.0529
-2.000 0.0536 0.03881 0.02889 -0.0382 0.7183 0.0535
-1.750 0.0785 0.03742 0.02712 -0.0371 0.7148 0.0538
-1.500 0.1042 0.03610 0.02548 -0.0363 0.7109 0.0530
-1.250 0.1311 0.03501 0.02407 -0.0361 0.7056 0.0524
-1.000 0.1601 0.03388 0.02261 -0.0358 0.7018 0.0525
-0.750 0.1927 0.03289 0.02118 -0.0358 0.6989 0.0560
-0.500 0.2355 0.03187 0.01995 -0.0389 0.6942 0.0606
-0.250 0.2688 0.03122 0.01911 -0.0399 0.6896 0.0641
0.000 0.2961 0.03071 0.01842 -0.0396 0.6861 0.0716
0.250 0.3213 0.03034 0.01791 -0.0387 0.6831 0.0817
0.500 0.3431 0.03055 0.01807 -0.0388 0.6768 0.0957
0.750 0.3666 0.03030 0.01785 -0.0382 0.6727 0.1326
1.000 0.4826 0.02800 0.01705 -0.0549 0.6706 1.0000
1.250 0.5039 0.02870 0.01756 -0.0545 0.6654 1.0000
1.500 0.5249 0.02936 0.01806 -0.0541 0.6598 1.0000
1.750 0.5472 0.02971 0.01826 -0.0530 0.6560 1.0000
2.000 0.5667 0.03056 0.01902 -0.0526 0.6498 1.0000
2.250 0.5871 0.03117 0.01955 -0.0519 0.6443 1.0000
2.500 0.6100 0.03142 0.01970 -0.0506 0.6407 1.0000
2.750 0.6264 0.03255 0.02082 -0.0503 0.6326 1.0000
3.000 0.6482 0.03292 0.02114 -0.0492 0.6279 1.0000
3.250 0.6660 0.03379 0.02201 -0.0485 0.6211 1.0000
3.500 0.6853 0.03443 0.02265 -0.0475 0.6149 1.0000
3.750 0.7094 0.03454 0.02274 -0.0462 0.6113 1.0000
4.000 0.7214 0.03595 0.02421 -0.0456 0.6015 1.0000
4.250 0.7455 0.03606 0.02438 -0.0443 0.5975 1.0000
4.500 0.7568 0.03746 0.02584 -0.0435 0.5877 1.0000
4.750 0.7812 0.03752 0.02595 -0.0422 0.5832 1.0000
5.000 0.7916 0.03893 0.02744 -0.0413 0.5730 1.0000
5.250 0.8173 0.03882 0.02745 -0.0399 0.5685 1.0000
5.500 0.8264 0.04027 0.02900 -0.0388 0.5575 1.0000
5.750 0.8547 0.03984 0.02867 -0.0374 0.5532 1.0000
6.000 0.8629 0.04130 0.03024 -0.0362 0.5414 1.0000
6.250 0.8745 0.04239 0.03145 -0.0349 0.5307 1.0000
6.500 0.9019 0.04187 0.03113 -0.0333 0.5251 1.0000
6.750 0.9107 0.04314 0.03254 -0.0319 0.5128 1.0000
7.250 0.9527 0.04302 0.03276 -0.0287 0.4956 1.0000
7.500 0.9610 0.04414 0.03408 -0.0271 0.4825 1.0000
7.750 0.9699 0.04513 0.03524 -0.0255 0.4695 1.0000
8.000 0.9792 0.04596 0.03622 -0.0236 0.4567 1.0000
8.250 0.9894 0.04650 0.03692 -0.0215 0.4440 1.0000
8.500 1.0020 0.04671 0.03733 -0.0195 0.4311 1.0000
8.750 1.0153 0.04675 0.03761 -0.0173 0.4169 1.0000
9.000 1.0296 0.04654 0.03762 -0.0151 0.4010 1.0000
9.250 1.0500 0.04556 0.03686 -0.0128 0.3833 1.0000
9.500 1.0647 0.04517 0.03663 -0.0105 0.3615 1.0000
9.750 1.0818 0.04444 0.03599 -0.0081 0.3333 1.0000
10.000 1.0921 0.04445 0.03593 -0.0056 0.2968 1.0000
10.250 1.0915 0.04599 0.03729 -0.0037 0.2599 1.0000
10.500 1.0872 0.04816 0.03921 -0.0021 0.2244 1.0000
10.750 1.0809 0.05081 0.04163 -0.0008 0.1934 1.0000
11.000 1.0735 0.05380 0.04441 0.0001 0.1662 1.0000
11.250 1.0671 0.05690 0.04738 0.0008 0.1430 1.0000
11.500 1.0606 0.06012 0.05042 0.0014 0.1253 1.0000
11.750 1.0555 0.06334 0.05350 0.0018 0.1100 1.0000
12.000 1.0517 0.06651 0.05659 0.0022 0.0982 1.0000
12.250 1.0510 0.06946 0.05956 0.0026 0.0878 1.0000
12.500 1.0527 0.07215 0.06229 0.0034 0.0804 1.0000
12.750 1.0574 0.07451 0.06471 0.0042 0.0751 1.0000
13.000 1.0652 0.07654 0.06684 0.0054 0.0708 1.0000
13.250 1.0737 0.07882 0.06940 0.0064 0.0664 1.0000
13.500 1.0793 0.08129 0.07197 0.0069 0.0624 1.0000
13.750 1.0854 0.08384 0.07462 0.0075 0.0588 1.0000
14.000 1.0863 0.08742 0.07856 0.0074 0.0565 1.0000
14.250 1.0864 0.09129 0.08273 0.0071 0.0548 1.0000
14.500 1.0829 0.09569 0.08740 0.0064 0.0536 1.0000
14.750 1.0756 0.10075 0.09272 0.0051 0.0528 1.0000
15.000 1.0641 0.10666 0.09886 0.0030 0.0522 1.0000
15.250 1.0478 0.11369 0.10612 0.0000 0.0522 1.0000
15.500 1.0261 0.12232 0.11494 -0.0044 0.0524 1.0000
15.750 0.9976 0.13340 0.12618 -0.0106 0.0532 1.0000
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Polar data table (+)
Polar graphs
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