NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 1,000,000 Max Cl/Cd: 99.67 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n6h10-il-1000000-n5.txt Download as CSV file: xf-n6h10-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 6-H-10 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.4168 0.08750 0.08471 0.0026 0.5805 0.0039
-7.250 -0.4162 0.08430 0.08151 0.0014 0.5791 0.0039
-7.000 -0.4105 0.08089 0.07808 -0.0002 0.5776 0.0039
-6.750 -0.4032 0.07725 0.07441 -0.0018 0.5760 0.0039
-6.500 -0.3937 0.07369 0.07082 -0.0034 0.5744 0.0039
-6.250 -0.3824 0.07008 0.06716 -0.0047 0.5727 0.0039
-6.000 -0.3747 0.06596 0.06299 -0.0055 0.5711 0.0040
-5.750 -0.3610 0.06257 0.05954 -0.0064 0.5693 0.0041
-5.500 -0.3462 0.06014 0.05705 -0.0071 0.5673 0.0043
-5.250 -0.3295 0.05711 0.05395 -0.0078 0.5658 0.0046
-5.000 -0.3115 0.05390 0.05067 -0.0083 0.5639 0.0049
-4.750 -0.2922 0.05065 0.04731 -0.0086 0.5621 0.0053
-4.500 -0.2711 0.04734 0.04389 -0.0087 0.5603 0.0056
-4.250 -0.2475 0.04400 0.04040 -0.0084 0.5583 0.0059
-4.000 -0.2253 0.04086 0.03712 -0.0078 0.5567 0.0060
-3.750 -0.2035 0.03786 0.03397 -0.0071 0.5552 0.0060
-3.250 -0.1641 0.02979 0.02545 -0.0040 0.5525 0.0038
-3.000 -0.1415 0.02665 0.02209 -0.0024 0.5512 0.0037
-2.750 -0.1182 0.02318 0.01835 -0.0005 0.5496 0.0038
-2.500 -0.0952 0.01902 0.01379 0.0019 0.5480 0.0041
-2.250 -0.0699 0.01691 0.01140 0.0030 0.5464 0.0042
-2.000 -0.0435 0.01522 0.00947 0.0037 0.5447 0.0044
-1.750 -0.0163 0.01430 0.00839 0.0040 0.5430 0.0048
-1.500 0.0117 0.01282 0.00666 0.0044 0.5413 0.0051
-1.250 0.0402 0.01131 0.00491 0.0048 0.5396 0.0052
-1.000 0.0677 0.01039 0.00385 0.0052 0.5382 0.0055
-0.750 0.0945 0.00981 0.00319 0.0055 0.5367 0.0058
-0.500 0.1214 0.00947 0.00280 0.0057 0.5351 0.0061
-0.250 0.1474 0.00905 0.00235 0.0061 0.5336 0.0069
0.000 0.1747 0.00891 0.00222 0.0061 0.5319 0.0078
0.250 0.2017 0.00873 0.00202 0.0063 0.5302 0.0089
0.500 0.2288 0.00858 0.00183 0.0064 0.5285 0.0095
0.750 0.2550 0.00835 0.00154 0.0067 0.5269 0.0119
1.000 0.2823 0.00827 0.00145 0.0068 0.5255 0.0149
1.250 0.3097 0.00821 0.00136 0.0068 0.5242 0.0175
1.500 0.3364 0.00806 0.00131 0.0069 0.5231 0.0543
1.750 0.3619 0.00784 0.00133 0.0072 0.5219 0.1493
2.000 0.3646 0.00612 0.00116 0.0118 0.5207 0.7427
2.250 0.3828 0.00589 0.00133 0.0141 0.5193 0.8808
2.500 0.4087 0.00593 0.00142 0.0146 0.5179 0.8984
2.750 0.4368 0.00597 0.00145 0.0144 0.5160 0.9020
3.000 0.4649 0.00601 0.00147 0.0142 0.5135 0.9047
3.250 0.4928 0.00607 0.00150 0.0140 0.5111 0.9072
3.500 0.5209 0.00610 0.00153 0.0138 0.5054 0.9096
3.750 0.5488 0.00614 0.00154 0.0136 0.4924 0.9122
4.000 0.5768 0.00622 0.00158 0.0134 0.4809 0.9141
4.250 0.6047 0.00630 0.00165 0.0131 0.4691 0.9163
4.500 0.6326 0.00640 0.00172 0.0128 0.4563 0.9172
4.750 0.6598 0.00662 0.00184 0.0125 0.4271 0.9178
5.000 0.6815 0.00767 0.00246 0.0125 0.3296 0.9188
5.250 0.7074 0.00804 0.00274 0.0123 0.3072 0.9196
5.500 0.7328 0.00844 0.00302 0.0122 0.2773 0.9205
5.750 0.7508 0.00967 0.00382 0.0125 0.1782 0.9220
6.000 0.7667 0.01091 0.00464 0.0132 0.0850 0.9237
6.250 0.7903 0.01130 0.00500 0.0133 0.0705 0.9250
6.500 0.8140 0.01165 0.00534 0.0133 0.0624 0.9264
6.750 0.8379 0.01197 0.00567 0.0134 0.0572 0.9279
7.000 0.8612 0.01229 0.00601 0.0136 0.0535 0.9296
7.250 0.8835 0.01267 0.00640 0.0138 0.0497 0.9316
7.500 0.9052 0.01307 0.00682 0.0141 0.0464 0.9340
7.750 0.9267 0.01343 0.00725 0.0145 0.0457 0.9366
8.000 0.9473 0.01385 0.00771 0.0148 0.0435 0.9398
8.250 0.9666 0.01441 0.00829 0.0151 0.0405 0.9442
8.500 0.9824 0.01530 0.00923 0.0151 0.0373 0.9517
8.750 1.0044 0.01611 0.01011 0.0142 0.0355 0.9587
9.000 1.0298 0.01694 0.01101 0.0125 0.0328 0.9651
9.250 1.0576 0.01805 0.01211 0.0097 0.0263 0.9699
9.500 1.0804 0.01995 0.01391 0.0065 0.0088 0.9760
9.750 1.1038 0.02144 0.01544 0.0039 0.0060 0.9805
10.000 1.1239 0.02285 0.01690 0.0023 0.0046 0.9870
10.250 1.1426 0.02418 0.01828 0.0011 0.0038 0.9932
10.500 1.1567 0.02571 0.01986 0.0006 0.0032 1.0000
10.750 1.1632 0.02686 0.02106 0.0022 0.0029 1.0000
11.000 1.1696 0.02811 0.02238 0.0036 0.0025 1.0000
11.250 1.1755 0.02947 0.02380 0.0049 0.0025 1.0000
11.500 1.1804 0.03094 0.02530 0.0061 0.0020 1.0000
11.750 1.1860 0.03240 0.02683 0.0073 0.0020 1.0000
12.000 1.1952 0.03364 0.02812 0.0081 0.0019 1.0000
12.250 1.2021 0.03507 0.02962 0.0091 0.0018 1.0000
12.500 1.2094 0.03649 0.03109 0.0099 0.0016 1.0000
12.750 1.2148 0.03817 0.03285 0.0107 0.0016 1.0000
13.000 1.2213 0.03980 0.03455 0.0114 0.0015 1.0000
13.250 1.2284 0.04142 0.03623 0.0120 0.0014 1.0000
13.500 1.2343 0.04315 0.03802 0.0126 0.0013 1.0000
13.750 1.2392 0.04508 0.04003 0.0131 0.0012 1.0000
14.000 1.2392 0.04750 0.04255 0.0138 0.0011 1.0000
14.250 1.2486 0.04905 0.04418 0.0141 0.0013 1.0000
14.500 1.2475 0.05170 0.04692 0.0146 0.0011 1.0000
14.750 1.2512 0.05392 0.04923 0.0148 0.0011 1.0000
15.000 1.2536 0.05629 0.05172 0.0150 0.0011 1.0000
15.250 1.2608 0.05819 0.05371 0.0150 0.0010 1.0000
15.500 1.2582 0.06125 0.05688 0.0151 0.0010 1.0000
15.750 1.2535 0.06466 0.06042 0.0151 0.0010 1.0000
16.000 1.2595 0.06682 0.06266 0.0149 0.0010 1.0000
16.250 1.2641 0.06919 0.06512 0.0146 0.0008 1.0000
16.500 1.2575 0.07306 0.06912 0.0142 0.0008 1.0000
16.750 1.2559 0.07634 0.07251 0.0137 0.0008 1.0000
17.000 1.2586 0.07911 0.07535 0.0131 0.0007 1.0000
17.250 1.2458 0.08414 0.08054 0.0122 0.0007 1.0000
17.500 1.2460 0.08741 0.08390 0.0113 0.0007 1.0000
17.750 1.2351 0.09242 0.08905 0.0100 0.0007 1.0000
18.000 1.2266 0.09717 0.09393 0.0086 0.0007 1.0000
18.250 1.2138 0.10282 0.09970 0.0068 0.0007 1.0000
18.500 1.2025 0.10839 0.10540 0.0048 0.0007 1.0000
18.750 1.1831 0.11553 0.11271 0.0020 0.0007 1.0000
19.000 1.1730 0.12121 0.11850 -0.0004 0.0007 1.0000
19.250 1.1603 0.12757 0.12497 -0.0032 0.0007 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA 6-H-10 AIRFOIL (n6h10-il)