NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 500,000 Max Cl/Cd: 87.93 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n6h10-il-500000-n5.txt Download as CSV file: xf-n6h10-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.4200 0.10587 0.10253 0.0093 0.6128 0.0061
-8.500 -0.4159 0.10253 0.09919 0.0075 0.6111 0.0061
-7.000 -0.4065 0.08177 0.07843 -0.0018 0.6005 0.0063
-6.750 -0.3996 0.07848 0.07512 -0.0022 0.5988 0.0065
-6.500 -0.3911 0.07555 0.07215 -0.0031 0.5972 0.0067
-6.250 -0.3809 0.07229 0.06883 -0.0044 0.5957 0.0069
-6.000 -0.3690 0.06891 0.06542 -0.0057 0.5941 0.0071
-5.750 -0.3554 0.06555 0.06200 -0.0069 0.5924 0.0072
-5.500 -0.3404 0.06218 0.05856 -0.0079 0.5907 0.0074
-5.250 -0.3239 0.05882 0.05512 -0.0087 0.5889 0.0076
-5.000 -0.3059 0.05555 0.05174 -0.0093 0.5873 0.0079
-4.750 -0.2862 0.05229 0.04836 -0.0097 0.5857 0.0083
-4.500 -0.2638 0.04918 0.04510 -0.0099 0.5841 0.0088
-4.250 -0.2385 0.04645 0.04218 -0.0097 0.5825 0.0092
-4.000 -0.2159 0.04365 0.03924 -0.0093 0.5809 0.0093
-2.750 -0.1119 0.02810 0.02274 -0.0044 0.5724 0.0077
-2.500 -0.0876 0.02448 0.01875 -0.0024 0.5709 0.0069
-2.250 -0.0624 0.02195 0.01592 -0.0010 0.5692 0.0070
-2.000 -0.0362 0.02054 0.01428 -0.0003 0.5677 0.0081
-1.750 -0.0095 0.01839 0.01183 0.0007 0.5660 0.0080
-1.500 0.0183 0.01642 0.00955 0.0015 0.5644 0.0077
-1.250 0.0472 0.01481 0.00766 0.0020 0.5625 0.0076
-1.000 0.0762 0.01362 0.00627 0.0022 0.5607 0.0079
-0.750 0.1047 0.01273 0.00524 0.0023 0.5591 0.0082
-0.500 0.1324 0.01207 0.00449 0.0025 0.5576 0.0089
-0.250 0.1588 0.01148 0.00383 0.0029 0.5562 0.0105
0.000 0.1847 0.01103 0.00334 0.0033 0.5548 0.0113
0.250 0.2106 0.01071 0.00297 0.0037 0.5534 0.0126
0.500 0.2369 0.01046 0.00269 0.0040 0.5519 0.0143
0.750 0.2637 0.01031 0.00252 0.0042 0.5501 0.0167
1.000 0.2901 0.01010 0.00226 0.0045 0.5483 0.0221
1.250 0.3158 0.00988 0.00217 0.0049 0.5463 0.0675
1.500 0.3305 0.00887 0.00217 0.0070 0.5447 0.4547
1.750 0.3310 0.00772 0.00224 0.0131 0.5433 0.8843
2.000 0.3552 0.00783 0.00238 0.0142 0.5419 0.9116
2.250 0.3819 0.00798 0.00253 0.0147 0.5407 0.9271
2.500 0.4104 0.00805 0.00256 0.0145 0.5394 0.9299
2.750 0.4384 0.00813 0.00261 0.0144 0.5382 0.9328
3.000 0.4652 0.00819 0.00267 0.0145 0.5367 0.9365
3.250 0.4925 0.00823 0.00273 0.0144 0.5350 0.9390
3.500 0.5221 0.00829 0.00281 0.0139 0.5331 0.9407
3.750 0.5510 0.00835 0.00288 0.0135 0.5310 0.9422
4.000 0.5792 0.00835 0.00285 0.0132 0.5262 0.9427
4.250 0.6075 0.00831 0.00283 0.0129 0.5157 0.9433
4.500 0.6355 0.00833 0.00283 0.0126 0.5076 0.9440
4.750 0.6637 0.00836 0.00288 0.0122 0.5004 0.9447
5.000 0.6916 0.00840 0.00295 0.0119 0.4919 0.9455
5.250 0.7193 0.00847 0.00301 0.0116 0.4795 0.9464
5.500 0.7468 0.00856 0.00310 0.0113 0.4637 0.9475
5.750 0.7729 0.00879 0.00325 0.0111 0.4269 0.9487
6.000 0.7897 0.01013 0.00407 0.0115 0.3234 0.9513
6.250 0.8102 0.01106 0.00476 0.0115 0.2583 0.9534
6.500 0.8272 0.01270 0.00596 0.0112 0.1624 0.9559
6.750 0.8427 0.01429 0.00719 0.0109 0.0804 0.9594
7.000 0.8657 0.01490 0.00780 0.0105 0.0698 0.9621
7.250 0.8873 0.01550 0.00842 0.0103 0.0637 0.9655
7.500 0.9131 0.01620 0.00916 0.0091 0.0587 0.9674
7.750 0.9401 0.01696 0.01003 0.0073 0.0562 0.9696
8.000 0.9654 0.01795 0.01108 0.0053 0.0535 0.9726
8.250 0.9862 0.01903 0.01222 0.0041 0.0515 0.9773
8.500 1.0094 0.02029 0.01352 0.0020 0.0475 0.9803
8.750 1.0329 0.02119 0.01447 0.0007 0.0444 0.9838
9.000 1.0545 0.02215 0.01551 -0.0005 0.0413 0.9880
9.250 1.0743 0.02340 0.01675 -0.0018 0.0368 0.9926
9.500 1.0953 0.02435 0.01775 -0.0028 0.0318 0.9979
9.750 1.1099 0.02556 0.01896 -0.0030 0.0260 1.0000
10.000 1.1053 0.02706 0.02035 -0.0001 0.0136 1.0000
10.250 1.1040 0.02866 0.02195 0.0022 0.0090 1.0000
10.500 1.1037 0.03035 0.02366 0.0041 0.0071 1.0000
10.750 1.1084 0.03176 0.02515 0.0056 0.0061 1.0000
11.000 1.1116 0.03334 0.02678 0.0071 0.0052 1.0000
11.250 1.1154 0.03493 0.02844 0.0084 0.0048 1.0000
11.500 1.1221 0.03633 0.02997 0.0095 0.0043 1.0000
11.750 1.1279 0.03785 0.03158 0.0106 0.0042 1.0000
12.000 1.1349 0.03932 0.03312 0.0115 0.0038 1.0000
12.250 1.1411 0.04092 0.03478 0.0123 0.0035 1.0000
12.500 1.1449 0.04276 0.03670 0.0132 0.0031 1.0000
12.750 1.1493 0.04466 0.03870 0.0140 0.0030 1.0000
13.000 1.1544 0.04654 0.04067 0.0147 0.0028 1.0000
13.250 1.1598 0.04842 0.04264 0.0152 0.0027 1.0000
13.500 1.1632 0.05052 0.04483 0.0158 0.0025 1.0000
13.750 1.1681 0.05255 0.04694 0.0162 0.0023 1.0000
14.000 1.1696 0.05497 0.04947 0.0166 0.0022 1.0000
14.250 1.1700 0.05755 0.05217 0.0171 0.0024 1.0000
14.500 1.1716 0.06010 0.05480 0.0172 0.0021 1.0000
14.750 1.1721 0.06279 0.05759 0.0173 0.0021 1.0000
15.000 1.1722 0.06560 0.06050 0.0173 0.0021 1.0000
15.250 1.1651 0.06936 0.06439 0.0174 0.0020 1.0000
15.500 1.1631 0.07257 0.06771 0.0172 0.0020 1.0000
15.750 1.1626 0.07575 0.07106 0.0169 0.0019 1.0000
16.000 1.1580 0.07948 0.07494 0.0164 0.0020 1.0000
16.250 1.1540 0.08336 0.07899 0.0157 0.0019 1.0000
16.500 1.1476 0.08763 0.08343 0.0148 0.0018 1.0000
16.750 1.1395 0.09232 0.08829 0.0137 0.0018 1.0000
17.000 1.1366 0.09646 0.09258 0.0122 0.0016 1.0000
17.250 1.1204 0.10274 0.09905 0.0104 0.0017 1.0000
17.500 1.1175 0.10715 0.10357 0.0085 0.0016 1.0000
17.750 1.1034 0.11353 0.11013 0.0061 0.0017 1.0000
18.000 1.0929 0.11952 0.11625 0.0034 0.0017 1.0000
18.250 1.0765 0.12697 0.12388 0.0001 0.0017 1.0000
18.500 1.0631 0.13410 0.13115 -0.0034 0.0017 1.0000
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Polar data table (+)
Polar graphs
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