NACA 6-H-10 AIRFOIL (n6h10-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 6-H-10 AIRFOIL (n6h10-il) Reynolds number: 200,000 Max Cl/Cd: 68.16 at α=6.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n6h10-il-200000-n5.txt Download as CSV file: xf-n6h10-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3954 0.09948 0.09530 0.0014 0.6482 0.0155
-7.750 -0.3947 0.09656 0.09240 -0.0012 0.6460 0.0156
-7.500 -0.3939 0.09390 0.08973 -0.0026 0.6440 0.0157
-7.250 -0.3891 0.09105 0.08685 -0.0044 0.6422 0.0158
-7.000 -0.3813 0.08797 0.08371 -0.0060 0.6405 0.0159
-6.750 -0.3719 0.08480 0.08048 -0.0075 0.6390 0.0160
-6.500 -0.3609 0.08151 0.07716 -0.0090 0.6366 0.0161
-6.250 -0.3485 0.07819 0.07378 -0.0104 0.6343 0.0161
-6.000 -0.3350 0.07483 0.07035 -0.0115 0.6321 0.0162
-5.750 -0.3205 0.07147 0.06689 -0.0124 0.6302 0.0162
-5.500 -0.3052 0.06810 0.06341 -0.0131 0.6284 0.0162
-5.250 -0.2890 0.06474 0.05993 -0.0135 0.6269 0.0163
-5.000 -0.2720 0.06146 0.05649 -0.0137 0.6255 0.0163
-4.750 -0.2598 0.05645 0.05136 -0.0138 0.6241 0.0166
-4.500 -0.2475 0.05181 0.04660 -0.0137 0.6221 0.0172
-4.250 -0.2313 0.04855 0.04324 -0.0135 0.6201 0.0177
-4.000 -0.2132 0.04572 0.04030 -0.0132 0.6181 0.0184
-3.750 -0.1936 0.04312 0.03755 -0.0128 0.6163 0.0192
-3.500 -0.1727 0.04074 0.03501 -0.0123 0.6146 0.0208
-3.250 -0.1502 0.03842 0.03248 -0.0117 0.6131 0.0238
-3.000 -0.1168 0.03748 0.03099 -0.0101 0.6116 0.0283
-2.750 -0.0936 0.03525 0.02844 -0.0089 0.6103 0.0284
-2.250 -0.0512 0.02881 0.02167 -0.0075 0.6061 0.0301
-2.000 -0.0219 0.02580 0.01832 -0.0058 0.6044 0.0151
-1.750 0.0044 0.02365 0.01582 -0.0049 0.6027 0.0145
-1.500 0.0322 0.02185 0.01368 -0.0043 0.6012 0.0143
-1.250 0.0610 0.02037 0.01188 -0.0038 0.5997 0.0146
-1.000 0.0895 0.01997 0.01126 -0.0035 0.5984 0.0163
-0.750 0.1194 0.01836 0.00937 -0.0034 0.5972 0.0167
-0.500 0.1487 0.01704 0.00788 -0.0034 0.5962 0.0173
-0.250 0.1761 0.01614 0.00696 -0.0032 0.5945 0.0184
0.000 0.2023 0.01567 0.00651 -0.0030 0.5925 0.0219
0.250 0.2277 0.01526 0.00606 -0.0025 0.5902 0.0248
0.500 0.2532 0.01498 0.00568 -0.0020 0.5880 0.0278
0.750 0.2784 0.01470 0.00536 -0.0015 0.5860 0.0387
1.000 0.2964 0.01386 0.00521 0.0003 0.5843 0.2641
1.250 0.3310 0.01327 0.00628 0.0012 0.5830 0.9391
1.500 0.4293 0.01402 0.00678 -0.0126 0.5818 0.9660
1.750 0.4772 0.01417 0.00689 -0.0171 0.5797 0.9706
2.000 0.5089 0.01430 0.00700 -0.0182 0.5775 0.9732
2.250 0.5346 0.01438 0.00707 -0.0181 0.5755 0.9735
2.500 0.5596 0.01446 0.00714 -0.0179 0.5737 0.9736
2.750 0.5845 0.01454 0.00721 -0.0176 0.5721 0.9738
3.000 0.6095 0.01459 0.00725 -0.0173 0.5705 0.9740
3.250 0.6344 0.01463 0.00728 -0.0170 0.5691 0.9743
3.500 0.6592 0.01467 0.00731 -0.0166 0.5677 0.9746
3.750 0.6843 0.01485 0.00759 -0.0165 0.5654 0.9750
4.000 0.7092 0.01505 0.00787 -0.0165 0.5630 0.9756
4.250 0.7337 0.01510 0.00798 -0.0161 0.5594 0.9762
4.500 0.7578 0.01472 0.00756 -0.0151 0.5538 0.9768
4.750 0.7839 0.01451 0.00742 -0.0150 0.5433 0.9774
5.000 0.8111 0.01429 0.00723 -0.0150 0.5349 0.9779
5.250 0.8384 0.01415 0.00717 -0.0151 0.5266 0.9785
5.500 0.8661 0.01415 0.00729 -0.0155 0.5183 0.9792
5.750 0.8937 0.01405 0.00725 -0.0157 0.5074 0.9800
6.000 0.9211 0.01402 0.00735 -0.0160 0.4926 0.9809
6.250 0.9478 0.01401 0.00741 -0.0162 0.4692 0.9818
6.500 0.9713 0.01425 0.00729 -0.0160 0.3823 0.9830
6.750 0.9804 0.01677 0.00899 -0.0161 0.2343 0.9857
7.000 0.9855 0.02000 0.01145 -0.0171 0.1042 0.9890
7.250 1.0019 0.02145 0.01286 -0.0178 0.0848 0.9916
7.500 1.0187 0.02275 0.01424 -0.0184 0.0776 0.9945
7.750 1.0347 0.02437 0.01593 -0.0198 0.0727 0.9977
8.000 1.0410 0.02555 0.01723 -0.0185 0.0706 1.0000
8.250 1.0256 0.02651 0.01825 -0.0128 0.0698 1.0000
8.500 1.0146 0.02750 0.01930 -0.0080 0.0688 1.0000
8.750 1.0088 0.02858 0.02043 -0.0041 0.0674 1.0000
9.000 1.0077 0.02974 0.02165 -0.0012 0.0661 1.0000
9.250 1.0080 0.03104 0.02305 0.0012 0.0645 1.0000
9.500 1.0101 0.03238 0.02445 0.0032 0.0631 1.0000
9.750 1.0111 0.03390 0.02599 0.0052 0.0609 1.0000
10.000 1.0168 0.03517 0.02732 0.0067 0.0582 1.0000
10.250 1.0281 0.03608 0.02834 0.0077 0.0541 1.0000
10.500 1.0367 0.03722 0.02948 0.0087 0.0479 1.0000
10.750 1.0475 0.03824 0.03059 0.0096 0.0431 1.0000
11.000 1.0586 0.03926 0.03167 0.0104 0.0378 1.0000
11.250 1.0703 0.04027 0.03270 0.0112 0.0275 1.0000
11.500 1.0743 0.04192 0.03425 0.0124 0.0154 1.0000
11.750 1.0766 0.04383 0.03613 0.0135 0.0127 1.0000
12.000 1.0796 0.04578 0.03815 0.0146 0.0111 1.0000
12.250 1.0830 0.04775 0.04030 0.0156 0.0100 1.0000
12.500 1.0871 0.04966 0.04235 0.0165 0.0088 1.0000
12.750 1.0902 0.05177 0.04456 0.0174 0.0079 1.0000
13.000 1.0921 0.05406 0.04698 0.0181 0.0074 1.0000
13.500 1.0947 0.05900 0.05221 0.0195 0.0068 1.0000
13.750 1.0975 0.06141 0.05478 0.0200 0.0066 1.0000
14.000 1.0992 0.06401 0.05755 0.0205 0.0063 1.0000
14.250 1.1006 0.06669 0.06039 0.0208 0.0062 1.0000
14.500 1.1010 0.06963 0.06349 0.0209 0.0058 1.0000
14.750 1.1005 0.07270 0.06673 0.0209 0.0057 1.0000
15.000 1.0998 0.07588 0.07007 0.0207 0.0056 1.0000
15.250 1.0953 0.07971 0.07401 0.0200 0.0052 1.0000
15.500 1.0919 0.08346 0.07794 0.0195 0.0052 1.0000
15.750 1.0833 0.08798 0.08254 0.0183 0.0049 1.0000
16.000 1.0790 0.09222 0.08701 0.0174 0.0050 1.0000
16.250 1.0691 0.09748 0.09244 0.0157 0.0047 1.0000
16.500 1.0595 0.10293 0.09812 0.0138 0.0047 1.0000
16.750 1.0479 0.10898 0.10438 0.0114 0.0046 1.0000
17.000 1.0352 0.11553 0.11113 0.0085 0.0046 1.0000
17.250 1.0219 0.12250 0.11830 0.0052 0.0046 1.0000
17.500 1.0101 0.12937 0.12531 0.0018 0.0046 1.0000
17.750 0.9965 0.13701 0.13310 -0.0022 0.0047 1.0000
18.000 0.9855 0.14425 0.14045 -0.0061 0.0048 1.0000
18.250 0.9691 0.15378 0.15012 -0.0114 0.0047 1.0000
18.500 0.9561 0.16279 0.15922 -0.0163 0.0048 1.0000
18.750 0.7194 0.15829 0.15530 -0.0129 0.0070 1.0000
19.000 0.7055 0.16579 0.16282 -0.0165 0.0071 1.0000
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