NACA 642-015A AIRFOIL (n64015a-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 642-015A AIRFOIL (n64015a-il) Reynolds number: 50,000 Max Cl/Cd: 28.61 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64015a-il-50000.txt Download as CSV file: xf-n64015a-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 642-015A 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6976 0.09939 0.09208 -0.0278 1.0000 0.1445
-11.500 -0.7049 0.09283 0.08553 -0.0301 1.0000 0.1418
-11.250 -0.8211 0.07971 0.07217 -0.0378 1.0000 0.1301
-11.000 -0.8309 0.07491 0.06728 -0.0375 1.0000 0.1286
-10.750 -0.8503 0.07069 0.06293 -0.0358 1.0000 0.1272
-10.500 -0.8694 0.06673 0.05877 -0.0336 1.0000 0.1260
-10.250 -0.8845 0.06288 0.05464 -0.0311 1.0000 0.1251
-10.000 -0.8938 0.05928 0.05074 -0.0285 1.0000 0.1249
-9.750 -0.8978 0.05593 0.04705 -0.0259 1.0000 0.1255
-9.500 -0.8979 0.05277 0.04352 -0.0233 1.0000 0.1266
-9.250 -0.8944 0.04975 0.04011 -0.0208 1.0000 0.1278
-9.000 -0.8868 0.04686 0.03679 -0.0183 1.0000 0.1292
-8.750 -0.8775 0.04434 0.03376 -0.0157 1.0000 0.1310
-8.500 -0.8514 0.04140 0.03099 -0.0158 1.0000 0.1375
-8.250 -0.8345 0.03925 0.02849 -0.0141 1.0000 0.1437
-8.000 -0.8059 0.03677 0.02604 -0.0140 1.0000 0.1524
-7.750 -0.7771 0.03466 0.02392 -0.0139 1.0000 0.1656
-7.500 -0.7456 0.03267 0.02205 -0.0139 1.0000 0.1851
-7.250 -0.7187 0.03074 0.02040 -0.0133 1.0000 0.2135
-7.000 -0.7076 0.02884 0.01889 -0.0105 1.0000 0.2528
-6.750 -0.7134 0.02674 0.01763 -0.0053 1.0000 0.3088
-6.500 -0.7304 0.02553 0.01807 0.0035 1.0000 0.4498
-6.250 -0.7315 0.02801 0.02087 0.0135 1.0000 0.5744
-6.000 -0.6806 0.03350 0.02609 0.0192 1.0000 0.6389
-5.750 -0.4984 0.04376 0.03528 0.0129 1.0000 0.7069
-5.500 -0.4401 0.04492 0.03607 0.0116 1.0000 0.7415
-5.250 -0.3733 0.04506 0.03587 0.0076 1.0000 0.7694
-5.000 -0.3558 0.04438 0.03506 0.0087 1.0000 0.7907
-4.750 -0.3166 0.04358 0.03409 0.0067 1.0000 0.8104
-4.500 -0.2877 0.04271 0.03311 0.0057 1.0000 0.8275
-4.250 -0.2715 0.04193 0.03226 0.0063 1.0000 0.8423
-4.000 -0.2497 0.04115 0.03141 0.0061 1.0000 0.8560
-3.750 -0.2243 0.04033 0.03053 0.0051 1.0000 0.8684
-3.500 -0.2160 0.03977 0.02994 0.0067 1.0000 0.8791
-3.250 -0.2194 0.03946 0.02961 0.0100 1.0000 0.8891
-3.000 -0.2021 0.03889 0.02900 0.0098 1.0000 0.8987
-2.750 -0.2126 0.03870 0.02881 0.0141 1.0000 0.9078
-2.500 -0.1898 0.03817 0.02822 0.0127 1.0000 0.9166
-2.250 -0.1901 0.03789 0.02793 0.0151 1.0000 0.9253
-2.000 -0.1704 0.03750 0.02748 0.0141 1.0000 0.9338
-1.750 -0.1622 0.03721 0.02717 0.0149 1.0000 0.9422
-1.500 -0.1428 0.03691 0.02683 0.0137 1.0000 0.9507
-1.250 -0.1259 0.03666 0.02655 0.0129 1.0000 0.9590
-1.000 -0.1062 0.03647 0.02633 0.0115 1.0000 0.9678
-0.750 -0.0800 0.03628 0.02612 0.0088 1.0000 0.9761
-0.500 -0.0536 0.03617 0.02600 0.0061 1.0000 0.9851
-0.250 -0.0218 0.03612 0.02593 0.0022 1.0000 0.9943
0.000 0.0000 0.03611 0.02592 0.0000 1.0000 1.0000
0.250 0.0218 0.03612 0.02593 -0.0022 0.9943 1.0000
0.500 0.0535 0.03617 0.02600 -0.0061 0.9851 1.0000
0.750 0.0800 0.03627 0.02612 -0.0088 0.9761 1.0000
1.000 0.1061 0.03646 0.02632 -0.0115 0.9678 1.0000
1.250 0.1259 0.03665 0.02654 -0.0129 0.9590 1.0000
1.500 0.1428 0.03690 0.02682 -0.0137 0.9507 1.0000
1.750 0.1621 0.03719 0.02715 -0.0149 0.9422 1.0000
2.000 0.1703 0.03748 0.02747 -0.0141 0.9338 1.0000
2.250 0.1899 0.03787 0.02791 -0.0151 0.9253 1.0000
2.500 0.1897 0.03815 0.02821 -0.0127 0.9166 1.0000
2.750 0.2124 0.03868 0.02878 -0.0140 0.9078 1.0000
3.000 0.2019 0.03887 0.02898 -0.0098 0.8987 1.0000
3.250 0.2194 0.03944 0.02959 -0.0100 0.8891 1.0000
3.500 0.2159 0.03975 0.02992 -0.0066 0.8791 1.0000
3.750 0.2243 0.04031 0.03050 -0.0051 0.8684 1.0000
4.000 0.2498 0.04113 0.03139 -0.0061 0.8560 1.0000
4.250 0.2714 0.04191 0.03224 -0.0063 0.8424 1.0000
4.500 0.2878 0.04269 0.03308 -0.0057 0.8275 1.0000
4.750 0.3168 0.04356 0.03407 -0.0067 0.8104 1.0000
5.000 0.3562 0.04436 0.03505 -0.0088 0.7907 1.0000
5.250 0.3737 0.04505 0.03585 -0.0077 0.7694 1.0000
5.500 0.4404 0.04490 0.03605 -0.0116 0.7415 1.0000
5.750 0.4987 0.04374 0.03526 -0.0130 0.7069 1.0000
6.000 0.6806 0.03350 0.02608 -0.0192 0.6390 1.0000
6.250 0.7315 0.02801 0.02087 -0.0135 0.5746 1.0000
6.500 0.7304 0.02553 0.01806 -0.0035 0.4502 1.0000
6.750 0.7134 0.02673 0.01763 0.0053 0.3090 1.0000
7.000 0.7076 0.02883 0.01889 0.0105 0.2528 1.0000
7.250 0.7187 0.03073 0.02039 0.0133 0.2136 1.0000
7.500 0.7455 0.03267 0.02204 0.0139 0.1851 1.0000
7.750 0.7771 0.03465 0.02392 0.0139 0.1657 1.0000
8.000 0.8059 0.03677 0.02603 0.0140 0.1524 1.0000
8.250 0.8345 0.03925 0.02849 0.0141 0.1437 1.0000
8.500 0.8514 0.04140 0.03099 0.0158 0.1375 1.0000
8.750 0.8775 0.04434 0.03376 0.0157 0.1310 1.0000
9.000 0.8869 0.04686 0.03678 0.0183 0.1292 1.0000
9.250 0.8945 0.04975 0.04011 0.0207 0.1278 1.0000
9.500 0.8980 0.05277 0.04352 0.0233 0.1266 1.0000
9.750 0.8980 0.05593 0.04705 0.0259 0.1255 1.0000
10.000 0.8940 0.05929 0.05075 0.0285 0.1249 1.0000
10.250 0.8847 0.06289 0.05466 0.0310 0.1251 1.0000
10.500 0.8697 0.06674 0.05879 0.0335 0.1260 1.0000
10.750 0.8508 0.07072 0.06296 0.0357 0.1273 1.0000
11.000 0.8314 0.07495 0.06733 0.0373 0.1286 1.0000
11.250 0.8220 0.07976 0.07222 0.0377 0.1302 1.0000
11.500 0.7056 0.09292 0.08562 0.0299 0.1418 1.0000
11.750 0.6985 0.09946 0.09215 0.0277 0.1445 1.0000
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Polar data table (+)
Polar graphs
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