NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 50,000 Max Cl/Cd: 26.74 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n64012-il-50000.txt Download as CSV file: xf-n64012-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY 64-012 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.7458 0.07703 0.07028 -0.0217 1.0000 0.1520
-9.250 -0.7656 0.07160 0.06472 -0.0225 1.0000 0.1443
-9.000 -0.8093 0.06602 0.05858 -0.0225 1.0000 0.1340
-8.750 -0.8017 0.06131 0.05375 -0.0222 1.0000 0.1319
-8.500 -0.7993 0.05688 0.04903 -0.0215 1.0000 0.1291
-8.250 -0.7982 0.05250 0.04414 -0.0203 1.0000 0.1253
-8.000 -0.7976 0.04891 0.03964 -0.0182 1.0000 0.1210
-7.750 -0.7819 0.04560 0.03587 -0.0170 1.0000 0.1201
-7.500 -0.7648 0.04246 0.03231 -0.0158 1.0000 0.1206
-7.250 -0.7419 0.03918 0.02908 -0.0155 1.0000 0.1250
-7.000 -0.7200 0.03669 0.02632 -0.0146 1.0000 0.1292
-6.750 -0.6956 0.03429 0.02353 -0.0137 1.0000 0.1327
-6.500 -0.6680 0.03193 0.02111 -0.0133 1.0000 0.1401
-6.250 -0.6391 0.02998 0.01909 -0.0127 1.0000 0.1514
-6.000 -0.6082 0.02809 0.01734 -0.0122 1.0000 0.1671
-5.750 -0.5850 0.02642 0.01583 -0.0109 1.0000 0.1921
-5.500 -0.5729 0.02444 0.01434 -0.0083 1.0000 0.2285
-5.250 -0.5797 0.02169 0.01289 -0.0035 1.0000 0.3274
-5.000 -0.5954 0.02280 0.01563 0.0096 1.0000 0.6194
-4.750 -0.5760 0.02640 0.01908 0.0196 1.0000 0.6952
-4.500 -0.3636 0.03552 0.02675 0.0129 1.0000 0.8205
-4.250 -0.3119 0.03476 0.02567 0.0094 1.0000 0.8462
-4.000 -0.2805 0.03389 0.02461 0.0080 1.0000 0.8658
-3.750 -0.2506 0.03298 0.02353 0.0065 1.0000 0.8823
-3.500 -0.2176 0.03200 0.02240 0.0043 1.0000 0.8966
-3.250 -0.1849 0.03101 0.02130 0.0020 1.0000 0.9090
-3.000 -0.1566 0.03011 0.02033 0.0002 1.0000 0.9203
-2.750 -0.1360 0.02942 0.01959 -0.0004 1.0000 0.9307
-2.500 -0.1147 0.02878 0.01891 -0.0012 1.0000 0.9409
-2.250 -0.0895 0.02810 0.01820 -0.0028 1.0000 0.9506
-2.000 -0.0712 0.02764 0.01772 -0.0034 1.0000 0.9605
-1.750 -0.0548 0.02733 0.01738 -0.0039 1.0000 0.9704
-1.500 -0.0282 0.02684 0.01687 -0.0063 1.0000 0.9797
-1.250 -0.0026 0.02647 0.01649 -0.0087 1.0000 0.9891
-1.000 0.0268 0.02614 0.01614 -0.0119 1.0000 0.9986
-0.750 0.0242 0.02630 0.01631 -0.0096 1.0000 1.0000
-0.500 0.0158 0.02647 0.01649 -0.0063 1.0000 1.0000
-0.250 0.0077 0.02656 0.01659 -0.0031 1.0000 1.0000
0.000 0.0000 0.02660 0.01662 0.0000 1.0000 1.0000
0.250 -0.0077 0.02657 0.01659 0.0031 1.0000 1.0000
0.500 -0.0157 0.02646 0.01648 0.0063 1.0000 1.0000
0.750 -0.0242 0.02629 0.01630 0.0096 1.0000 1.0000
1.000 -0.0268 0.02614 0.01614 0.0119 0.9986 1.0000
1.250 0.0025 0.02647 0.01648 0.0087 0.9891 1.0000
1.500 0.0282 0.02684 0.01687 0.0063 0.9797 1.0000
1.750 0.0548 0.02732 0.01738 0.0039 0.9704 1.0000
2.000 0.0710 0.02763 0.01772 0.0035 0.9605 1.0000
2.250 0.0893 0.02809 0.01819 0.0029 0.9507 1.0000
2.500 0.1147 0.02877 0.01891 0.0012 0.9409 1.0000
2.750 0.1360 0.02941 0.01957 0.0004 0.9308 1.0000
3.000 0.1568 0.03010 0.02032 -0.0002 0.9203 1.0000
3.250 0.1851 0.03099 0.02129 -0.0020 0.9091 1.0000
3.500 0.2178 0.03199 0.02239 -0.0043 0.8966 1.0000
3.750 0.2509 0.03297 0.02352 -0.0065 0.8823 1.0000
4.000 0.2807 0.03388 0.02460 -0.0080 0.8658 1.0000
4.250 0.3123 0.03475 0.02566 -0.0094 0.8461 1.0000
4.500 0.3638 0.03550 0.02676 -0.0130 0.8205 1.0000
4.750 0.5760 0.02641 0.01909 -0.0197 0.6955 1.0000
5.000 0.5954 0.02281 0.01564 -0.0097 0.6198 1.0000
5.250 0.5798 0.02168 0.01289 0.0035 0.3281 1.0000
5.500 0.5730 0.02443 0.01433 0.0083 0.2290 1.0000
5.750 0.5850 0.02641 0.01583 0.0109 0.1920 1.0000
6.000 0.6082 0.02808 0.01734 0.0122 0.1671 1.0000
6.250 0.6391 0.02996 0.01908 0.0127 0.1516 1.0000
6.500 0.6680 0.03193 0.02111 0.0133 0.1401 1.0000
6.750 0.6956 0.03430 0.02353 0.0137 0.1326 1.0000
7.000 0.7200 0.03668 0.02631 0.0146 0.1293 1.0000
7.250 0.7419 0.03918 0.02909 0.0155 0.1251 1.0000
7.500 0.7648 0.04251 0.03236 0.0158 0.1205 1.0000
7.750 0.7819 0.04560 0.03587 0.0170 0.1201 1.0000
8.000 0.7975 0.04890 0.03963 0.0182 0.1211 1.0000
8.250 0.7982 0.05250 0.04416 0.0203 0.1253 1.0000
8.500 0.7993 0.05686 0.04901 0.0215 0.1291 1.0000
8.750 0.8021 0.06129 0.05372 0.0222 0.1320 1.0000
9.000 0.8093 0.06604 0.05860 0.0225 0.1340 1.0000
9.250 0.7657 0.07163 0.06478 0.0225 0.1444 1.0000
9.500 0.7459 0.07703 0.07028 0.0217 0.1520 1.0000
9.750 0.7192 0.08436 0.07766 0.0185 0.1650 1.0000
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Polar data table (+)
Polar graphs
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