NACA 0012 AIRFOILS (n0012-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: NACA 0012 AIRFOILS (n0012-il) Reynolds number: 50,000 Max Cl/Cd: 25.67 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n0012-il-50000.txt Download as CSV file: xf-n0012-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 0012 AIRFOILS
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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.8195 0.07577 0.06813 -0.0225 1.0000 0.1547
-9.750 -0.8355 0.07025 0.06249 -0.0224 1.0000 0.1540
-9.500 -0.8519 0.06488 0.05690 -0.0217 1.0000 0.1535
-9.250 -0.8654 0.05977 0.05144 -0.0204 1.0000 0.1535
-9.000 -0.8743 0.05500 0.04622 -0.0186 1.0000 0.1540
-8.750 -0.8723 0.05095 0.04184 -0.0169 1.0000 0.1566
-8.500 -0.8537 0.04867 0.03960 -0.0158 1.0000 0.1632
-8.250 -0.8499 0.04511 0.03550 -0.0138 1.0000 0.1676
-8.000 -0.8371 0.04218 0.03233 -0.0122 1.0000 0.1739
-7.750 -0.8225 0.03983 0.02970 -0.0105 1.0000 0.1828
-7.500 -0.8044 0.03756 0.02736 -0.0092 1.0000 0.1924
-7.250 -0.7876 0.03540 0.02498 -0.0076 1.0000 0.2045
-7.000 -0.7691 0.03350 0.02292 -0.0061 1.0000 0.2190
-6.750 -0.7491 0.03186 0.02130 -0.0047 1.0000 0.2365
-6.500 -0.7296 0.03038 0.01986 -0.0031 1.0000 0.2578
-6.250 -0.7105 0.02892 0.01844 -0.0014 1.0000 0.2830
-6.000 -0.6919 0.02765 0.01725 0.0005 1.0000 0.3133
-5.750 -0.6730 0.02658 0.01641 0.0025 1.0000 0.3480
-5.500 -0.6548 0.02564 0.01570 0.0047 1.0000 0.3875
-5.250 -0.6369 0.02482 0.01509 0.0072 1.0000 0.4301
-5.000 -0.6195 0.02413 0.01455 0.0098 1.0000 0.4753
-4.750 -0.6002 0.02367 0.01438 0.0126 1.0000 0.5186
-4.500 -0.5816 0.02327 0.01415 0.0155 1.0000 0.5624
-4.250 -0.5628 0.02300 0.01400 0.0186 1.0000 0.6051
-4.000 -0.5431 0.02285 0.01398 0.0218 1.0000 0.6455
-3.750 -0.5227 0.02282 0.01405 0.0250 1.0000 0.6846
-3.500 -0.5030 0.02279 0.01404 0.0283 1.0000 0.7237
-3.250 -0.4793 0.02298 0.01425 0.0313 1.0000 0.7602
-3.000 -0.4493 0.02334 0.01458 0.0333 1.0000 0.7958
-2.750 -0.4085 0.02382 0.01495 0.0333 1.0000 0.8311
-2.500 -0.3500 0.02437 0.01531 0.0298 1.0000 0.8657
-2.250 -0.2694 0.02476 0.01543 0.0214 1.0000 0.8987
-2.000 -0.1829 0.02464 0.01505 0.0107 1.0000 0.9296
-1.750 -0.1037 0.02407 0.01431 0.0003 1.0000 0.9591
-1.500 -0.0223 0.02312 0.01322 -0.0113 1.0000 0.9863
-1.250 0.0255 0.02220 0.01226 -0.0174 1.0000 1.0000
-1.000 0.0285 0.02168 0.01177 -0.0155 1.0000 1.0000
-0.750 0.0275 0.02128 0.01142 -0.0127 1.0000 1.0000
-0.500 0.0225 0.02100 0.01117 -0.0092 1.0000 1.0000
-0.250 0.0128 0.02083 0.01103 -0.0049 1.0000 1.0000
0.000 0.0000 0.02078 0.01099 0.0000 1.0000 1.0000
0.250 -0.0128 0.02083 0.01103 0.0049 1.0000 1.0000
0.500 -0.0225 0.02099 0.01117 0.0092 1.0000 1.0000
0.750 -0.0275 0.02128 0.01141 0.0127 1.0000 1.0000
1.000 -0.0285 0.02168 0.01177 0.0155 1.0000 1.0000
1.250 -0.0255 0.02219 0.01225 0.0174 1.0000 1.0000
1.500 0.0223 0.02311 0.01322 0.0113 0.9864 1.0000
1.750 0.1036 0.02406 0.01430 -0.0003 0.9592 1.0000
2.000 0.1829 0.02463 0.01505 -0.0106 0.9296 1.0000
2.250 0.2694 0.02476 0.01543 -0.0214 0.8987 1.0000
2.500 0.3499 0.02436 0.01530 -0.0298 0.8657 1.0000
2.750 0.4085 0.02382 0.01494 -0.0333 0.8311 1.0000
3.000 0.4492 0.02334 0.01458 -0.0333 0.7958 1.0000
3.250 0.4792 0.02298 0.01425 -0.0313 0.7602 1.0000
3.500 0.5030 0.02279 0.01404 -0.0283 0.7237 1.0000
3.750 0.5227 0.02282 0.01405 -0.0250 0.6847 1.0000
4.000 0.5431 0.02284 0.01398 -0.0218 0.6456 1.0000
4.250 0.5627 0.02300 0.01400 -0.0186 0.6051 1.0000
4.500 0.5816 0.02327 0.01415 -0.0155 0.5624 1.0000
4.750 0.6001 0.02367 0.01438 -0.0126 0.5187 1.0000
5.000 0.6194 0.02413 0.01455 -0.0098 0.4754 1.0000
5.250 0.6368 0.02482 0.01509 -0.0072 0.4301 1.0000
5.500 0.6547 0.02564 0.01570 -0.0047 0.3875 1.0000
5.750 0.6730 0.02658 0.01641 -0.0025 0.3481 1.0000
6.000 0.6919 0.02765 0.01724 -0.0005 0.3133 1.0000
6.250 0.7105 0.02892 0.01844 0.0014 0.2830 1.0000
6.500 0.7295 0.03038 0.01986 0.0031 0.2578 1.0000
6.750 0.7491 0.03186 0.02130 0.0047 0.2365 1.0000
7.000 0.7691 0.03350 0.02292 0.0061 0.2190 1.0000
7.250 0.7876 0.03540 0.02498 0.0076 0.2045 1.0000
7.500 0.8044 0.03756 0.02736 0.0092 0.1924 1.0000
7.750 0.8226 0.03983 0.02970 0.0105 0.1828 1.0000
8.000 0.8372 0.04218 0.03233 0.0122 0.1739 1.0000
8.250 0.8499 0.04511 0.03550 0.0138 0.1676 1.0000
8.500 0.8538 0.04867 0.03960 0.0158 0.1632 1.0000
8.750 0.8727 0.05094 0.04181 0.0168 0.1565 1.0000
9.000 0.8743 0.05501 0.04623 0.0185 0.1540 1.0000
9.250 0.8655 0.05978 0.05145 0.0203 0.1535 1.0000
9.500 0.8520 0.06490 0.05691 0.0216 0.1535 1.0000
9.750 0.8357 0.07027 0.06251 0.0223 0.1540 1.0000
10.000 0.8198 0.07580 0.06817 0.0224 0.1547 1.0000
10.250 0.6499 0.10541 0.09769 -0.0021 0.2047 1.0000
10.500 0.6743 0.10839 0.10074 0.0003 0.2006 1.0000
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Polar data table (+)
Polar graphs
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