N-12 (n12-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: N-12 (n12-il) Reynolds number: 50,000 Max Cl/Cd: 27.66 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n12-il-50000-n5.txt Download as CSV file: xf-n12-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: N-12
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6760 0.09257 0.08496 -0.0258 1.0000 0.0597
-11.000 -0.7001 0.08388 0.07632 -0.0316 1.0000 0.0590
-10.750 -0.7270 0.07679 0.06918 -0.0358 1.0000 0.0582
-10.500 -0.7546 0.07116 0.06346 -0.0375 1.0000 0.0575
-10.250 -0.7803 0.06678 0.05894 -0.0365 1.0000 0.0573
-10.000 -0.8023 0.06315 0.05514 -0.0339 1.0000 0.0572
-9.750 -0.8178 0.05943 0.05117 -0.0313 1.0000 0.0572
-9.500 -0.8278 0.05582 0.04725 -0.0286 1.0000 0.0573
-9.250 -0.8325 0.05238 0.04346 -0.0259 1.0000 0.0576
-9.000 -0.8322 0.04924 0.03996 -0.0233 1.0000 0.0584
-8.750 -0.8285 0.04633 0.03664 -0.0207 1.0000 0.0599
-8.500 -0.8215 0.04356 0.03338 -0.0183 1.0000 0.0617
-8.250 -0.8092 0.04095 0.03035 -0.0163 1.0000 0.0630
-8.000 -0.7919 0.03873 0.02799 -0.0149 1.0000 0.0643
-7.750 -0.7736 0.03681 0.02590 -0.0136 1.0000 0.0657
-7.500 -0.7552 0.03520 0.02414 -0.0121 1.0000 0.0683
-7.250 -0.7356 0.03367 0.02237 -0.0108 1.0000 0.0717
-7.000 -0.7142 0.03215 0.02066 -0.0095 1.0000 0.0743
-6.750 -0.6947 0.03086 0.01931 -0.0081 1.0000 0.0766
-6.500 -0.6751 0.02969 0.01805 -0.0066 1.0000 0.0795
-6.250 -0.6559 0.02858 0.01680 -0.0050 1.0000 0.0830
-6.000 -0.6377 0.02757 0.01571 -0.0033 1.0000 0.0875
-5.750 -0.6197 0.02665 0.01473 -0.0016 1.0000 0.0963
-5.500 -0.6024 0.02567 0.01373 0.0002 1.0000 0.1078
-5.250 -0.5851 0.02462 0.01270 0.0019 1.0000 0.1247
-5.000 -0.5682 0.02354 0.01172 0.0037 1.0000 0.1527
-4.750 -0.5519 0.02246 0.01089 0.0054 1.0000 0.1975
-4.500 -0.5379 0.02137 0.01030 0.0074 1.0000 0.2696
-4.250 -0.5222 0.02068 0.00988 0.0095 1.0000 0.3446
-4.000 -0.5057 0.02014 0.00945 0.0116 1.0000 0.4062
-3.750 -0.4918 0.01951 0.00917 0.0143 1.0000 0.4725
-3.500 -0.4757 0.01903 0.00899 0.0169 1.0000 0.5341
-3.250 -0.4572 0.01870 0.00898 0.0194 1.0000 0.6055
-3.000 -0.4355 0.01856 0.00910 0.0216 1.0000 0.6841
-2.750 -0.4052 0.01873 0.00943 0.0228 1.0000 0.7604
-2.500 -0.3695 0.01916 0.00987 0.0230 1.0000 0.8286
-2.250 -0.3223 0.01965 0.01023 0.0207 1.0000 0.8761
-2.000 -0.2672 0.02000 0.01035 0.0162 1.0000 0.9055
-1.750 -0.2146 0.02013 0.01028 0.0115 1.0000 0.9233
-1.500 -0.1684 0.02014 0.01012 0.0076 1.0000 0.9402
-1.250 -0.1250 0.02010 0.00996 0.0040 1.0000 0.9570
-1.000 -0.0825 0.02005 0.00981 0.0004 1.0000 0.9741
-0.750 -0.0364 0.01990 0.00960 -0.0040 1.0000 0.9887
-0.500 0.0021 0.01978 0.00942 -0.0073 1.0000 1.0000
-0.250 0.0048 0.01972 0.00938 -0.0039 1.0000 1.0000
0.000 0.0061 0.01973 0.00939 -0.0002 1.0000 1.0000
0.250 0.0069 0.01980 0.00946 0.0035 1.0000 1.0000
0.500 0.0258 0.01991 0.00959 0.0038 0.9946 1.0000
0.750 0.0942 0.01985 0.00958 -0.0047 0.9678 1.0000
1.000 0.1473 0.01970 0.00950 -0.0099 0.9425 1.0000
1.250 0.1971 0.01954 0.00943 -0.0144 0.9224 1.0000
1.500 0.2460 0.01937 0.00938 -0.0185 0.9017 1.0000
1.750 0.2942 0.01919 0.00932 -0.0223 0.8784 1.0000
2.000 0.3361 0.01904 0.00928 -0.0248 0.8504 1.0000
2.250 0.3733 0.01893 0.00928 -0.0262 0.8188 1.0000
2.500 0.4065 0.01883 0.00924 -0.0266 0.7821 1.0000
2.750 0.4341 0.01880 0.00925 -0.0259 0.7403 1.0000
3.000 0.4590 0.01883 0.00926 -0.0246 0.6960 1.0000
3.250 0.4820 0.01894 0.00928 -0.0229 0.6490 1.0000
3.500 0.5026 0.01917 0.00941 -0.0209 0.6036 1.0000
3.750 0.5218 0.01947 0.00963 -0.0188 0.5615 1.0000
4.000 0.5408 0.01983 0.00994 -0.0168 0.5251 1.0000
4.250 0.5583 0.02024 0.01025 -0.0146 0.4829 1.0000
4.500 0.5728 0.02071 0.01052 -0.0119 0.4309 1.0000
4.750 0.5851 0.02131 0.01077 -0.0089 0.3703 1.0000
5.000 0.5973 0.02204 0.01117 -0.0063 0.3094 1.0000
5.250 0.6121 0.02276 0.01175 -0.0041 0.2552 1.0000
5.500 0.6269 0.02361 0.01247 -0.0019 0.1876 1.0000
5.750 0.6397 0.02491 0.01328 0.0003 0.1380 1.0000
6.000 0.6529 0.02638 0.01451 0.0025 0.1204 1.0000
6.250 0.6678 0.02774 0.01584 0.0045 0.1081 1.0000
6.500 0.6844 0.02899 0.01718 0.0064 0.0984 1.0000
6.750 0.7012 0.03036 0.01852 0.0081 0.0924 1.0000
7.000 0.7219 0.03165 0.02005 0.0096 0.0877 1.0000
7.250 0.7427 0.03307 0.02163 0.0110 0.0836 1.0000
7.500 0.7621 0.03463 0.02316 0.0122 0.0795 1.0000
7.750 0.7805 0.03615 0.02503 0.0138 0.0749 1.0000
8.000 0.7986 0.03789 0.02700 0.0153 0.0717 1.0000
8.250 0.8162 0.03983 0.02914 0.0167 0.0694 1.0000
8.500 0.8321 0.04197 0.03136 0.0181 0.0669 1.0000
8.750 0.8407 0.04432 0.03423 0.0206 0.0643 1.0000
9.000 0.8479 0.04673 0.03702 0.0230 0.0619 1.0000
9.250 0.8533 0.04946 0.04007 0.0253 0.0605 1.0000
9.500 0.8560 0.05227 0.04318 0.0277 0.0592 1.0000
9.750 0.8576 0.05495 0.04608 0.0300 0.0580 1.0000
10.000 0.8612 0.05745 0.04865 0.0318 0.0563 1.0000
10.250 0.8539 0.06065 0.05204 0.0342 0.0554 1.0000
10.500 0.8353 0.06399 0.05567 0.0372 0.0550 1.0000
10.750 0.8133 0.06741 0.05928 0.0398 0.0549 1.0000
11.000 0.7908 0.07142 0.06346 0.0405 0.0549 1.0000
11.250 0.7697 0.07610 0.06826 0.0395 0.0551 1.0000
11.500 0.7473 0.08187 0.07411 0.0368 0.0553 1.0000
11.750 0.7263 0.08850 0.08079 0.0329 0.0556 1.0000
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Polar data table (+)
Polar graphs
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