NACA 0012 AIRFOILS (n0012-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 0012 AIRFOILS (n0012-il) Reynolds number: 50,000 Max Cl/Cd: 26.54 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n0012-il-50000-n5.txt Download as CSV file: xf-n0012-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7967 0.08718 0.07926 -0.0211 1.0000 0.0718
-11.750 -0.8409 0.07647 0.06839 -0.0284 1.0000 0.0708
-11.500 -0.8751 0.06932 0.06102 -0.0318 1.0000 0.0705
-11.250 -0.9008 0.06422 0.05569 -0.0322 1.0000 0.0706
-11.000 -0.9203 0.06020 0.05141 -0.0308 1.0000 0.0712
-10.750 -0.9332 0.05655 0.04743 -0.0289 1.0000 0.0721
-10.500 -0.9408 0.05300 0.04344 -0.0270 1.0000 0.0736
-10.250 -0.9394 0.05004 0.04017 -0.0253 1.0000 0.0757
-10.000 -0.9268 0.04819 0.03830 -0.0242 1.0000 0.0785
-9.750 -0.9169 0.04592 0.03579 -0.0228 1.0000 0.0815
-9.500 -0.9070 0.04340 0.03287 -0.0212 1.0000 0.0849
-9.250 -0.8921 0.04144 0.03079 -0.0200 1.0000 0.0887
-9.000 -0.8759 0.03979 0.02903 -0.0189 1.0000 0.0933
-8.750 -0.8596 0.03785 0.02668 -0.0176 1.0000 0.0987
-8.500 -0.8413 0.03641 0.02536 -0.0166 1.0000 0.1041
-8.250 -0.8228 0.03486 0.02354 -0.0154 1.0000 0.1113
-8.000 -0.8043 0.03353 0.02227 -0.0143 1.0000 0.1185
-7.750 -0.7850 0.03218 0.02078 -0.0131 1.0000 0.1277
-7.500 -0.7664 0.03098 0.01955 -0.0119 1.0000 0.1384
-7.250 -0.7478 0.02982 0.01843 -0.0106 1.0000 0.1500
-7.000 -0.7294 0.02872 0.01733 -0.0093 1.0000 0.1644
-6.750 -0.7109 0.02766 0.01630 -0.0080 1.0000 0.1813
-6.500 -0.6925 0.02667 0.01535 -0.0066 1.0000 0.2017
-6.250 -0.6748 0.02574 0.01453 -0.0051 1.0000 0.2241
-6.000 -0.6567 0.02488 0.01376 -0.0036 1.0000 0.2509
-5.750 -0.6386 0.02410 0.01309 -0.0021 1.0000 0.2810
-5.500 -0.6206 0.02339 0.01250 -0.0005 1.0000 0.3132
-5.250 -0.6025 0.02275 0.01198 0.0011 1.0000 0.3470
-5.000 -0.5842 0.02218 0.01154 0.0028 1.0000 0.3820
-4.750 -0.5655 0.02168 0.01116 0.0046 1.0000 0.4176
-4.500 -0.5464 0.02123 0.01083 0.0063 1.0000 0.4537
-4.250 -0.5270 0.02083 0.01051 0.0080 1.0000 0.4896
-4.000 -0.5073 0.02048 0.01026 0.0098 1.0000 0.5251
-3.750 -0.4879 0.02018 0.01007 0.0117 1.0000 0.5595
-3.500 -0.4681 0.01993 0.00990 0.0136 1.0000 0.5934
-3.250 -0.4481 0.01972 0.00979 0.0155 1.0000 0.6270
-3.000 -0.4277 0.01956 0.00971 0.0174 1.0000 0.6601
-2.750 -0.4071 0.01945 0.00966 0.0193 1.0000 0.6928
-2.500 -0.3861 0.01938 0.00966 0.0211 1.0000 0.7253
-2.250 -0.3639 0.01937 0.00970 0.0228 1.0000 0.7567
-2.000 -0.3402 0.01942 0.00979 0.0242 1.0000 0.7876
-1.750 -0.3144 0.01952 0.00990 0.0251 1.0000 0.8179
-1.500 -0.2858 0.01966 0.01005 0.0253 1.0000 0.8478
-1.250 -0.2542 0.01984 0.01020 0.0248 1.0000 0.8771
-1.000 -0.2167 0.02005 0.01036 0.0229 1.0000 0.9040
-0.750 -0.1745 0.02026 0.01053 0.0199 1.0000 0.9288
-0.500 -0.1277 0.02045 0.01068 0.0156 1.0000 0.9515
-0.250 -0.0696 0.02062 0.01081 0.0089 0.9951 0.9702
0.000 0.0000 0.02068 0.01086 0.0000 0.9836 0.9836
0.250 0.0696 0.02061 0.01081 -0.0089 0.9702 0.9951
0.500 0.1277 0.02045 0.01068 -0.0156 0.9515 1.0000
0.750 0.1745 0.02026 0.01053 -0.0199 0.9288 1.0000
1.000 0.2166 0.02005 0.01036 -0.0229 0.9041 1.0000
1.250 0.2542 0.01983 0.01019 -0.0248 0.8771 1.0000
1.500 0.2858 0.01966 0.01004 -0.0253 0.8478 1.0000
1.750 0.3144 0.01952 0.00990 -0.0251 0.8179 1.0000
2.000 0.3402 0.01942 0.00979 -0.0242 0.7876 1.0000
2.250 0.3639 0.01937 0.00970 -0.0228 0.7568 1.0000
2.500 0.3860 0.01938 0.00965 -0.0211 0.7253 1.0000
2.750 0.4070 0.01944 0.00966 -0.0193 0.6929 1.0000
3.000 0.4277 0.01956 0.00971 -0.0174 0.6601 1.0000
3.250 0.4480 0.01972 0.00979 -0.0155 0.6270 1.0000
3.500 0.4680 0.01993 0.00990 -0.0136 0.5934 1.0000
3.750 0.4878 0.02018 0.01007 -0.0117 0.5595 1.0000
4.250 0.5269 0.02083 0.01051 -0.0080 0.4896 1.0000
4.500 0.5464 0.02123 0.01082 -0.0063 0.4537 1.0000
4.750 0.5654 0.02167 0.01116 -0.0046 0.4177 1.0000
5.000 0.5841 0.02218 0.01154 -0.0028 0.3820 1.0000
5.250 0.6025 0.02275 0.01198 -0.0011 0.3471 1.0000
5.500 0.6206 0.02338 0.01250 0.0005 0.3132 1.0000
5.750 0.6386 0.02409 0.01308 0.0021 0.2810 1.0000
6.000 0.6566 0.02487 0.01376 0.0036 0.2509 1.0000
6.250 0.6748 0.02573 0.01453 0.0051 0.2241 1.0000
6.500 0.6925 0.02667 0.01535 0.0066 0.2017 1.0000
6.750 0.7109 0.02766 0.01630 0.0080 0.1813 1.0000
7.000 0.7294 0.02871 0.01733 0.0093 0.1644 1.0000
7.250 0.7479 0.02982 0.01843 0.0106 0.1500 1.0000
7.500 0.7664 0.03098 0.01955 0.0119 0.1384 1.0000
7.750 0.7851 0.03218 0.02078 0.0131 0.1277 1.0000
8.000 0.8044 0.03353 0.02227 0.0143 0.1185 1.0000
8.250 0.8228 0.03486 0.02354 0.0154 0.1113 1.0000
8.500 0.8414 0.03641 0.02536 0.0166 0.1041 1.0000
8.750 0.8597 0.03785 0.02668 0.0175 0.0987 1.0000
9.000 0.8760 0.03979 0.02903 0.0189 0.0932 1.0000
9.250 0.8922 0.04144 0.03079 0.0200 0.0887 1.0000
9.500 0.9071 0.04340 0.03287 0.0212 0.0849 1.0000
9.750 0.9171 0.04592 0.03580 0.0227 0.0815 1.0000
10.000 0.9270 0.04819 0.03830 0.0241 0.0785 1.0000
10.250 0.9397 0.05004 0.04016 0.0252 0.0757 1.0000
10.500 0.9411 0.05302 0.04346 0.0269 0.0736 1.0000
10.750 0.9335 0.05656 0.04745 0.0288 0.0721 1.0000
11.000 0.9206 0.06023 0.05144 0.0307 0.0711 1.0000
11.250 0.9012 0.06425 0.05572 0.0321 0.0706 1.0000
11.500 0.8756 0.06936 0.06106 0.0316 0.0704 1.0000
11.750 0.8414 0.07654 0.06847 0.0283 0.0708 1.0000
12.000 0.7971 0.08731 0.07940 0.0209 0.0718 1.0000
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Polar data table (+)
Polar graphs
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