NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 16-012 (naca16012-il) Reynolds number: 500,000 Max Cl/Cd: 42.48 at α=2.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-naca16012-il-500000.txt Download as CSV file: xf-naca16012-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA 16-012
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6737 0.08087 0.07852 -0.0536 1.0000 0.0159
-11.750 -0.7026 0.07454 0.07213 -0.0557 1.0000 0.0157
-11.500 -0.7317 0.06975 0.06725 -0.0551 1.0000 0.0155
-11.250 -0.7607 0.06596 0.06338 -0.0526 1.0000 0.0154
-11.000 -0.7892 0.06288 0.06021 -0.0485 1.0000 0.0153
-10.750 -0.8183 0.06023 0.05746 -0.0430 1.0000 0.0151
-10.500 -0.8480 0.05798 0.05511 -0.0362 1.0000 0.0149
-10.250 -0.8786 0.05625 0.05328 -0.0282 1.0000 0.0148
-10.000 -1.0090 0.04025 0.03611 -0.0048 0.9993 0.0126
-9.750 -0.9418 0.04882 0.04540 -0.0117 0.9992 0.0140
-9.500 -0.9668 0.03941 0.03520 -0.0052 0.9957 0.0136
-9.250 -0.9563 0.03648 0.03193 -0.0035 0.9940 0.0142
-9.000 -0.9494 0.03323 0.02827 -0.0004 0.9911 0.0147
-8.750 -0.9369 0.02997 0.02456 0.0019 0.9888 0.0151
-8.500 -0.9179 0.02767 0.02189 0.0031 0.9872 0.0156
-8.250 -0.8943 0.02618 0.02012 0.0035 0.9859 0.0161
-8.000 -0.8717 0.02352 0.01713 0.0039 0.9853 0.0170
-7.750 -0.8442 0.02253 0.01608 0.0032 0.9844 0.0179
-7.500 -0.8141 0.02214 0.01565 0.0021 0.9835 0.0192
-7.250 -0.7971 0.02157 0.01500 0.0040 0.9799 0.0204
-7.000 -0.7720 0.02092 0.01423 0.0043 0.9778 0.0215
-6.750 -0.7460 0.01957 0.01273 0.0043 0.9766 0.0229
-6.500 -0.7199 0.01862 0.01176 0.0041 0.9751 0.0247
-6.250 -0.6910 0.01821 0.01133 0.0034 0.9737 0.0269
-6.000 -0.6610 0.01786 0.01092 0.0025 0.9727 0.0291
-5.750 -0.6345 0.01685 0.00982 0.0023 0.9717 0.0315
-5.500 -0.6051 0.01627 0.00923 0.0015 0.9709 0.0347
-5.250 -0.5945 0.01597 0.00890 0.0050 0.9653 0.0370
-5.000 -0.5684 0.01568 0.00857 0.0050 0.9630 0.0393
-4.750 -0.5434 0.01500 0.00784 0.0052 0.9611 0.0432
-4.500 -0.5136 0.01464 0.00747 0.0044 0.9598 0.0478
-4.250 -0.4826 0.01434 0.00713 0.0034 0.9588 0.0520
-4.000 -0.4523 0.01395 0.00678 0.0025 0.9578 0.0630
-3.750 -0.4439 0.01353 0.00653 0.0065 0.9519 0.0950
-3.500 -0.4273 0.01258 0.00617 0.0082 0.9488 0.2245
-3.250 -0.4137 0.01124 0.00574 0.0103 0.9464 0.4340
-3.000 -0.4069 0.00977 0.00534 0.0145 0.9441 0.6743
-2.750 -0.4020 0.00946 0.00546 0.0198 0.9381 0.7754
-2.500 -0.3749 0.00944 0.00556 0.0201 0.9357 0.8241
-2.250 -0.3440 0.00950 0.00564 0.0195 0.9341 0.8496
-2.000 -0.3116 0.00959 0.00572 0.0186 0.9328 0.8677
-1.750 -0.2794 0.00973 0.00587 0.0178 0.9316 0.8849
-1.500 -0.2428 0.01000 0.00612 0.0161 0.9310 0.8966
-1.250 -0.2057 0.01079 0.00692 0.0150 0.9306 0.9155
-1.000 -0.1500 0.01163 0.00776 0.0094 0.9315 0.9188
-0.750 -0.1063 0.01181 0.00791 0.0057 0.9313 0.9199
-0.500 -0.0636 0.01191 0.00800 0.0022 0.9310 0.9210
-0.250 -0.0526 0.01193 0.00800 0.0058 0.9253 0.9283
0.000 0.0000 0.01198 0.00805 0.0000 0.9266 0.9266
0.250 0.0525 0.01193 0.00800 -0.0058 0.9283 0.9253
0.500 0.0636 0.01191 0.00800 -0.0022 0.9210 0.9310
0.750 0.1063 0.01181 0.00791 -0.0057 0.9199 0.9313
1.000 0.1500 0.01163 0.00776 -0.0094 0.9188 0.9315
1.250 0.2052 0.01081 0.00695 -0.0150 0.9156 0.9306
1.500 0.2428 0.01000 0.00612 -0.0161 0.8967 0.9310
1.750 0.2793 0.00973 0.00586 -0.0178 0.8848 0.9317
2.000 0.3115 0.00958 0.00572 -0.0185 0.8676 0.9328
2.250 0.3440 0.00949 0.00564 -0.0195 0.8495 0.9341
2.500 0.3748 0.00944 0.00556 -0.0201 0.8243 0.9357
2.750 0.4019 0.00946 0.00546 -0.0197 0.7756 0.9381
3.000 0.4070 0.00977 0.00534 -0.0145 0.6751 0.9441
3.250 0.4136 0.01124 0.00574 -0.0103 0.4337 0.9464
3.500 0.4273 0.01258 0.00616 -0.0082 0.2244 0.9488
3.750 0.4437 0.01353 0.00653 -0.0064 0.0948 0.9520
4.000 0.4523 0.01394 0.00677 -0.0025 0.0629 0.9578
4.250 0.4826 0.01433 0.00712 -0.0034 0.0520 0.9588
4.500 0.5136 0.01464 0.00747 -0.0044 0.0478 0.9598
4.750 0.5434 0.01500 0.00784 -0.0053 0.0433 0.9612
5.000 0.5684 0.01568 0.00857 -0.0050 0.0393 0.9631
5.250 0.5945 0.01596 0.00889 -0.0050 0.0370 0.9653
5.500 0.6052 0.01626 0.00923 -0.0015 0.0347 0.9709
5.750 0.6345 0.01684 0.00981 -0.0024 0.0316 0.9717
6.000 0.6611 0.01785 0.01091 -0.0025 0.0291 0.9727
6.250 0.6911 0.01821 0.01133 -0.0034 0.0269 0.9738
6.500 0.7201 0.01861 0.01175 -0.0042 0.0247 0.9751
6.750 0.7462 0.01955 0.01271 -0.0044 0.0229 0.9766
7.000 0.7721 0.02091 0.01422 -0.0043 0.0215 0.9779
7.250 0.7970 0.02156 0.01498 -0.0040 0.0204 0.9800
7.500 0.8141 0.02211 0.01562 -0.0021 0.0192 0.9835
7.750 0.8441 0.02255 0.01611 -0.0032 0.0180 0.9844
8.000 0.8718 0.02350 0.01711 -0.0039 0.0170 0.9853
8.250 0.8944 0.02619 0.02013 -0.0035 0.0161 0.9859
8.500 0.9181 0.02765 0.02188 -0.0031 0.0156 0.9872
8.750 0.9372 0.02996 0.02456 -0.0019 0.0151 0.9889
9.000 0.9497 0.03322 0.02826 0.0003 0.0147 0.9912
9.250 0.9562 0.03654 0.03201 0.0035 0.0142 0.9941
9.500 0.9687 0.03909 0.03484 0.0049 0.0136 0.9957
9.750 0.9937 0.03886 0.03457 0.0041 0.0130 0.9973
10.000 1.0099 0.04017 0.03603 0.0047 0.0126 0.9993
10.250 0.8787 0.05626 0.05330 0.0282 0.0148 1.0000
10.500 0.8478 0.05801 0.05514 0.0363 0.0149 1.0000
10.750 0.8184 0.06020 0.05743 0.0430 0.0151 1.0000
11.000 0.7892 0.06291 0.06024 0.0485 0.0152 1.0000
11.250 0.7605 0.06600 0.06342 0.0526 0.0154 1.0000
11.500 0.7314 0.06984 0.06735 0.0552 0.0155 1.0000
11.750 0.7027 0.07453 0.07212 0.0557 0.0157 1.0000
12.000 0.6737 0.08092 0.07858 0.0535 0.0159 1.0000
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Polar data table (+)
Polar graphs
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