NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 16-012 (naca16012-il) Reynolds number: 200,000 Max Cl/Cd: 32.51 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca16012-il-200000-n5.txt Download as CSV file: xf-naca16012-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.6765 0.08272 0.07916 -0.0523 1.0000 0.0121
-11.750 -0.7221 0.07176 0.06805 -0.0566 1.0000 0.0116
-11.500 -0.7631 0.06495 0.06107 -0.0555 1.0000 0.0114
-11.250 -0.7978 0.06025 0.05619 -0.0518 1.0000 0.0113
-11.000 -0.8291 0.05659 0.05235 -0.0465 1.0000 0.0112
-10.750 -0.8581 0.05364 0.04923 -0.0401 1.0000 0.0112
-10.500 -0.8857 0.05126 0.04666 -0.0327 1.0000 0.0112
-10.250 -0.9121 0.04928 0.04451 -0.0247 1.0000 0.0112
-10.000 -0.9323 0.04693 0.04193 -0.0175 1.0000 0.0113
-9.750 -0.9467 0.04445 0.03919 -0.0112 1.0000 0.0114
-9.500 -0.9559 0.04196 0.03641 -0.0055 0.9999 0.0116
-9.250 -0.9481 0.03865 0.03265 -0.0033 0.9974 0.0120
-9.000 -0.9359 0.03572 0.02929 -0.0016 0.9954 0.0125
-8.750 -0.9202 0.03351 0.02670 -0.0001 0.9936 0.0134
-8.500 -0.9020 0.03162 0.02443 0.0012 0.9916 0.0144
-8.250 -0.8816 0.02955 0.02199 0.0021 0.9899 0.0152
-8.000 -0.8599 0.02772 0.01995 0.0025 0.9884 0.0162
-7.750 -0.8353 0.02667 0.01877 0.0025 0.9869 0.0173
-7.500 -0.8096 0.02574 0.01765 0.0024 0.9858 0.0189
-7.250 -0.7877 0.02489 0.01662 0.0032 0.9835 0.0211
-7.000 -0.7630 0.02404 0.01556 0.0035 0.9815 0.0228
-6.750 -0.7401 0.02276 0.01425 0.0039 0.9797 0.0251
-6.500 -0.7144 0.02221 0.01364 0.0038 0.9780 0.0284
-6.250 -0.6875 0.02161 0.01289 0.0036 0.9766 0.0317
-6.000 -0.6658 0.02073 0.01196 0.0043 0.9745 0.0350
-5.750 -0.6452 0.02024 0.01143 0.0053 0.9713 0.0388
-5.500 -0.6214 0.01977 0.01085 0.0058 0.9689 0.0421
-5.250 -0.5973 0.01915 0.01020 0.0060 0.9669 0.0464
-5.000 -0.5703 0.01868 0.00969 0.0057 0.9653 0.0510
-4.500 -0.5291 0.01787 0.00879 0.0081 0.9590 0.0624
-4.250 -0.5051 0.01747 0.00838 0.0085 0.9564 0.0732
-4.000 -0.4796 0.01700 0.00801 0.0086 0.9542 0.0986
-3.750 -0.4548 0.01640 0.00764 0.0086 0.9525 0.1548
-3.500 -0.4393 0.01573 0.00733 0.0106 0.9490 0.2376
-3.250 -0.4306 0.01479 0.00698 0.0140 0.9441 0.3692
-3.000 -0.4252 0.01345 0.00661 0.0182 0.9405 0.5691
-2.750 -0.3953 0.01305 0.00710 0.0185 0.9401 0.7685
-2.500 -0.3515 0.01352 0.00767 0.0156 0.9401 0.8282
-2.250 -0.3171 0.01407 0.00820 0.0149 0.9391 0.8661
-2.000 -0.2716 0.01537 0.00947 0.0126 0.9394 0.8953
-1.750 -0.2222 0.01593 0.00995 0.0082 0.9397 0.8997
-1.250 -0.1613 0.01604 0.00992 0.0065 0.9356 0.9097
-1.000 -0.1289 0.01616 0.01000 0.0052 0.9325 0.9117
-0.750 -0.0952 0.01622 0.01002 0.0036 0.9301 0.9140
-0.500 -0.0625 0.01623 0.01001 0.0022 0.9278 0.9170
-0.250 -0.0396 0.01610 0.00986 0.0030 0.9249 0.9230
0.000 0.0000 0.01611 0.00986 0.0000 0.9239 0.9239
0.250 0.0396 0.01610 0.00986 -0.0029 0.9230 0.9249
0.500 0.0625 0.01623 0.01001 -0.0022 0.9170 0.9278
0.750 0.0952 0.01622 0.01002 -0.0036 0.9140 0.9301
1.000 0.1288 0.01616 0.01000 -0.0052 0.9117 0.9326
1.250 0.1612 0.01604 0.00992 -0.0065 0.9097 0.9356
1.500 0.1857 0.01603 0.00997 -0.0061 0.9038 0.9386
1.750 0.2222 0.01593 0.00995 -0.0082 0.8997 0.9398
2.000 0.2717 0.01536 0.00946 -0.0126 0.8953 0.9394
2.250 0.3175 0.01406 0.00820 -0.0150 0.8664 0.9390
2.500 0.3518 0.01351 0.00765 -0.0157 0.8277 0.9401
2.750 0.3953 0.01305 0.00710 -0.0185 0.7686 0.9401
3.000 0.4253 0.01345 0.00660 -0.0183 0.5703 0.9405
3.250 0.4306 0.01478 0.00698 -0.0140 0.3694 0.9442
3.500 0.4392 0.01572 0.00733 -0.0105 0.2383 0.9491
3.750 0.4548 0.01640 0.00764 -0.0086 0.1546 0.9525
4.000 0.4797 0.01700 0.00800 -0.0086 0.0988 0.9542
4.250 0.5051 0.01747 0.00838 -0.0085 0.0731 0.9564
4.500 0.5290 0.01787 0.00879 -0.0081 0.0624 0.9591
5.000 0.5704 0.01868 0.00968 -0.0058 0.0510 0.9653
5.250 0.5973 0.01915 0.01020 -0.0060 0.0464 0.9670
5.500 0.6214 0.01976 0.01084 -0.0058 0.0421 0.9690
5.750 0.6452 0.02024 0.01142 -0.0053 0.0388 0.9713
6.000 0.6657 0.02073 0.01195 -0.0043 0.0350 0.9746
6.250 0.6876 0.02161 0.01289 -0.0036 0.0317 0.9766
6.500 0.7145 0.02220 0.01363 -0.0038 0.0284 0.9780
6.750 0.7402 0.02277 0.01425 -0.0039 0.0251 0.9797
7.000 0.7631 0.02403 0.01555 -0.0036 0.0228 0.9815
7.250 0.7878 0.02488 0.01661 -0.0032 0.0211 0.9836
7.500 0.8098 0.02575 0.01766 -0.0024 0.0189 0.9858
7.750 0.8355 0.02667 0.01876 -0.0025 0.0173 0.9870
8.000 0.8601 0.02772 0.01995 -0.0026 0.0162 0.9884
8.250 0.8818 0.02956 0.02199 -0.0021 0.0152 0.9899
8.500 0.9021 0.03163 0.02443 -0.0012 0.0144 0.9916
8.750 0.9203 0.03351 0.02670 0.0000 0.0134 0.9936
9.000 0.9361 0.03571 0.02928 0.0015 0.0125 0.9954
9.250 0.9484 0.03866 0.03266 0.0033 0.0120 0.9975
9.500 0.9562 0.04198 0.03643 0.0054 0.0116 0.9999
9.750 0.9467 0.04444 0.03918 0.0112 0.0114 1.0000
10.000 0.9325 0.04691 0.04191 0.0175 0.0113 1.0000
10.250 0.9123 0.04927 0.04450 0.0247 0.0112 1.0000
10.500 0.8859 0.05126 0.04666 0.0327 0.0112 1.0000
10.750 0.8582 0.05365 0.04923 0.0401 0.0112 1.0000
11.000 0.8292 0.05659 0.05236 0.0465 0.0112 1.0000
11.250 0.7980 0.06024 0.05619 0.0518 0.0113 1.0000
11.500 0.7632 0.06498 0.06110 0.0555 0.0114 1.0000
11.750 0.7227 0.07172 0.06801 0.0566 0.0116 1.0000
12.000 0.6766 0.08284 0.07927 0.0522 0.0121 1.0000
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Polar data table (+)
Polar graphs
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