NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY 64-012 AIRFOIL (n64012-il) Reynolds number: 100,000 Max Cl/Cd: 32.85 at α=4.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64012-il-100000-n5.txt Download as CSV file: xf-n64012-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY 64-012 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.000 -0.7831 0.08140 0.07610 -0.0201 1.0000 0.0288
-11.750 -0.8111 0.07422 0.06874 -0.0245 1.0000 0.0287
-11.500 -0.8385 0.06822 0.06250 -0.0272 1.0000 0.0285
-11.250 -0.8540 0.06415 0.05827 -0.0280 1.0000 0.0286
-11.000 -0.8687 0.06052 0.05445 -0.0279 1.0000 0.0287
-10.750 -0.8773 0.05770 0.05151 -0.0270 1.0000 0.0292
-10.500 -0.8838 0.05514 0.04880 -0.0255 1.0000 0.0296
-10.250 -0.8872 0.05252 0.04598 -0.0240 1.0000 0.0300
-10.000 -0.8869 0.04983 0.04306 -0.0226 1.0000 0.0305
-9.750 -0.8849 0.04685 0.03977 -0.0211 1.0000 0.0309
-9.500 -0.8791 0.04390 0.03648 -0.0197 1.0000 0.0312
-9.250 -0.8695 0.04112 0.03335 -0.0184 1.0000 0.0317
-9.000 -0.8564 0.03855 0.03044 -0.0173 1.0000 0.0322
-8.750 -0.8402 0.03619 0.02773 -0.0163 1.0000 0.0329
-8.500 -0.8215 0.03405 0.02528 -0.0154 1.0000 0.0338
-8.250 -0.8014 0.03233 0.02325 -0.0147 1.0000 0.0354
-8.000 -0.7798 0.03051 0.02122 -0.0141 1.0000 0.0369
-7.750 -0.7577 0.02883 0.01953 -0.0137 1.0000 0.0384
-7.500 -0.7353 0.02741 0.01804 -0.0131 1.0000 0.0399
-7.250 -0.7132 0.02612 0.01669 -0.0125 1.0000 0.0416
-7.000 -0.6918 0.02497 0.01546 -0.0117 1.0000 0.0437
-6.750 -0.6707 0.02407 0.01442 -0.0108 1.0000 0.0465
-6.500 -0.6538 0.02288 0.01334 -0.0096 1.0000 0.0497
-6.250 -0.6366 0.02200 0.01244 -0.0083 1.0000 0.0531
-6.000 -0.6207 0.02124 0.01161 -0.0066 1.0000 0.0569
-5.750 -0.6092 0.02045 0.01088 -0.0044 1.0000 0.0620
-5.500 -0.5910 0.01978 0.01014 -0.0034 0.9947 0.0693
-5.250 -0.5583 0.01878 0.00918 -0.0053 0.9821 0.0846
-5.000 -0.5276 0.01759 0.00826 -0.0070 0.9700 0.1269
-4.750 -0.5039 0.01573 0.00730 -0.0081 0.9573 0.2818
-4.500 -0.4812 0.01464 0.00718 -0.0079 0.9452 0.4807
-4.250 -0.4511 0.01455 0.00716 -0.0081 0.9347 0.5449
-4.000 -0.4243 0.01463 0.00731 -0.0075 0.9230 0.5941
-3.750 -0.3984 0.01490 0.00762 -0.0063 0.9119 0.6362
-3.500 -0.3719 0.01519 0.00787 -0.0052 0.9019 0.6653
-3.250 -0.3455 0.01532 0.00794 -0.0044 0.8915 0.6819
-3.000 -0.3191 0.01525 0.00773 -0.0040 0.8809 0.6922
-2.750 -0.2917 0.01523 0.00761 -0.0036 0.8718 0.6991
-2.500 -0.2655 0.01515 0.00741 -0.0033 0.8619 0.7081
-2.250 -0.2391 0.01512 0.00732 -0.0028 0.8524 0.7151
-2.000 -0.2126 0.01502 0.00711 -0.0025 0.8443 0.7230
-1.750 -0.1860 0.01499 0.00702 -0.0021 0.8345 0.7288
-1.500 -0.1594 0.01491 0.00684 -0.0019 0.8263 0.7368
-1.250 -0.1327 0.01488 0.00679 -0.0015 0.8174 0.7425
-1.000 -0.1061 0.01483 0.00668 -0.0013 0.8094 0.7497
-0.750 -0.0795 0.01480 0.00663 -0.0009 0.8015 0.7564
-0.500 -0.0531 0.01479 0.00659 -0.0006 0.7935 0.7638
-0.250 -0.0265 0.01476 0.00654 -0.0003 0.7858 0.7709
0.000 0.0000 0.01476 0.00655 0.0000 0.7781 0.7780
0.250 0.0266 0.01476 0.00654 0.0003 0.7709 0.7858
0.500 0.0531 0.01479 0.00659 0.0006 0.7638 0.7935
0.750 0.0796 0.01480 0.00663 0.0009 0.7563 0.8015
1.000 0.1061 0.01483 0.00668 0.0013 0.7497 0.8094
1.250 0.1327 0.01488 0.00679 0.0015 0.7425 0.8174
1.500 0.1594 0.01491 0.00684 0.0019 0.7368 0.8264
1.750 0.1860 0.01499 0.00702 0.0021 0.7288 0.8345
2.000 0.2126 0.01502 0.00711 0.0025 0.7230 0.8443
2.250 0.2391 0.01512 0.00732 0.0028 0.7151 0.8524
2.500 0.2656 0.01515 0.00741 0.0032 0.7081 0.8619
2.750 0.2917 0.01523 0.00761 0.0036 0.6991 0.8718
3.000 0.3192 0.01525 0.00773 0.0040 0.6922 0.8809
3.250 0.3455 0.01532 0.00794 0.0044 0.6819 0.8915
3.500 0.3719 0.01519 0.00787 0.0052 0.6653 0.9019
3.750 0.3984 0.01490 0.00762 0.0063 0.6361 0.9119
4.000 0.4243 0.01463 0.00731 0.0075 0.5942 0.9230
4.250 0.4511 0.01455 0.00716 0.0081 0.5449 0.9347
4.500 0.4812 0.01465 0.00718 0.0079 0.4797 0.9452
4.750 0.5038 0.01575 0.00730 0.0081 0.2796 0.9573
5.000 0.5276 0.01759 0.00826 0.0070 0.1270 0.9700
5.250 0.5583 0.01878 0.00918 0.0053 0.0845 0.9821
5.500 0.5910 0.01978 0.01014 0.0034 0.0693 0.9948
5.750 0.6092 0.02045 0.01088 0.0044 0.0620 1.0000
6.000 0.6207 0.02124 0.01161 0.0066 0.0569 1.0000
6.250 0.6366 0.02200 0.01243 0.0083 0.0531 1.0000
6.500 0.6538 0.02288 0.01334 0.0096 0.0497 1.0000
6.750 0.6708 0.02407 0.01442 0.0108 0.0465 1.0000
7.000 0.6919 0.02497 0.01546 0.0117 0.0437 1.0000
7.250 0.7133 0.02613 0.01669 0.0125 0.0416 1.0000
7.500 0.7354 0.02742 0.01804 0.0131 0.0398 1.0000
7.750 0.7578 0.02883 0.01953 0.0137 0.0384 1.0000
8.000 0.7800 0.03053 0.02124 0.0140 0.0369 1.0000
8.250 0.8015 0.03232 0.02324 0.0146 0.0354 1.0000
8.500 0.8216 0.03405 0.02528 0.0154 0.0337 1.0000
8.750 0.8404 0.03619 0.02773 0.0162 0.0329 1.0000
9.000 0.8566 0.03855 0.03044 0.0172 0.0322 1.0000
9.250 0.8697 0.04113 0.03336 0.0184 0.0317 1.0000
9.500 0.8793 0.04391 0.03649 0.0196 0.0312 1.0000
9.750 0.8851 0.04688 0.03981 0.0211 0.0309 1.0000
10.000 0.8868 0.04993 0.04317 0.0226 0.0306 1.0000
10.250 0.8868 0.05264 0.04612 0.0240 0.0300 1.0000
10.500 0.8856 0.05504 0.04868 0.0254 0.0295 1.0000
10.750 0.8798 0.05751 0.05128 0.0268 0.0290 1.0000
11.000 0.8665 0.06083 0.05480 0.0278 0.0289 1.0000
11.250 0.8547 0.06419 0.05831 0.0279 0.0286 1.0000
11.500 0.8380 0.06840 0.06268 0.0270 0.0285 1.0000
11.750 0.8150 0.07386 0.06835 0.0246 0.0286 1.0000
12.000 0.7258 0.09393 0.08886 0.0104 0.0313 1.0000
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Polar data table (+)
Polar graphs
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