NACA 0012 AIRFOILS (n0012-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA 0012 AIRFOILS (n0012-il) Reynolds number: 100,000 Max Cl/Cd: 36.68 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-n0012-il-100000.txt Download as CSV file: xf-n0012-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.750 -0.5002 0.13756 0.13233 -0.0001 1.0000 0.1541
-12.500 -0.5184 0.13358 0.12838 -0.0027 1.0000 0.1558
-12.250 -0.7724 0.10051 0.09513 -0.0153 1.0000 0.0866
-12.000 -0.7713 0.09532 0.08992 -0.0162 1.0000 0.0856
-11.750 -0.7896 0.08685 0.08144 -0.0207 1.0000 0.0842
-11.500 -0.8377 0.07511 0.06959 -0.0287 1.0000 0.0813
-11.250 -0.9695 0.06291 0.05654 -0.0300 1.0000 0.0753
-11.000 -0.9842 0.05891 0.05220 -0.0278 1.0000 0.0751
-10.750 -0.9915 0.05467 0.04764 -0.0258 1.0000 0.0753
-10.500 -0.9911 0.05060 0.04330 -0.0241 1.0000 0.0759
-10.250 -0.9765 0.04806 0.04077 -0.0233 1.0000 0.0778
-10.000 -0.9659 0.04582 0.03837 -0.0219 1.0000 0.0802
-9.750 -0.9597 0.04303 0.03523 -0.0199 1.0000 0.0822
-9.500 -0.9523 0.04019 0.03192 -0.0177 1.0000 0.0841
-9.250 -0.9430 0.03777 0.02894 -0.0153 1.0000 0.0861
-9.000 -0.9245 0.03526 0.02644 -0.0145 1.0000 0.0896
-8.750 -0.9056 0.03367 0.02473 -0.0134 1.0000 0.0938
-8.500 -0.8883 0.03189 0.02255 -0.0117 1.0000 0.0978
-8.250 -0.8673 0.02998 0.02066 -0.0109 1.0000 0.1029
-8.000 -0.8471 0.02869 0.01923 -0.0096 1.0000 0.1091
-7.750 -0.8256 0.02707 0.01754 -0.0086 1.0000 0.1157
-7.500 -0.8050 0.02598 0.01635 -0.0074 1.0000 0.1242
-7.250 -0.7836 0.02464 0.01512 -0.0064 1.0000 0.1333
-7.000 -0.7631 0.02348 0.01396 -0.0051 1.0000 0.1447
-6.750 -0.7430 0.02242 0.01293 -0.0038 1.0000 0.1584
-6.500 -0.7235 0.02143 0.01204 -0.0023 1.0000 0.1748
-6.250 -0.7047 0.02050 0.01124 -0.0008 1.0000 0.1953
-6.000 -0.6862 0.01958 0.01046 0.0008 1.0000 0.2203
-5.750 -0.6682 0.01875 0.00981 0.0025 1.0000 0.2509
-5.500 -0.6504 0.01798 0.00927 0.0042 1.0000 0.2860
-5.250 -0.6324 0.01733 0.00880 0.0060 1.0000 0.3256
-5.000 -0.6141 0.01674 0.00844 0.0077 1.0000 0.3656
-4.750 -0.5953 0.01627 0.00813 0.0095 1.0000 0.4067
-4.500 -0.5761 0.01585 0.00789 0.0112 1.0000 0.4463
-4.250 -0.5565 0.01550 0.00770 0.0129 1.0000 0.4854
-4.000 -0.5366 0.01520 0.00752 0.0145 1.0000 0.5246
-3.750 -0.5162 0.01493 0.00743 0.0162 1.0000 0.5615
-3.500 -0.4957 0.01471 0.00736 0.0179 1.0000 0.5981
-3.250 -0.4751 0.01453 0.00732 0.0196 1.0000 0.6345
-3.000 -0.4545 0.01440 0.00731 0.0214 1.0000 0.6707
-2.750 -0.4340 0.01432 0.00735 0.0232 1.0000 0.7062
-2.500 -0.4137 0.01429 0.00742 0.0252 1.0000 0.7415
-2.250 -0.3935 0.01433 0.00757 0.0272 1.0000 0.7754
-2.000 -0.3737 0.01444 0.00777 0.0294 1.0000 0.8089
-1.750 -0.3539 0.01463 0.00801 0.0317 1.0000 0.8421
-1.500 -0.3320 0.01489 0.00831 0.0335 1.0000 0.8748
-1.250 -0.3021 0.01527 0.00870 0.0338 1.0000 0.9052
-1.000 -0.2578 0.01574 0.00913 0.0309 1.0000 0.9317
-0.750 -0.2041 0.01619 0.00952 0.0259 1.0000 0.9557
-0.500 -0.1345 0.01658 0.00986 0.0174 1.0000 0.9712
-0.250 -0.0651 0.01682 0.01007 0.0086 1.0000 0.9860
0.000 0.0000 0.01693 0.01017 0.0000 1.0000 1.0000
0.250 0.0651 0.01682 0.01007 -0.0086 0.9860 1.0000
0.500 0.1344 0.01658 0.00986 -0.0174 0.9712 1.0000
0.750 0.2040 0.01619 0.00952 -0.0258 0.9557 1.0000
1.000 0.2578 0.01574 0.00913 -0.0309 0.9317 1.0000
1.250 0.3020 0.01527 0.00870 -0.0338 0.9053 1.0000
1.500 0.3320 0.01489 0.00831 -0.0335 0.8748 1.0000
1.750 0.3538 0.01462 0.00801 -0.0317 0.8421 1.0000
2.000 0.3737 0.01444 0.00777 -0.0294 0.8089 1.0000
2.250 0.3934 0.01433 0.00757 -0.0272 0.7754 1.0000
2.500 0.4136 0.01429 0.00742 -0.0252 0.7415 1.0000
2.750 0.4339 0.01432 0.00735 -0.0232 0.7063 1.0000
3.000 0.4544 0.01440 0.00731 -0.0214 0.6707 1.0000
3.250 0.4750 0.01453 0.00731 -0.0196 0.6346 1.0000
3.500 0.4957 0.01471 0.00736 -0.0179 0.5981 1.0000
3.750 0.5162 0.01493 0.00742 -0.0162 0.5615 1.0000
4.000 0.5365 0.01520 0.00752 -0.0145 0.5247 1.0000
4.250 0.5565 0.01549 0.00770 -0.0128 0.4854 1.0000
4.500 0.5761 0.01585 0.00789 -0.0111 0.4463 1.0000
4.750 0.5953 0.01627 0.00813 -0.0094 0.4068 1.0000
5.000 0.6141 0.01674 0.00843 -0.0077 0.3656 1.0000
5.250 0.6323 0.01732 0.00880 -0.0060 0.3257 1.0000
5.500 0.6503 0.01798 0.00927 -0.0042 0.2860 1.0000
5.750 0.6681 0.01874 0.00981 -0.0025 0.2509 1.0000
6.000 0.6862 0.01958 0.01046 -0.0008 0.2203 1.0000
6.250 0.7046 0.02050 0.01124 0.0008 0.1953 1.0000
6.500 0.7235 0.02143 0.01204 0.0023 0.1748 1.0000
6.750 0.7430 0.02242 0.01293 0.0038 0.1584 1.0000
7.000 0.7631 0.02348 0.01396 0.0051 0.1447 1.0000
7.250 0.7836 0.02464 0.01512 0.0064 0.1333 1.0000
7.500 0.8050 0.02598 0.01635 0.0074 0.1242 1.0000
7.750 0.8256 0.02706 0.01754 0.0086 0.1157 1.0000
8.000 0.8471 0.02869 0.01923 0.0096 0.1091 1.0000
8.250 0.8673 0.02998 0.02066 0.0109 0.1029 1.0000
8.500 0.8884 0.03189 0.02255 0.0117 0.0978 1.0000
8.750 0.9057 0.03367 0.02473 0.0134 0.0938 1.0000
9.000 0.9246 0.03526 0.02644 0.0145 0.0896 1.0000
9.250 0.9431 0.03777 0.02894 0.0153 0.0861 1.0000
9.500 0.9524 0.04019 0.03192 0.0176 0.0841 1.0000
9.750 0.9598 0.04303 0.03523 0.0198 0.0822 1.0000
10.000 0.9661 0.04582 0.03837 0.0219 0.0801 1.0000
10.250 0.9768 0.04806 0.04076 0.0233 0.0778 1.0000
10.500 0.9913 0.05061 0.04331 0.0241 0.0759 1.0000
10.750 0.9917 0.05469 0.04765 0.0257 0.0753 1.0000
11.000 0.9845 0.05893 0.05222 0.0278 0.0751 1.0000
11.250 0.9697 0.06293 0.05656 0.0300 0.0753 1.0000
11.500 0.8378 0.07522 0.06970 0.0285 0.0813 1.0000
11.750 0.7898 0.08703 0.08162 0.0204 0.0842 1.0000
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Polar data table (+)
Polar graphs
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