NACA 0010-65 (naca001065-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 0010-65 (naca001065-il) Reynolds number: 100,000 Max Cl/Cd: 28.28 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca001065-il-100000-n5.txt Download as CSV file: xf-naca001065-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 0010-65
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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6699 0.08957 0.08420 -0.0384 1.0000 0.0345
-11.000 -0.6927 0.08061 0.07523 -0.0442 1.0000 0.0341
-10.750 -0.7193 0.07409 0.06864 -0.0462 1.0000 0.0338
-10.500 -0.7468 0.06921 0.06369 -0.0451 1.0000 0.0336
-10.250 -0.7742 0.06542 0.05980 -0.0415 1.0000 0.0335
-10.000 -0.8012 0.06234 0.05659 -0.0361 1.0000 0.0335
-9.750 -0.8244 0.05879 0.05284 -0.0308 1.0000 0.0337
-9.500 -0.8446 0.05513 0.04890 -0.0254 1.0000 0.0341
-9.250 -0.8621 0.05142 0.04479 -0.0197 1.0000 0.0349
-9.000 -0.8758 0.04793 0.04078 -0.0140 1.0000 0.0358
-8.750 -0.8787 0.04511 0.03772 -0.0099 1.0000 0.0367
-8.500 -0.8737 0.04329 0.03577 -0.0068 1.0000 0.0378
-8.250 -0.8685 0.04126 0.03347 -0.0035 1.0000 0.0388
-8.000 -0.8608 0.03943 0.03139 -0.0004 1.0000 0.0402
-7.750 -0.8520 0.03763 0.02928 0.0026 1.0000 0.0423
-7.500 -0.8423 0.03550 0.02669 0.0058 1.0000 0.0443
-7.250 -0.8294 0.03342 0.02413 0.0085 1.0000 0.0457
-7.000 -0.8140 0.03157 0.02204 0.0105 1.0000 0.0474
-6.750 -0.7974 0.03052 0.02092 0.0122 1.0000 0.0496
-6.500 -0.7793 0.02935 0.01954 0.0138 1.0000 0.0517
-6.250 -0.7591 0.02801 0.01797 0.0153 1.0000 0.0534
-6.000 -0.7382 0.02684 0.01658 0.0166 1.0000 0.0553
-5.750 -0.7174 0.02600 0.01548 0.0179 1.0000 0.0578
-5.500 -0.6964 0.02483 0.01432 0.0190 1.0000 0.0602
-5.250 -0.6755 0.02395 0.01339 0.0202 1.0000 0.0621
-5.000 -0.6549 0.02317 0.01254 0.0215 1.0000 0.0644
-4.750 -0.6346 0.02244 0.01173 0.0229 1.0000 0.0670
-4.500 -0.6147 0.02180 0.01099 0.0243 1.0000 0.0699
-4.250 -0.5961 0.02111 0.01030 0.0259 1.0000 0.0739
-4.000 -0.5769 0.02060 0.00978 0.0274 1.0000 0.0804
-3.750 -0.5579 0.02004 0.00922 0.0290 1.0000 0.0877
-3.500 -0.5386 0.01953 0.00869 0.0305 1.0000 0.0982
-3.250 -0.5192 0.01901 0.00825 0.0320 1.0000 0.1144
-3.000 -0.4994 0.01853 0.00790 0.0334 1.0000 0.1417
-2.750 -0.4796 0.01802 0.00761 0.0346 1.0000 0.1865
-2.500 -0.4607 0.01726 0.00735 0.0359 1.0000 0.2829
-2.250 -0.3998 0.01650 0.00919 0.0321 1.0000 0.8756
-2.000 -0.3602 0.01733 0.00988 0.0308 1.0000 0.9192
-1.750 -0.2905 0.01827 0.01061 0.0230 1.0000 0.9437
-1.500 -0.1889 0.01927 0.01138 0.0087 1.0000 0.9738
-1.250 -0.1201 0.01941 0.01139 0.0000 1.0000 0.9867
-1.000 -0.0843 0.01933 0.01124 -0.0023 1.0000 0.9919
-0.750 -0.0524 0.01926 0.01113 -0.0039 1.0000 0.9962
-0.500 -0.0212 0.01921 0.01104 -0.0054 1.0000 1.0000
-0.250 -0.0106 0.01918 0.01100 -0.0027 1.0000 1.0000
0.000 0.0000 0.01917 0.01099 0.0000 1.0000 1.0000
0.250 0.0106 0.01918 0.01100 0.0027 1.0000 1.0000
0.500 0.0212 0.01921 0.01104 0.0054 1.0000 1.0000
0.750 0.0524 0.01926 0.01112 0.0040 0.9962 1.0000
1.000 0.0843 0.01932 0.01123 0.0024 0.9919 1.0000
1.250 0.1199 0.01940 0.01138 0.0000 0.9867 1.0000
1.500 0.1889 0.01927 0.01138 -0.0087 0.9737 1.0000
1.750 0.2904 0.01827 0.01061 -0.0230 0.9438 1.0000
2.000 0.3602 0.01732 0.00988 -0.0308 0.9193 1.0000
2.250 0.3999 0.01649 0.00919 -0.0321 0.8758 1.0000
2.500 0.4606 0.01725 0.00735 -0.0359 0.2834 1.0000
2.750 0.4795 0.01802 0.00761 -0.0346 0.1869 1.0000
3.000 0.4993 0.01852 0.00789 -0.0333 0.1418 1.0000
3.250 0.5191 0.01901 0.00824 -0.0320 0.1145 1.0000
3.500 0.5385 0.01953 0.00869 -0.0305 0.0983 1.0000
3.750 0.5578 0.02004 0.00921 -0.0290 0.0878 1.0000
4.000 0.5767 0.02060 0.00978 -0.0274 0.0804 1.0000
4.250 0.5959 0.02111 0.01029 -0.0259 0.0740 1.0000
4.500 0.6146 0.02180 0.01098 -0.0243 0.0699 1.0000
4.750 0.6345 0.02244 0.01173 -0.0228 0.0670 1.0000
5.000 0.6548 0.02316 0.01253 -0.0214 0.0644 1.0000
5.250 0.6754 0.02395 0.01338 -0.0201 0.0621 1.0000
5.500 0.6963 0.02482 0.01431 -0.0189 0.0602 1.0000
5.750 0.7173 0.02600 0.01547 -0.0179 0.0578 1.0000
6.000 0.7381 0.02684 0.01657 -0.0165 0.0553 1.0000
6.250 0.7590 0.02801 0.01797 -0.0152 0.0534 1.0000
6.500 0.7791 0.02934 0.01953 -0.0138 0.0517 1.0000
6.750 0.7972 0.03051 0.02092 -0.0121 0.0496 1.0000
7.000 0.8139 0.03157 0.02203 -0.0105 0.0474 1.0000
7.250 0.8293 0.03342 0.02412 -0.0085 0.0457 1.0000
7.500 0.8422 0.03549 0.02668 -0.0057 0.0443 1.0000
7.750 0.8519 0.03762 0.02928 -0.0026 0.0423 1.0000
8.000 0.8607 0.03942 0.03139 0.0004 0.0402 1.0000
8.250 0.8684 0.04125 0.03346 0.0035 0.0388 1.0000
8.500 0.8737 0.04328 0.03576 0.0068 0.0378 1.0000
8.750 0.8786 0.04512 0.03774 0.0099 0.0367 1.0000
9.000 0.8759 0.04792 0.04077 0.0140 0.0358 1.0000
9.250 0.8622 0.05140 0.04477 0.0197 0.0349 1.0000
9.500 0.8447 0.05512 0.04889 0.0254 0.0341 1.0000
9.750 0.8245 0.05879 0.05284 0.0308 0.0337 1.0000
10.000 0.8013 0.06235 0.05660 0.0361 0.0335 1.0000
10.250 0.7743 0.06544 0.05981 0.0415 0.0335 1.0000
10.500 0.7469 0.06924 0.06372 0.0450 0.0336 1.0000
10.750 0.7196 0.07414 0.06869 0.0461 0.0338 1.0000
11.000 0.6931 0.08067 0.07529 0.0441 0.0341 1.0000
11.250 0.6704 0.08969 0.08431 0.0382 0.0345 1.0000
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Polar data table (+)
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