NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA 16-012 (naca16012-il) Reynolds number: 100,000 Max Cl/Cd: 30.81 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-naca16012-il-100000-n5.txt Download as CSV file: xf-naca16012-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.750 -0.6483 0.08781 0.08263 -0.0524 1.0000 0.0227
-11.500 -0.6695 0.08196 0.07673 -0.0545 1.0000 0.0225
-11.250 -0.6923 0.07712 0.07183 -0.0548 1.0000 0.0224
-11.000 -0.7161 0.07307 0.06769 -0.0535 1.0000 0.0223
-10.750 -0.7402 0.06962 0.06414 -0.0506 1.0000 0.0222
-10.500 -0.7643 0.06668 0.06108 -0.0465 1.0000 0.0221
-10.250 -0.7887 0.06418 0.05846 -0.0411 1.0000 0.0221
-10.000 -0.8132 0.06208 0.05623 -0.0347 1.0000 0.0221
-9.750 -0.8361 0.05990 0.05390 -0.0279 1.0000 0.0221
-9.500 -0.8546 0.05721 0.05099 -0.0218 1.0000 0.0221
-9.250 -0.8701 0.05441 0.04792 -0.0157 1.0000 0.0222
-9.000 -0.8824 0.05154 0.04473 -0.0099 1.0000 0.0223
-8.750 -0.8911 0.04867 0.04147 -0.0043 1.0000 0.0225
-8.500 -0.8957 0.04591 0.03829 0.0009 1.0000 0.0228
-8.250 -0.8968 0.04301 0.03509 0.0053 1.0000 0.0234
-8.000 -0.8919 0.04127 0.03322 0.0085 1.0000 0.0245
-7.750 -0.8844 0.03986 0.03162 0.0116 1.0000 0.0259
-7.500 -0.8750 0.03803 0.02946 0.0147 1.0000 0.0274
-7.250 -0.8626 0.03579 0.02680 0.0175 1.0000 0.0287
-7.000 -0.8462 0.03364 0.02423 0.0198 1.0000 0.0303
-6.750 -0.8282 0.03207 0.02226 0.0218 1.0000 0.0328
-6.500 -0.8104 0.03051 0.02066 0.0231 1.0000 0.0355
-6.250 -0.7905 0.02920 0.01919 0.0245 1.0000 0.0382
-6.000 -0.7706 0.02818 0.01793 0.0260 1.0000 0.0422
-5.750 -0.7515 0.02688 0.01659 0.0273 1.0000 0.0461
-5.500 -0.7336 0.02597 0.01560 0.0290 1.0000 0.0500
-5.250 -0.7156 0.02533 0.01477 0.0308 1.0000 0.0552
-5.000 -0.7002 0.02434 0.01381 0.0328 1.0000 0.0594
-4.750 -0.6820 0.02371 0.01311 0.0344 0.9996 0.0654
-4.500 -0.6577 0.02299 0.01233 0.0347 0.9975 0.0726
-4.250 -0.6324 0.02239 0.01161 0.0348 0.9953 0.0809
-4.000 -0.6077 0.02175 0.01098 0.0350 0.9930 0.0948
-3.750 -0.5840 0.02105 0.01045 0.0353 0.9905 0.1282
-3.500 -0.5636 0.01987 0.01000 0.0360 0.9883 0.2548
-3.250 -0.5520 0.01792 0.00970 0.0387 0.9867 0.5632
-3.000 -0.4865 0.01903 0.01168 0.0338 0.9906 0.8188
-2.750 -0.4508 0.02034 0.01288 0.0339 0.9895 0.8789
-2.500 -0.3719 0.02300 0.01530 0.0264 0.9949 0.9239
-2.250 -0.3225 0.02364 0.01575 0.0221 0.9952 0.9378
-2.000 -0.2801 0.02373 0.01566 0.0185 0.9940 0.9420
-1.750 -0.2449 0.02377 0.01557 0.0163 0.9919 0.9476
-1.500 -0.2113 0.02379 0.01547 0.0143 0.9895 0.9529
-1.250 -0.1726 0.02380 0.01539 0.0112 0.9876 0.9562
-1.000 -0.1369 0.02383 0.01534 0.0087 0.9856 0.9604
-0.750 -0.1069 0.02381 0.01526 0.0075 0.9823 0.9649
-0.500 -0.0696 0.02379 0.01520 0.0046 0.9796 0.9675
-0.250 -0.0330 0.02380 0.01518 0.0019 0.9773 0.9707
0.000 0.0000 0.02382 0.01519 0.0000 0.9744 0.9744
0.250 0.0330 0.02380 0.01518 -0.0019 0.9707 0.9773
0.500 0.0695 0.02378 0.01520 -0.0046 0.9675 0.9796
0.750 0.1068 0.02381 0.01526 -0.0074 0.9650 0.9823
1.000 0.1369 0.02383 0.01534 -0.0087 0.9605 0.9856
1.250 0.1726 0.02379 0.01538 -0.0112 0.9562 0.9876
1.500 0.2113 0.02378 0.01546 -0.0143 0.9529 0.9895
1.750 0.2450 0.02376 0.01556 -0.0163 0.9477 0.9919
2.000 0.2801 0.02373 0.01566 -0.0185 0.9420 0.9940
2.250 0.3225 0.02363 0.01574 -0.0221 0.9379 0.9952
2.500 0.3720 0.02299 0.01529 -0.0265 0.9239 0.9949
2.750 0.4509 0.02033 0.01288 -0.0339 0.8790 0.9896
3.000 0.4866 0.01902 0.01167 -0.0338 0.8186 0.9906
3.250 0.5521 0.01792 0.00970 -0.0387 0.5633 0.9867
3.500 0.5637 0.01986 0.01000 -0.0360 0.2554 0.9884
3.750 0.5840 0.02104 0.01045 -0.0353 0.1282 0.9905
4.000 0.6078 0.02175 0.01097 -0.0350 0.0949 0.9930
4.250 0.6324 0.02239 0.01161 -0.0348 0.0809 0.9954
4.500 0.6577 0.02298 0.01232 -0.0347 0.0726 0.9975
4.750 0.6821 0.02370 0.01310 -0.0344 0.0654 0.9996
5.000 0.7001 0.02434 0.01380 -0.0328 0.0594 1.0000
5.250 0.7154 0.02532 0.01476 -0.0307 0.0553 1.0000
5.500 0.7335 0.02596 0.01559 -0.0290 0.0500 1.0000
5.750 0.7514 0.02687 0.01658 -0.0273 0.0461 1.0000
6.000 0.7705 0.02817 0.01792 -0.0260 0.0422 1.0000
6.250 0.7904 0.02919 0.01918 -0.0245 0.0382 1.0000
6.500 0.8103 0.03050 0.02065 -0.0231 0.0355 1.0000
6.750 0.8281 0.03204 0.02223 -0.0217 0.0328 1.0000
7.000 0.8461 0.03363 0.02422 -0.0198 0.0303 1.0000
7.250 0.8625 0.03578 0.02679 -0.0175 0.0287 1.0000
7.500 0.8749 0.03803 0.02945 -0.0146 0.0274 1.0000
7.750 0.8844 0.03984 0.03160 -0.0116 0.0259 1.0000
8.000 0.8918 0.04128 0.03322 -0.0085 0.0245 1.0000
8.250 0.8968 0.04299 0.03507 -0.0052 0.0234 1.0000
8.500 0.8956 0.04590 0.03827 -0.0009 0.0228 1.0000
8.750 0.8910 0.04866 0.04146 0.0043 0.0225 1.0000
9.000 0.8825 0.05153 0.04471 0.0099 0.0223 1.0000
9.250 0.8702 0.05440 0.04791 0.0157 0.0222 1.0000
9.500 0.8547 0.05720 0.05098 0.0218 0.0221 1.0000
9.750 0.8362 0.05989 0.05389 0.0279 0.0221 1.0000
10.000 0.8134 0.06208 0.05623 0.0346 0.0221 1.0000
10.250 0.7887 0.06419 0.05847 0.0411 0.0221 1.0000
10.500 0.7645 0.06669 0.06108 0.0464 0.0221 1.0000
10.750 0.7404 0.06962 0.06414 0.0506 0.0222 1.0000
11.000 0.7164 0.07306 0.06768 0.0534 0.0223 1.0000
11.250 0.6927 0.07712 0.07183 0.0548 0.0224 1.0000
11.500 0.6698 0.08198 0.07675 0.0545 0.0225 1.0000
11.750 0.6486 0.08787 0.08269 0.0523 0.0227 1.0000
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Polar data table (+)
Polar graphs
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