NACA M26 AIRFOIL (m26-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M26 AIRFOIL (m26-il) Reynolds number: 500,000 Max Cl/Cd: 115.74 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m26-il-500000.txt Download as CSV file: xf-m26-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M26 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2027 0.10044 0.09700 0.0017 0.5511 0.0134
-8.250 -0.1971 0.09738 0.09394 0.0008 0.5496 0.0136
-8.000 -0.2929 0.10678 0.10319 0.0110 0.5525 0.0131
-7.750 -0.2854 0.10368 0.10011 0.0098 0.5510 0.0133
-7.500 -0.2777 0.10080 0.09723 0.0081 0.5494 0.0135
-7.250 -0.2712 0.09812 0.09456 0.0063 0.5478 0.0137
-7.000 -0.2605 0.09519 0.09162 0.0039 0.5462 0.0141
-6.750 -0.2479 0.09215 0.08855 0.0012 0.5447 0.0146
-6.500 -0.2336 0.08901 0.08540 -0.0017 0.5433 0.0153
-6.250 -0.2149 0.08592 0.08226 -0.0055 0.5420 0.0161
-6.000 -0.1914 0.08314 0.07941 -0.0108 0.5406 0.0163
-5.750 -0.1656 0.08045 0.07662 -0.0163 0.5392 0.0164
-5.500 -0.1486 0.07626 0.07242 -0.0186 0.5381 0.0165
-5.250 -0.1370 0.07238 0.06854 -0.0186 0.5369 0.0167
-5.000 -0.1190 0.06937 0.06551 -0.0203 0.5356 0.0170
-4.750 -0.0977 0.06656 0.06267 -0.0226 0.5343 0.0175
-4.500 -0.0743 0.06376 0.05981 -0.0250 0.5329 0.0182
-4.250 -0.0484 0.06094 0.05691 -0.0276 0.5314 0.0194
-4.000 -0.0068 0.05901 0.05482 -0.0324 0.5298 0.0204
-3.750 0.0284 0.05662 0.05226 -0.0357 0.5284 0.0205
-3.500 0.0541 0.03407 0.03001 -0.0361 0.5257 0.0210
-3.250 0.0728 0.03142 0.02729 -0.0367 0.5244 0.0214
-3.000 0.0955 0.02907 0.02489 -0.0378 0.5232 0.0221
-2.750 0.1204 0.02679 0.02254 -0.0390 0.5221 0.0230
-2.500 0.1554 0.02538 0.02100 -0.0408 0.5207 0.0255
-2.250 0.1890 0.02379 0.01924 -0.0422 0.5193 0.0258
-2.000 0.2165 0.02181 0.01714 -0.0429 0.5179 0.0259
-1.750 0.2362 0.01828 0.01351 -0.0434 0.5168 0.0264
-1.500 0.2575 0.01658 0.01177 -0.0436 0.5155 0.0270
-1.250 0.2817 0.01523 0.01035 -0.0440 0.5142 0.0278
-1.000 0.3081 0.01400 0.00900 -0.0444 0.5129 0.0289
-0.750 0.3357 0.01287 0.00774 -0.0447 0.5115 0.0305
-0.500 0.3712 0.01269 0.00733 -0.0448 0.5102 0.0330
-0.250 0.3995 0.01189 0.00636 -0.0448 0.5087 0.0332
0.000 0.4375 0.02559 0.01953 -0.0462 0.5090 0.0340
0.250 0.4631 0.02438 0.01829 -0.0467 0.5078 0.0349
0.500 0.4906 0.02335 0.01722 -0.0470 0.5062 0.0360
0.750 0.5190 0.02238 0.01617 -0.0472 0.5042 0.0376
1.000 0.5486 0.02144 0.01512 -0.0472 0.5021 0.0405
1.250 0.5810 0.02138 0.01485 -0.0468 0.5003 0.0430
1.500 0.6064 0.01927 0.01266 -0.0472 0.4987 0.0460
1.750 0.6342 0.01873 0.01203 -0.0474 0.4971 0.0492
2.000 0.6652 0.01923 0.01232 -0.0472 0.4955 0.0551
2.250 0.6909 0.01762 0.01066 -0.0476 0.4938 0.0594
2.500 0.7196 0.01722 0.01024 -0.0478 0.4924 0.0652
2.750 0.7482 0.01655 0.00950 -0.0480 0.4907 0.0717
3.000 0.7764 0.01617 0.00915 -0.0483 0.4889 0.0769
3.250 0.8049 0.01570 0.00861 -0.0484 0.4870 0.0860
3.500 0.8333 0.01526 0.00813 -0.0486 0.4851 0.0894
3.750 0.8630 0.01366 0.00618 -0.0475 0.4835 0.0432
4.000 0.8895 0.01330 0.00577 -0.0473 0.4819 0.0442
4.250 0.9163 0.01317 0.00562 -0.0473 0.4802 0.0454
4.500 0.9435 0.01317 0.00564 -0.0473 0.4783 0.0455
4.750 0.9711 0.01307 0.00562 -0.0476 0.4763 0.0468
5.000 0.9988 0.01304 0.00565 -0.0478 0.4740 0.0499
5.250 1.0267 0.01296 0.00561 -0.0480 0.4711 0.0587
5.500 1.0542 0.01267 0.00549 -0.0481 0.4669 0.1647
5.750 1.1491 0.01116 0.00545 -0.0633 0.4558 1.0000
6.000 1.1748 0.01107 0.00529 -0.0630 0.4495 1.0000
6.250 1.2009 0.01094 0.00525 -0.0630 0.4414 1.0000
6.500 1.2268 0.01084 0.00515 -0.0629 0.4295 1.0000
6.750 1.2523 0.01082 0.00504 -0.0628 0.3988 1.0000
7.000 1.2709 0.01250 0.00606 -0.0631 0.2963 1.0000
7.250 1.2646 0.01735 0.00974 -0.0622 0.0965 1.0000
7.500 1.2652 0.01969 0.01181 -0.0602 0.0261 1.0000
7.750 1.2774 0.02064 0.01282 -0.0588 0.0217 1.0000
8.000 1.2868 0.02178 0.01405 -0.0575 0.0203 1.0000
8.250 1.2863 0.02341 0.01578 -0.0554 0.0197 1.0000
8.500 1.2831 0.02523 0.01769 -0.0529 0.0191 1.0000
8.750 1.2840 0.02729 0.01984 -0.0516 0.0185 1.0000
9.000 1.2852 0.02958 0.02224 -0.0507 0.0179 1.0000
9.250 1.2852 0.03212 0.02487 -0.0500 0.0174 1.0000
9.500 1.2836 0.03487 0.02771 -0.0493 0.0170 1.0000
9.750 1.2801 0.03785 0.03079 -0.0486 0.0165 1.0000
10.000 1.2743 0.04110 0.03413 -0.0479 0.0160 1.0000
10.250 1.2656 0.04466 0.03779 -0.0471 0.0156 1.0000
10.500 1.2539 0.04856 0.04179 -0.0464 0.0152 1.0000
10.750 1.2416 0.05262 0.04596 -0.0458 0.0149 1.0000
11.000 1.2404 0.05557 0.04898 -0.0454 0.0147 1.0000
11.250 1.2402 0.05847 0.05194 -0.0450 0.0146 1.0000
11.500 1.2395 0.06144 0.05498 -0.0447 0.0144 1.0000
11.750 1.2385 0.06447 0.05810 -0.0444 0.0142 1.0000
12.000 1.2381 0.06741 0.06110 -0.0440 0.0140 1.0000
12.250 1.2382 0.07030 0.06403 -0.0436 0.0139 1.0000
12.500 1.2393 0.07301 0.06680 -0.0432 0.0137 1.0000
12.750 1.2421 0.07539 0.06922 -0.0425 0.0136 1.0000
13.000 1.2473 0.07733 0.07119 -0.0417 0.0136 1.0000
13.250 1.2553 0.07886 0.07274 -0.0407 0.0134 1.0000
13.500 1.2645 0.08031 0.07422 -0.0398 0.0130 1.0000
13.750 1.2752 0.08144 0.07539 -0.0387 0.0126 1.0000
14.000 1.2945 0.08079 0.07476 -0.0360 0.0130 1.0000
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Polar data table (+)
Polar graphs
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