AG11 (ag11-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: AG11 (ag11-il) Reynolds number: 500,000 Max Cl/Cd: 75.71 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ag11-il-500000-n5.txt Download as CSV file: xf-ag11-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: AG11
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5584 0.11896 0.11666 0.0236 1.0000 0.0066
-9.250 -0.5537 0.11523 0.11294 0.0218 1.0000 0.0066
-8.500 -0.5414 0.10339 0.10116 0.0167 1.0000 0.0067
-8.250 -0.5358 0.10008 0.09786 0.0165 1.0000 0.0068
-8.000 -0.5306 0.09679 0.09459 0.0152 1.0000 0.0070
-7.750 -0.5255 0.09349 0.09131 0.0135 1.0000 0.0071
-7.500 -0.5205 0.09010 0.08794 0.0113 1.0000 0.0073
-7.250 -0.5141 0.08647 0.08433 0.0083 1.0000 0.0075
-7.000 -0.5035 0.08244 0.08031 0.0039 1.0000 0.0078
-6.750 -0.4903 0.07828 0.07615 -0.0008 1.0000 0.0083
-6.500 -0.4694 0.07317 0.07100 -0.0083 1.0000 0.0094
-6.250 -0.4471 0.06803 0.06581 -0.0149 1.0000 0.0096
-6.000 -0.4263 0.06334 0.06104 -0.0194 1.0000 0.0097
-5.750 -0.4048 0.05888 0.05651 -0.0230 1.0000 0.0097
-5.500 -0.3824 0.05454 0.05207 -0.0261 1.0000 0.0097
-5.250 -0.3594 0.05032 0.04774 -0.0285 1.0000 0.0097
-5.000 -0.3357 0.04626 0.04355 -0.0305 1.0000 0.0098
-4.750 -0.3117 0.04237 0.03951 -0.0319 1.0000 0.0098
-3.250 -0.1447 0.01947 0.01505 -0.0378 0.9799 0.0062
-3.000 -0.1101 0.01659 0.01178 -0.0389 0.9668 0.0060
-2.750 -0.0765 0.01403 0.00879 -0.0394 0.9504 0.0058
-2.500 -0.0460 0.01215 0.00654 -0.0393 0.9296 0.0058
-2.250 -0.0189 0.01095 0.00505 -0.0385 0.9046 0.0060
-2.000 0.0069 0.01019 0.00407 -0.0375 0.8766 0.0064
-1.750 0.0324 0.00955 0.00325 -0.0366 0.8466 0.0070
-1.500 0.0583 0.00926 0.00283 -0.0360 0.8151 0.0078
-1.250 0.0848 0.00914 0.00256 -0.0354 0.7829 0.0093
-1.000 0.1116 0.00889 0.00214 -0.0350 0.7505 0.0104
-0.750 0.1388 0.00870 0.00181 -0.0346 0.7186 0.0114
-0.500 0.1662 0.00862 0.00155 -0.0343 0.6868 0.0137
-0.250 0.1938 0.00851 0.00133 -0.0341 0.6555 0.0243
0.000 0.2212 0.00822 0.00124 -0.0341 0.6237 0.1293
0.250 0.2443 0.00647 0.00125 -0.0341 0.5944 0.7383
0.500 0.2715 0.00600 0.00119 -0.0332 0.5622 1.0000
0.750 0.2989 0.00618 0.00116 -0.0331 0.5283 1.0000
1.000 0.3263 0.00637 0.00116 -0.0329 0.4955 1.0000
1.250 0.3538 0.00656 0.00117 -0.0328 0.4627 1.0000
1.500 0.3812 0.00677 0.00120 -0.0327 0.4300 1.0000
1.750 0.4087 0.00698 0.00124 -0.0326 0.3995 1.0000
2.000 0.4361 0.00719 0.00131 -0.0326 0.3697 1.0000
2.250 0.4635 0.00742 0.00138 -0.0325 0.3395 1.0000
2.500 0.4909 0.00765 0.00148 -0.0324 0.3111 1.0000
2.750 0.5183 0.00789 0.00159 -0.0324 0.2841 1.0000
3.000 0.5455 0.00814 0.00170 -0.0323 0.2573 1.0000
3.250 0.5728 0.00839 0.00184 -0.0323 0.2335 1.0000
3.500 0.6001 0.00863 0.00198 -0.0322 0.2120 1.0000
3.750 0.6272 0.00890 0.00216 -0.0321 0.1902 1.0000
4.000 0.6545 0.00914 0.00233 -0.0321 0.1727 1.0000
4.250 0.6816 0.00940 0.00252 -0.0320 0.1570 1.0000
4.500 0.7086 0.00967 0.00272 -0.0319 0.1408 1.0000
4.750 0.7355 0.00998 0.00297 -0.0319 0.1242 1.0000
5.000 0.7624 0.01027 0.00321 -0.0318 0.1108 1.0000
5.250 0.7892 0.01056 0.00347 -0.0317 0.0995 1.0000
5.500 0.8160 0.01085 0.00374 -0.0316 0.0897 1.0000
5.750 0.8426 0.01118 0.00405 -0.0315 0.0801 1.0000
6.000 0.8690 0.01153 0.00438 -0.0313 0.0712 1.0000
6.250 0.8956 0.01183 0.00470 -0.0312 0.0657 1.0000
6.500 0.9218 0.01222 0.00510 -0.0310 0.0597 1.0000
6.750 0.9482 0.01253 0.00545 -0.0309 0.0559 1.0000
7.000 0.9742 0.01292 0.00585 -0.0307 0.0513 1.0000
7.250 1.0000 0.01335 0.00631 -0.0305 0.0476 1.0000
7.500 1.0260 0.01369 0.00674 -0.0303 0.0454 1.0000
7.750 1.0517 0.01410 0.00721 -0.0301 0.0426 1.0000
8.000 1.0769 0.01459 0.00772 -0.0298 0.0393 1.0000
8.250 1.1019 0.01510 0.00830 -0.0295 0.0369 1.0000
8.500 1.1271 0.01552 0.00882 -0.0293 0.0353 1.0000
8.750 1.1521 0.01598 0.00938 -0.0290 0.0330 1.0000
9.000 1.1765 0.01651 0.00995 -0.0287 0.0305 1.0000
9.250 1.1998 0.01726 0.01077 -0.0282 0.0281 1.0000
9.500 1.2243 0.01773 0.01135 -0.0279 0.0269 1.0000
9.750 1.2483 0.01826 0.01199 -0.0276 0.0252 1.0000
10.000 1.2718 0.01885 0.01266 -0.0272 0.0234 1.0000
10.250 1.2941 0.01961 0.01349 -0.0267 0.0212 1.0000
10.500 1.3170 0.02024 0.01424 -0.0263 0.0196 1.0000
10.750 1.3398 0.02082 0.01491 -0.0259 0.0176 1.0000
11.000 1.3614 0.02157 0.01571 -0.0254 0.0154 1.0000
11.250 1.3823 0.02241 0.01666 -0.0248 0.0139 1.0000
11.500 1.4030 0.02320 0.01757 -0.0242 0.0122 1.0000
11.750 1.4220 0.02420 0.01864 -0.0235 0.0103 1.0000
12.000 1.4405 0.02521 0.01981 -0.0227 0.0091 1.0000
12.250 1.4576 0.02634 0.02107 -0.0218 0.0078 1.0000
12.500 1.4720 0.02773 0.02258 -0.0208 0.0067 1.0000
12.750 1.4861 0.02905 0.02408 -0.0197 0.0059 1.0000
13.000 1.4970 0.03062 0.02580 -0.0184 0.0051 1.0000
13.250 1.5033 0.03253 0.02788 -0.0168 0.0045 1.0000
13.500 1.5062 0.03435 0.02989 -0.0148 0.0041 1.0000
13.750 1.5041 0.03644 0.03216 -0.0128 0.0038 1.0000
14.000 1.4991 0.03912 0.03502 -0.0117 0.0036 1.0000
14.250 1.4912 0.04261 0.03871 -0.0119 0.0034 1.0000
14.500 1.4804 0.04722 0.04352 -0.0138 0.0033 1.0000
14.750 1.4656 0.05335 0.04986 -0.0176 0.0033 1.0000
15.000 1.4454 0.06130 0.05803 -0.0231 0.0033 1.0000
15.250 1.4184 0.07107 0.06801 -0.0295 0.0034 1.0000
15.500 1.3829 0.08257 0.07970 -0.0363 0.0036 1.0000
15.750 1.3402 0.09530 0.09259 -0.0431 0.0040 1.0000
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Polar data table (+)
Polar graphs
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