AG12 (ag12-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: AG12 (ag12-il) Reynolds number: 50,000 Max Cl/Cd: 33.19 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-ag12-il-50000.txt Download as CSV file: xf-ag12-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: AG12
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.5444 0.10114 0.09472 0.0178 1.0000 0.2148
-7.250 -0.5505 0.09947 0.09317 0.0137 1.0000 0.2251
-7.000 -0.5426 0.09576 0.08951 0.0130 1.0000 0.2387
-6.750 -0.5365 0.09238 0.08618 0.0108 1.0000 0.2528
-6.500 -0.5245 0.08801 0.08186 0.0121 1.0000 0.2675
-6.000 -0.5015 0.08019 0.07412 0.0128 1.0000 0.2977
-5.750 -0.4901 0.07658 0.07055 0.0127 1.0000 0.3155
-5.500 -0.4816 0.07319 0.06716 0.0108 1.0000 0.3374
-5.250 -0.4678 0.06947 0.06347 0.0119 1.0000 0.3567
-5.000 -0.4563 0.06596 0.06001 0.0117 1.0000 0.3808
-4.750 -0.3578 0.04183 0.03383 -0.0426 1.0000 0.1174
-4.500 -0.3295 0.03730 0.02890 -0.0440 1.0000 0.1118
-4.250 -0.2992 0.03327 0.02412 -0.0455 1.0000 0.1136
-4.000 -0.2704 0.02998 0.02038 -0.0461 1.0000 0.1197
-3.750 -0.2405 0.02700 0.01690 -0.0461 1.0000 0.1250
-3.500 -0.2114 0.02466 0.01415 -0.0459 1.0000 0.1419
-3.250 -0.1836 0.02247 0.01188 -0.0453 1.0000 0.1679
-3.000 -0.1570 0.02060 0.01018 -0.0445 1.0000 0.2227
-2.750 -0.1310 0.01872 0.00881 -0.0434 1.0000 0.3298
-2.500 -0.1093 0.01701 0.00792 -0.0412 1.0000 0.5032
-2.250 -0.0960 0.01544 0.00726 -0.0364 1.0000 0.6829
-2.000 -0.0640 0.01387 0.00601 -0.0351 1.0000 1.0000
-1.750 -0.0335 0.01382 0.00534 -0.0363 1.0000 1.0000
-1.500 -0.0058 0.01380 0.00490 -0.0366 1.0000 1.0000
-1.250 0.0209 0.01381 0.00454 -0.0366 1.0000 1.0000
-1.000 0.0470 0.01384 0.00430 -0.0365 1.0000 1.0000
-0.750 0.0727 0.01389 0.00414 -0.0363 1.0000 1.0000
-0.500 0.0982 0.01396 0.00405 -0.0362 1.0000 1.0000
-0.250 0.1233 0.01406 0.00401 -0.0360 1.0000 1.0000
0.000 0.1482 0.01418 0.00401 -0.0358 1.0000 1.0000
0.250 0.1728 0.01434 0.00410 -0.0356 1.0000 1.0000
0.500 0.1969 0.01455 0.00426 -0.0355 1.0000 1.0000
0.750 0.2206 0.01480 0.00450 -0.0355 1.0000 1.0000
1.000 0.2439 0.01513 0.00484 -0.0356 1.0000 1.0000
1.250 0.2667 0.01556 0.00528 -0.0359 1.0000 1.0000
1.500 0.2886 0.01611 0.00586 -0.0365 1.0000 1.0000
1.750 0.3094 0.01683 0.00664 -0.0373 1.0000 1.0000
2.000 0.3288 0.01776 0.00763 -0.0386 1.0000 1.0000
2.250 0.4211 0.01852 0.00871 -0.0523 0.9556 1.0000
2.500 0.5003 0.01864 0.00915 -0.0612 0.9066 1.0000
2.750 0.5527 0.01855 0.00926 -0.0634 0.8589 1.0000
3.000 0.5824 0.01869 0.00954 -0.0614 0.8097 1.0000
3.250 0.6062 0.01890 0.00974 -0.0580 0.7634 1.0000
3.500 0.6272 0.01928 0.01008 -0.0547 0.7154 1.0000
3.750 0.6484 0.01970 0.01042 -0.0515 0.6686 1.0000
4.000 0.6700 0.02024 0.01094 -0.0486 0.6202 1.0000
4.250 0.6920 0.02085 0.01143 -0.0460 0.5722 1.0000
4.500 0.7143 0.02158 0.01206 -0.0438 0.5229 1.0000
4.750 0.7367 0.02236 0.01268 -0.0416 0.4748 1.0000
5.000 0.7593 0.02330 0.01354 -0.0398 0.4251 1.0000
5.250 0.7820 0.02433 0.01436 -0.0380 0.3786 1.0000
5.500 0.8047 0.02557 0.01563 -0.0366 0.3318 1.0000
5.750 0.8277 0.02699 0.01699 -0.0352 0.2894 1.0000
6.000 0.8503 0.02851 0.01843 -0.0339 0.2497 1.0000
6.250 0.8735 0.03040 0.02030 -0.0327 0.2167 1.0000
6.500 0.8964 0.03249 0.02232 -0.0316 0.1876 1.0000
6.750 0.9185 0.03502 0.02516 -0.0306 0.1631 1.0000
7.000 0.9398 0.03808 0.02850 -0.0296 0.1444 1.0000
7.250 0.9597 0.04177 0.03270 -0.0286 0.1312 1.0000
7.500 0.9768 0.04521 0.03653 -0.0278 0.1176 1.0000
7.750 0.9888 0.05069 0.04279 -0.0273 0.1137 1.0000
8.000 0.9972 0.05643 0.04910 -0.0271 0.1116 1.0000
8.250 1.0077 0.06103 0.05385 -0.0267 0.1060 1.0000
8.500 1.0042 0.06748 0.06082 -0.0278 0.1055 1.0000
8.750 0.9980 0.07397 0.06763 -0.0293 0.1058 1.0000
9.000 0.9915 0.08037 0.07420 -0.0310 0.1065 1.0000
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Polar data table (+)
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