NACA M12 AIRFOIL (m12-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M12 AIRFOIL (m12-il) Reynolds number: 50,000 Max Cl/Cd: 33.59 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m12-il-50000-n5.txt Download as CSV file: xf-m12-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M12 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.250 -0.5339 0.10217 0.09456 -0.0281 1.0000 0.0752
-10.000 -0.5327 0.09771 0.09012 -0.0297 1.0000 0.0746
-9.750 -0.5351 0.09268 0.08512 -0.0323 1.0000 0.0740
-9.500 -0.5422 0.08684 0.07930 -0.0361 1.0000 0.0733
-9.250 -0.5557 0.08110 0.07360 -0.0394 1.0000 0.0728
-9.000 -0.5729 0.07622 0.06870 -0.0405 1.0000 0.0724
-8.750 -0.5879 0.07152 0.06391 -0.0409 1.0000 0.0723
-8.500 -0.6003 0.06707 0.05932 -0.0405 1.0000 0.0725
-8.250 -0.6096 0.06290 0.05494 -0.0395 1.0000 0.0728
-8.000 -0.6157 0.05898 0.05076 -0.0379 1.0000 0.0732
-7.750 -0.6183 0.05534 0.04682 -0.0360 1.0000 0.0735
-7.500 -0.6170 0.05203 0.04319 -0.0338 1.0000 0.0737
-7.250 -0.6131 0.04896 0.03978 -0.0316 1.0000 0.0739
-7.000 -0.6070 0.04610 0.03657 -0.0293 1.0000 0.0742
-6.750 -0.5986 0.04347 0.03355 -0.0270 1.0000 0.0747
-6.500 -0.5886 0.04108 0.03073 -0.0247 1.0000 0.0755
-6.250 -0.5756 0.03922 0.02876 -0.0229 1.0000 0.0770
-6.000 -0.5614 0.03776 0.02720 -0.0212 1.0000 0.0790
-5.750 -0.5465 0.03629 0.02555 -0.0194 1.0000 0.0810
-5.500 -0.5306 0.03473 0.02374 -0.0177 1.0000 0.0827
-5.250 -0.5134 0.03323 0.02198 -0.0161 1.0000 0.0842
-5.000 -0.4950 0.03186 0.02035 -0.0146 1.0000 0.0860
-4.750 -0.4756 0.03065 0.01885 -0.0132 1.0000 0.0880
-4.500 -0.4566 0.02964 0.01782 -0.0120 1.0000 0.0909
-4.250 -0.4380 0.02889 0.01707 -0.0108 1.0000 0.0952
-4.000 -0.4180 0.02816 0.01620 -0.0096 1.0000 0.1003
-3.750 -0.3819 0.02722 0.01522 -0.0114 0.9944 0.1067
-3.500 -0.3423 0.02638 0.01424 -0.0138 0.9867 0.1169
-3.250 -0.3039 0.02556 0.01348 -0.0163 0.9785 0.1346
-3.000 -0.2652 0.02466 0.01269 -0.0190 0.9706 0.1657
-2.750 -0.2310 0.02360 0.01205 -0.0210 0.9614 0.2236
-2.500 -0.2036 0.02204 0.01158 -0.0220 0.9518 0.3885
-2.250 -0.1594 0.02102 0.01241 -0.0216 0.9490 0.8115
-2.000 -0.0845 0.02170 0.01278 -0.0285 0.9476 0.9352
-1.750 0.0202 0.02188 0.01253 -0.0430 0.9506 0.9932
-1.500 0.0735 0.02171 0.01209 -0.0488 0.9416 1.0000
-1.250 0.1133 0.02157 0.01174 -0.0517 0.9294 1.0000
-1.000 0.1475 0.02147 0.01148 -0.0535 0.9153 1.0000
-0.750 0.1802 0.02139 0.01126 -0.0549 0.9012 1.0000
-0.500 0.2113 0.02134 0.01108 -0.0558 0.8871 1.0000
-0.250 0.2409 0.02130 0.01093 -0.0563 0.8730 1.0000
0.000 0.2693 0.02128 0.01081 -0.0566 0.8591 1.0000
0.250 0.2968 0.02126 0.01070 -0.0565 0.8455 1.0000
0.500 0.3234 0.02126 0.01062 -0.0562 0.8322 1.0000
0.750 0.3466 0.02134 0.01063 -0.0553 0.8176 1.0000
1.000 0.3693 0.02143 0.01067 -0.0543 0.8032 1.0000
1.250 0.3919 0.02154 0.01072 -0.0533 0.7891 1.0000
1.500 0.4146 0.02165 0.01080 -0.0522 0.7754 1.0000
1.750 0.4379 0.02176 0.01086 -0.0512 0.7622 1.0000
2.000 0.4624 0.02183 0.01090 -0.0502 0.7497 1.0000
2.250 0.4851 0.02198 0.01103 -0.0491 0.7361 1.0000
2.500 0.5068 0.02217 0.01121 -0.0478 0.7220 1.0000
2.750 0.5288 0.02237 0.01141 -0.0466 0.7080 1.0000
3.000 0.5510 0.02255 0.01159 -0.0453 0.6941 1.0000
3.250 0.5736 0.02272 0.01176 -0.0441 0.6803 1.0000
3.500 0.5964 0.02287 0.01192 -0.0428 0.6665 1.0000
3.750 0.6195 0.02301 0.01206 -0.0415 0.6526 1.0000
4.000 0.6418 0.02319 0.01226 -0.0402 0.6382 1.0000
4.250 0.6631 0.02344 0.01256 -0.0388 0.6233 1.0000
4.500 0.6845 0.02370 0.01286 -0.0374 0.6085 1.0000
4.750 0.7059 0.02397 0.01318 -0.0360 0.5937 1.0000
5.000 0.7275 0.02425 0.01351 -0.0347 0.5789 1.0000
5.250 0.7491 0.02454 0.01387 -0.0333 0.5640 1.0000
5.500 0.7707 0.02484 0.01423 -0.0319 0.5488 1.0000
5.750 0.7922 0.02515 0.01460 -0.0305 0.5331 1.0000
6.000 0.8134 0.02547 0.01501 -0.0291 0.5168 1.0000
6.250 0.8343 0.02578 0.01539 -0.0276 0.4995 1.0000
6.500 0.8550 0.02607 0.01572 -0.0260 0.4808 1.0000
6.750 0.8746 0.02635 0.01602 -0.0242 0.4603 1.0000
7.000 0.8912 0.02674 0.01650 -0.0222 0.4367 1.0000
7.250 0.9087 0.02705 0.01678 -0.0201 0.4128 1.0000
7.500 0.9233 0.02752 0.01730 -0.0179 0.3865 1.0000
7.750 0.9371 0.02804 0.01782 -0.0156 0.3591 1.0000
8.000 0.9496 0.02865 0.01841 -0.0132 0.3307 1.0000
8.250 0.9604 0.02939 0.01909 -0.0107 0.3014 1.0000
8.500 0.9695 0.03030 0.01990 -0.0081 0.2722 1.0000
8.750 0.9765 0.03141 0.02090 -0.0055 0.2436 1.0000
9.000 0.9821 0.03271 0.02211 -0.0029 0.2160 1.0000
9.250 0.9863 0.03416 0.02349 -0.0003 0.1920 1.0000
9.500 0.9883 0.03571 0.02496 0.0025 0.1720 1.0000
9.750 0.9907 0.03739 0.02657 0.0050 0.1559 1.0000
10.000 0.9942 0.03917 0.02831 0.0071 0.1424 1.0000
10.250 0.9988 0.04102 0.03014 0.0089 0.1312 1.0000
10.500 1.0043 0.04290 0.03197 0.0105 0.1228 1.0000
10.750 1.0124 0.04474 0.03390 0.0119 0.1149 1.0000
11.000 1.0210 0.04662 0.03582 0.0132 0.1084 1.0000
11.250 1.0300 0.04853 0.03785 0.0143 0.1020 1.0000
11.500 1.0421 0.05033 0.03960 0.0155 0.0972 1.0000
11.750 1.0525 0.05247 0.04204 0.0165 0.0924 1.0000
12.000 1.0613 0.05454 0.04420 0.0174 0.0881 1.0000
12.250 1.0730 0.05658 0.04626 0.0183 0.0843 1.0000
12.500 1.0757 0.05950 0.04953 0.0191 0.0813 1.0000
12.750 1.0772 0.06245 0.05273 0.0196 0.0785 1.0000
13.000 1.0795 0.06522 0.05562 0.0201 0.0758 1.0000
13.250 1.0896 0.06750 0.05784 0.0207 0.0730 1.0000
13.500 1.0765 0.07199 0.06270 0.0203 0.0718 1.0000
13.750 1.0612 0.07702 0.06807 0.0193 0.0708 1.0000
14.000 1.0422 0.08282 0.07415 0.0175 0.0702 1.0000
14.250 1.0185 0.08968 0.08128 0.0147 0.0698 1.0000
14.500 0.9883 0.09819 0.09003 0.0104 0.0699 1.0000
14.750 0.9503 0.10916 0.10121 0.0042 0.0705 1.0000
15.000 0.9049 0.12358 0.11574 -0.0044 0.0713 1.0000
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Polar data table (+)
Polar graphs
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