NACA M12 AIRFOIL (m12-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M12 AIRFOIL (m12-il) Reynolds number: 200,000 Max Cl/Cd: 65.88 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m12-il-200000.txt Download as CSV file: xf-m12-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.4327 0.08113 0.07765 -0.0355 1.0000 0.0838
-9.250 -0.4222 0.07864 0.07517 -0.0342 1.0000 0.0851
-9.000 -0.6179 0.07666 0.07267 -0.0411 1.0000 0.0820
-8.750 -0.6003 0.07037 0.06658 -0.0411 1.0000 0.0836
-8.500 -0.5840 0.06834 0.06462 -0.0398 1.0000 0.0848
-8.250 -0.5788 0.06572 0.06200 -0.0389 1.0000 0.0862
-8.000 -0.5790 0.06272 0.05895 -0.0380 1.0000 0.0881
-7.750 -0.5851 0.05940 0.05551 -0.0365 1.0000 0.0912
-7.500 -0.6206 0.05504 0.05060 -0.0321 1.0000 0.0964
-7.250 -0.6119 0.05221 0.04787 -0.0302 1.0000 0.0976
-7.000 -0.6061 0.05027 0.04596 -0.0275 1.0000 0.0988
-6.750 -0.6029 0.04866 0.04433 -0.0245 1.0000 0.1009
-6.500 -0.6423 0.03321 0.02662 -0.0142 1.0000 0.0594
-6.250 -0.6261 0.03006 0.02326 -0.0132 0.9990 0.0588
-6.000 -0.5907 0.02747 0.02028 -0.0152 0.9950 0.0588
-5.750 -0.5538 0.02599 0.01841 -0.0173 0.9904 0.0596
-5.500 -0.5174 0.02353 0.01568 -0.0194 0.9862 0.0601
-5.250 -0.4780 0.02170 0.01370 -0.0220 0.9830 0.0609
-5.000 -0.4418 0.02046 0.01240 -0.0239 0.9771 0.0621
-4.750 -0.4017 0.01951 0.01141 -0.0265 0.9727 0.0641
-4.500 -0.3608 0.01866 0.01050 -0.0292 0.9686 0.0668
-4.250 -0.3239 0.01784 0.00962 -0.0310 0.9619 0.0688
-4.000 -0.2827 0.01678 0.00855 -0.0337 0.9581 0.0715
-3.750 -0.2458 0.01598 0.00783 -0.0357 0.9522 0.0758
-3.500 -0.2097 0.01543 0.00725 -0.0372 0.9449 0.0819
-3.250 -0.1746 0.01464 0.00656 -0.0387 0.9388 0.0937
-3.000 -0.1465 0.01381 0.00595 -0.0387 0.9288 0.1263
-2.750 -0.1197 0.01295 0.00548 -0.0386 0.9194 0.2023
-2.500 -0.0965 0.01191 0.00513 -0.0379 0.9099 0.3528
-2.250 -0.0831 0.01072 0.00497 -0.0350 0.8978 0.5874
-2.000 -0.0625 0.01016 0.00513 -0.0321 0.8883 0.7761
-1.750 -0.0318 0.01027 0.00539 -0.0310 0.8786 0.8650
-1.500 0.0014 0.01062 0.00570 -0.0305 0.8685 0.9126
-1.250 0.0396 0.01096 0.00592 -0.0312 0.8597 0.9404
-1.000 0.0879 0.01122 0.00607 -0.0346 0.8488 0.9555
-0.750 0.1412 0.01137 0.00608 -0.0392 0.8389 0.9672
-0.500 0.1951 0.01149 0.00608 -0.0441 0.8280 0.9820
-0.250 0.2545 0.01144 0.00592 -0.0505 0.8154 0.9944
0.000 0.2967 0.01130 0.00569 -0.0536 0.8020 1.0000
0.250 0.3200 0.01122 0.00551 -0.0528 0.7881 1.0000
0.500 0.3435 0.01116 0.00534 -0.0520 0.7746 1.0000
0.750 0.3673 0.01111 0.00519 -0.0513 0.7610 1.0000
1.000 0.3914 0.01108 0.00509 -0.0507 0.7463 1.0000
1.250 0.4155 0.01106 0.00500 -0.0500 0.7315 1.0000
1.500 0.4397 0.01105 0.00492 -0.0493 0.7164 1.0000
1.750 0.4640 0.01106 0.00485 -0.0487 0.7012 1.0000
2.000 0.4884 0.01108 0.00481 -0.0480 0.6864 1.0000
2.250 0.5130 0.01112 0.00478 -0.0474 0.6723 1.0000
2.500 0.5376 0.01117 0.00477 -0.0468 0.6587 1.0000
2.750 0.5620 0.01124 0.00477 -0.0461 0.6454 1.0000
3.000 0.5862 0.01131 0.00480 -0.0454 0.6317 1.0000
3.250 0.6103 0.01140 0.00487 -0.0447 0.6177 1.0000
3.500 0.6342 0.01150 0.00495 -0.0440 0.6039 1.0000
3.750 0.6581 0.01161 0.00504 -0.0432 0.5902 1.0000
4.000 0.6818 0.01175 0.00514 -0.0424 0.5767 1.0000
4.250 0.7054 0.01189 0.00525 -0.0416 0.5630 1.0000
4.500 0.7287 0.01205 0.00537 -0.0407 0.5488 1.0000
4.750 0.7515 0.01220 0.00549 -0.0397 0.5325 1.0000
5.000 0.7736 0.01235 0.00563 -0.0385 0.5140 1.0000
5.250 0.7955 0.01251 0.00577 -0.0374 0.4946 1.0000
5.500 0.8168 0.01269 0.00590 -0.0361 0.4737 1.0000
5.750 0.8379 0.01290 0.00605 -0.0348 0.4520 1.0000
6.000 0.8589 0.01312 0.00626 -0.0335 0.4299 1.0000
6.250 0.8795 0.01338 0.00648 -0.0321 0.4085 1.0000
6.500 0.8999 0.01366 0.00673 -0.0308 0.3856 1.0000
6.750 0.9196 0.01398 0.00701 -0.0293 0.3611 1.0000
7.000 0.9390 0.01434 0.00732 -0.0278 0.3341 1.0000
7.250 0.9571 0.01478 0.00767 -0.0262 0.3047 1.0000
7.500 0.9742 0.01531 0.00808 -0.0244 0.2730 1.0000
7.750 0.9898 0.01597 0.00860 -0.0225 0.2379 1.0000
8.000 1.0034 0.01680 0.00922 -0.0203 0.1945 1.0000
8.250 1.0132 0.01797 0.01006 -0.0177 0.1423 1.0000
8.500 1.0222 0.01925 0.01107 -0.0150 0.1111 1.0000
8.750 1.0326 0.02038 0.01210 -0.0124 0.0961 1.0000
9.000 1.0444 0.02140 0.01312 -0.0101 0.0868 1.0000
9.250 1.0527 0.02264 0.01427 -0.0074 0.0804 1.0000
9.500 1.0669 0.02347 0.01519 -0.0055 0.0752 1.0000
9.750 1.0762 0.02453 0.01622 -0.0030 0.0711 1.0000
10.000 1.0854 0.02573 0.01745 -0.0005 0.0677 1.0000
10.250 1.0977 0.02669 0.01850 0.0015 0.0645 1.0000
10.500 1.1090 0.02778 0.01959 0.0034 0.0615 1.0000
10.750 1.1216 0.02940 0.02115 0.0050 0.0586 1.0000
11.000 1.1351 0.03046 0.02236 0.0066 0.0564 1.0000
11.250 1.1483 0.03163 0.02363 0.0081 0.0540 1.0000
11.500 1.1615 0.03289 0.02488 0.0094 0.0518 1.0000
11.750 1.1800 0.03480 0.02678 0.0101 0.0495 1.0000
12.000 1.1912 0.03616 0.02834 0.0117 0.0480 1.0000
12.250 1.2028 0.03764 0.02997 0.0130 0.0463 1.0000
12.500 1.2131 0.03914 0.03157 0.0143 0.0447 1.0000
12.750 1.2269 0.04071 0.03313 0.0152 0.0431 1.0000
13.000 1.2411 0.04327 0.03580 0.0160 0.0416 1.0000
13.250 1.2424 0.04530 0.03810 0.0177 0.0408 1.0000
13.500 1.2434 0.04764 0.04069 0.0192 0.0400 1.0000
13.750 1.2431 0.05014 0.04342 0.0205 0.0392 1.0000
14.000 1.2423 0.05271 0.04619 0.0215 0.0384 1.0000
14.250 1.2430 0.05517 0.04877 0.0223 0.0375 1.0000
14.500 1.2422 0.05776 0.05148 0.0228 0.0368 1.0000
14.750 1.2446 0.06040 0.05417 0.0231 0.0360 1.0000
15.250 1.2248 0.06898 0.06312 0.0232 0.0351 1.0000
15.500 1.2054 0.07376 0.06816 0.0223 0.0350 1.0000
15.750 1.1850 0.07910 0.07375 0.0207 0.0350 1.0000
16.000 1.1628 0.08506 0.07995 0.0184 0.0349 1.0000
16.250 1.1397 0.09160 0.08670 0.0155 0.0350 1.0000
16.500 1.1152 0.09891 0.09421 0.0118 0.0351 1.0000
16.750 1.0903 0.10686 0.10234 0.0075 0.0352 1.0000
17.000 1.0625 0.11591 0.11162 0.0015 0.0354 1.0000
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