NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 200,000 Max Cl/Cd: 65.7 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m11-il-200000.txt Download as CSV file: xf-m11-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M11 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5194 0.09565 0.09204 -0.0226 1.0000 0.0477
-9.000 -0.5268 0.09130 0.08773 -0.0273 1.0000 0.0479
-8.750 -0.5360 0.08748 0.08389 -0.0293 1.0000 0.0480
-8.500 -0.5277 0.08181 0.07831 -0.0260 1.0000 0.0495
-8.250 -0.5219 0.07922 0.07572 -0.0246 1.0000 0.0505
-8.000 -0.5206 0.07633 0.07284 -0.0244 1.0000 0.0519
-7.750 -0.5204 0.07285 0.06936 -0.0255 1.0000 0.0537
-7.500 -0.5204 0.06903 0.06550 -0.0269 1.0000 0.0555
-7.250 -0.5275 0.06477 0.06085 -0.0307 1.0000 0.0599
-7.000 -0.5328 0.05932 0.05520 -0.0300 1.0000 0.0609
-6.750 -0.5201 0.05606 0.05216 -0.0287 1.0000 0.0627
-6.500 -0.5095 0.05416 0.05029 -0.0271 1.0000 0.0654
-6.250 -0.5052 0.05334 0.04882 -0.0252 1.0000 0.0734
-6.000 -0.5026 0.04725 0.04280 -0.0240 1.0000 0.0751
-5.750 -0.4911 0.04482 0.04046 -0.0224 1.0000 0.0769
-5.500 -0.4808 0.04290 0.03851 -0.0203 1.0000 0.0803
-5.250 -0.4784 0.04055 0.03565 -0.0168 1.0000 0.0889
-5.000 -0.4676 0.03822 0.03344 -0.0150 1.0000 0.0912
-4.750 -0.4577 0.03689 0.03206 -0.0124 1.0000 0.0974
-4.500 -0.4526 0.03475 0.02967 -0.0093 1.0000 0.1049
-4.250 -0.4427 0.03321 0.02812 -0.0069 1.0000 0.1087
-4.000 -0.4155 0.03097 0.02562 -0.0079 0.9960 0.1205
-3.750 -0.3756 0.02393 0.01722 -0.0069 0.9909 0.0705
-3.500 -0.3373 0.02081 0.01333 -0.0077 0.9858 0.0613
-3.250 -0.2972 0.01986 0.01215 -0.0099 0.9800 0.0604
-3.000 -0.2589 0.01818 0.01027 -0.0119 0.9741 0.0607
-2.750 -0.2177 0.01649 0.00851 -0.0147 0.9702 0.0625
-2.500 -0.1810 0.01549 0.00746 -0.0164 0.9623 0.0630
-2.250 -0.1391 0.01457 0.00657 -0.0192 0.9574 0.0645
-2.000 -0.1012 0.01386 0.00589 -0.0211 0.9498 0.0668
-1.750 -0.0606 0.01322 0.00524 -0.0236 0.9434 0.0699
-1.500 -0.0253 0.01269 0.00474 -0.0250 0.9339 0.0754
-1.250 0.0129 0.01214 0.00422 -0.0268 0.9264 0.0855
-1.000 0.0401 0.01130 0.00380 -0.0265 0.9134 0.1734
-0.750 0.2016 0.00924 0.00431 -0.0532 0.9400 1.0000
-0.500 0.2391 0.00909 0.00404 -0.0551 0.9206 1.0000
-0.250 0.2709 0.00898 0.00382 -0.0557 0.9000 1.0000
0.000 0.2966 0.00893 0.00367 -0.0550 0.8788 1.0000
0.250 0.3193 0.00893 0.00356 -0.0537 0.8577 1.0000
0.500 0.3417 0.00895 0.00346 -0.0523 0.8385 1.0000
0.750 0.3635 0.00899 0.00341 -0.0509 0.8192 1.0000
1.000 0.3855 0.00904 0.00339 -0.0495 0.8006 1.0000
1.250 0.4078 0.00911 0.00337 -0.0482 0.7828 1.0000
1.500 0.4304 0.00919 0.00336 -0.0469 0.7658 1.0000
1.750 0.4532 0.00928 0.00338 -0.0456 0.7490 1.0000
2.000 0.4760 0.00937 0.00343 -0.0445 0.7318 1.0000
2.250 0.4989 0.00947 0.00348 -0.0433 0.7143 1.0000
2.500 0.5219 0.00958 0.00355 -0.0421 0.6969 1.0000
2.750 0.5446 0.00969 0.00362 -0.0409 0.6780 1.0000
3.000 0.5672 0.00980 0.00368 -0.0396 0.6573 1.0000
3.250 0.5897 0.00991 0.00375 -0.0383 0.6351 1.0000
3.500 0.6118 0.01002 0.00380 -0.0369 0.6092 1.0000
3.750 0.6338 0.01016 0.00388 -0.0354 0.5816 1.0000
4.000 0.6559 0.01033 0.00399 -0.0341 0.5549 1.0000
4.250 0.6779 0.01053 0.00412 -0.0328 0.5272 1.0000
4.500 0.6999 0.01075 0.00429 -0.0315 0.4979 1.0000
4.750 0.7216 0.01099 0.00451 -0.0302 0.4651 1.0000
5.000 0.7424 0.01130 0.00471 -0.0287 0.4247 1.0000
5.250 0.7622 0.01171 0.00497 -0.0272 0.3722 1.0000
5.500 0.7792 0.01240 0.00533 -0.0253 0.2957 1.0000
5.750 0.7903 0.01385 0.00603 -0.0228 0.1634 1.0000
6.000 0.8036 0.01521 0.00700 -0.0206 0.1044 1.0000
6.250 0.8195 0.01627 0.00792 -0.0186 0.0838 1.0000
6.500 0.8361 0.01726 0.00887 -0.0167 0.0710 1.0000
6.750 0.8533 0.01819 0.00980 -0.0148 0.0621 1.0000
7.000 0.8700 0.01925 0.01092 -0.0129 0.0552 1.0000
7.250 0.8860 0.02044 0.01208 -0.0110 0.0494 1.0000
7.500 0.9036 0.02183 0.01354 -0.0092 0.0461 1.0000
7.750 0.9229 0.02289 0.01469 -0.0077 0.0425 1.0000
8.000 0.9413 0.02410 0.01589 -0.0063 0.0394 1.0000
8.250 0.9603 0.02644 0.01835 -0.0050 0.0371 1.0000
8.500 0.9799 0.02798 0.02015 -0.0035 0.0357 1.0000
8.750 0.9982 0.02990 0.02232 -0.0019 0.0344 1.0000
9.000 1.0149 0.03165 0.02429 -0.0002 0.0328 1.0000
9.250 1.0305 0.03328 0.02607 0.0014 0.0314 1.0000
9.500 1.0444 0.03528 0.02818 0.0029 0.0301 1.0000
9.750 1.0539 0.03865 0.03183 0.0049 0.0295 1.0000
10.000 1.0592 0.04208 0.03560 0.0073 0.0294 1.0000
10.250 1.0608 0.04547 0.03936 0.0101 0.0294 1.0000
10.500 1.0604 0.04829 0.04253 0.0131 0.0298 1.0000
10.750 1.0140 0.05499 0.05012 0.0201 0.0330 1.0000
11.000 0.9900 0.05883 0.05420 0.0237 0.0340 1.0000
11.250 0.9672 0.06291 0.05846 0.0252 0.0347 1.0000
11.500 0.9438 0.06757 0.06330 0.0250 0.0352 1.0000
11.750 0.9233 0.07272 0.06856 0.0234 0.0360 1.0000
12.000 0.9019 0.07865 0.07460 0.0205 0.0365 1.0000
12.250 0.7783 0.07805 0.07431 0.0207 0.0350 1.0000
12.500 0.7456 0.08763 0.08398 0.0155 0.0356 1.0000
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Polar data table (+)
Polar graphs
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