NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M21 AIRFOIL (m21-il) Reynolds number: 200,000 Max Cl/Cd: 73.67 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m21-il-200000-n5.txt Download as CSV file: xf-m21-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M21 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.2675 0.09443 0.09014 -0.0284 0.6949 0.0424
-8.000 -0.2680 0.09099 0.08667 -0.0326 0.6889 0.0425
-7.750 -0.2635 0.08740 0.08302 -0.0363 0.6836 0.0426
-7.250 -0.2449 0.07966 0.07511 -0.0434 0.6729 0.0427
-7.000 -0.2342 0.07669 0.07215 -0.0410 0.6674 0.0433
-6.750 -0.2211 0.07433 0.06976 -0.0406 0.6620 0.0441
-6.500 -0.2075 0.07158 0.06695 -0.0420 0.6563 0.0450
-6.250 -0.1932 0.06851 0.06379 -0.0440 0.6513 0.0456
-6.000 -0.1776 0.06334 0.05841 -0.0483 0.6474 0.0419
-5.750 -0.1618 0.06039 0.05541 -0.0491 0.6420 0.0409
-5.500 -0.1430 0.05694 0.05179 -0.0511 0.6373 0.0415
-5.250 -0.1238 0.05370 0.04836 -0.0525 0.6330 0.0415
-5.000 -0.1042 0.05075 0.04529 -0.0533 0.6280 0.0409
-4.750 -0.0835 0.04771 0.04208 -0.0541 0.6230 0.0405
-4.500 -0.0621 0.04472 0.03889 -0.0547 0.6187 0.0403
-4.250 -0.0399 0.04182 0.03574 -0.0551 0.6147 0.0403
-4.000 -0.0164 0.03877 0.03246 -0.0554 0.6098 0.0412
-3.750 0.0070 0.03551 0.02890 -0.0553 0.6052 0.0418
-3.500 0.0302 0.03250 0.02555 -0.0549 0.6012 0.0419
-3.250 0.0536 0.02908 0.02173 -0.0543 0.5972 0.0423
-3.000 0.0773 0.02570 0.01790 -0.0536 0.5927 0.0430
-2.750 0.1032 0.02509 0.01718 -0.0535 0.5879 0.0437
-2.500 0.1294 0.02463 0.01659 -0.0534 0.5837 0.0447
-2.250 0.1560 0.02353 0.01533 -0.0533 0.5789 0.0461
-2.000 0.1828 0.02177 0.01323 -0.0529 0.5743 0.0472
-1.750 0.2101 0.01997 0.01101 -0.0525 0.5701 0.0489
-1.500 0.2380 0.01865 0.00931 -0.0523 0.5662 0.0505
-1.250 0.2657 0.01838 0.00906 -0.0524 0.5611 0.0518
-1.000 0.2934 0.01820 0.00884 -0.0525 0.5564 0.0537
-0.750 0.3213 0.01778 0.00826 -0.0524 0.5523 0.0562
-0.500 0.3497 0.01719 0.00747 -0.0524 0.5479 0.0587
-0.250 0.3775 0.01680 0.00709 -0.0525 0.5431 0.0607
0.000 0.4052 0.01663 0.00689 -0.0525 0.5387 0.0630
0.250 0.4329 0.01649 0.00664 -0.0524 0.5349 0.0667
0.500 0.4608 0.01626 0.00637 -0.0525 0.5301 0.0701
0.750 0.4881 0.01602 0.00617 -0.0525 0.5254 0.0729
1.000 0.5155 0.01592 0.00605 -0.0524 0.5214 0.0770
1.250 0.5428 0.01581 0.00587 -0.0524 0.5175 0.0808
1.500 0.5697 0.01560 0.00571 -0.0523 0.5126 0.0839
1.750 0.5965 0.01549 0.00564 -0.0522 0.5081 0.0881
2.000 0.6233 0.01541 0.00552 -0.0521 0.5043 0.0919
2.250 0.6501 0.01536 0.00546 -0.0520 0.5002 0.0950
2.500 0.6763 0.01524 0.00542 -0.0518 0.4954 0.0996
2.750 0.7028 0.01520 0.00540 -0.0517 0.4910 0.1045
3.000 0.7293 0.01521 0.00535 -0.0515 0.4874 0.1086
3.250 0.7555 0.01520 0.00539 -0.0514 0.4831 0.1140
3.500 0.7819 0.01522 0.00546 -0.0513 0.4784 0.1223
3.750 0.8081 0.01524 0.00550 -0.0511 0.4742 0.1345
4.000 0.8336 0.01519 0.00560 -0.0508 0.4705 0.2077
4.500 0.9720 0.01435 0.00637 -0.0691 0.4589 1.0000
5.000 1.0205 0.01471 0.00669 -0.0681 0.4504 1.0000
5.250 1.0445 0.01490 0.00689 -0.0676 0.4460 1.0000
5.500 1.0683 0.01509 0.00706 -0.0670 0.4422 1.0000
5.750 1.0921 0.01530 0.00724 -0.0665 0.4383 1.0000
6.000 1.1155 0.01552 0.00753 -0.0659 0.4335 1.0000
6.250 1.1387 0.01574 0.00777 -0.0653 0.4290 1.0000
6.500 1.1619 0.01597 0.00798 -0.0646 0.4254 1.0000
6.750 1.1849 0.01622 0.00827 -0.0640 0.4214 1.0000
7.000 1.2074 0.01648 0.00861 -0.0633 0.4167 1.0000
7.250 1.2299 0.01673 0.00889 -0.0626 0.4124 1.0000
7.500 1.2524 0.01700 0.00914 -0.0619 0.4088 1.0000
7.750 1.2742 0.01730 0.00955 -0.0611 0.4045 1.0000
8.000 1.2956 0.01759 0.00993 -0.0603 0.4000 1.0000
8.250 1.3169 0.01788 0.01024 -0.0594 0.3954 1.0000
8.500 1.3371 0.01819 0.01060 -0.0584 0.3894 1.0000
8.750 1.3560 0.01849 0.01097 -0.0572 0.3817 1.0000
9.000 1.3741 0.01881 0.01133 -0.0559 0.3740 1.0000
9.250 1.3911 0.01917 0.01174 -0.0544 0.3655 1.0000
9.500 1.4070 0.01956 0.01218 -0.0528 0.3571 1.0000
9.750 1.4215 0.02000 0.01266 -0.0509 0.3485 1.0000
10.000 1.4346 0.02049 0.01322 -0.0490 0.3392 1.0000
10.250 1.4443 0.02106 0.01382 -0.0465 0.3290 1.0000
10.500 1.4498 0.02174 0.01453 -0.0435 0.3183 1.0000
10.750 1.4527 0.02258 0.01541 -0.0404 0.3065 1.0000
11.000 1.4553 0.02366 0.01651 -0.0378 0.2938 1.0000
11.250 1.4552 0.02510 0.01797 -0.0353 0.2776 1.0000
11.500 1.4515 0.02702 0.01985 -0.0331 0.2582 1.0000
11.750 1.4437 0.02952 0.02227 -0.0311 0.2340 1.0000
12.000 1.4289 0.03286 0.02549 -0.0293 0.2056 1.0000
12.250 1.4099 0.03683 0.02932 -0.0277 0.1774 1.0000
12.500 1.3891 0.04119 0.03356 -0.0264 0.1508 1.0000
12.750 1.3641 0.04618 0.03839 -0.0252 0.1214 1.0000
13.000 1.3362 0.05168 0.04373 -0.0244 0.0917 1.0000
13.250 1.3103 0.05731 0.04922 -0.0240 0.0652 1.0000
13.500 1.2894 0.06263 0.05443 -0.0239 0.0450 1.0000
13.750 1.2772 0.06709 0.05889 -0.0239 0.0356 1.0000
14.000 1.2694 0.07110 0.06294 -0.0240 0.0310 1.0000
14.250 1.2630 0.07505 0.06695 -0.0242 0.0282 1.0000
14.500 1.2588 0.07876 0.07075 -0.0245 0.0263 1.0000
14.750 1.2546 0.08252 0.07461 -0.0249 0.0248 1.0000
15.000 1.2495 0.08646 0.07863 -0.0253 0.0236 1.0000
15.250 1.2432 0.09062 0.08287 -0.0259 0.0226 1.0000
15.500 1.2401 0.09439 0.08676 -0.0264 0.0218 1.0000
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Polar data table (+)
Polar graphs
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