NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M21 AIRFOIL (m21-il) Reynolds number: 100,000 Max Cl/Cd: 48.52 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m21-il-100000-n5.txt Download as CSV file: xf-m21-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.2654 0.11119 0.10624 -0.0297 0.7941 0.0587
-9.000 -0.2659 0.10867 0.10369 -0.0331 0.7849 0.0589
-8.750 -0.2666 0.10591 0.10091 -0.0374 0.7762 0.0591
-8.500 -0.2505 0.10084 0.09580 -0.0344 0.7686 0.0598
-8.250 -0.2371 0.09733 0.09223 -0.0328 0.7611 0.0611
-8.000 -0.2293 0.09436 0.08924 -0.0337 0.7532 0.0624
-7.750 -0.2237 0.09143 0.08626 -0.0353 0.7464 0.0634
-7.500 -0.2194 0.08858 0.08339 -0.0373 0.7387 0.0645
-7.250 -0.2139 0.08575 0.08049 -0.0394 0.7326 0.0659
-7.000 -0.2056 0.08290 0.07758 -0.0431 0.7257 0.0677
-6.750 -0.1948 0.08058 0.07508 -0.0494 0.7193 0.0692
-6.500 -0.1807 0.07824 0.07247 -0.0539 0.7140 0.0697
-6.250 -0.1690 0.07372 0.06799 -0.0537 0.7073 0.0702
-6.000 -0.1567 0.07023 0.06449 -0.0528 0.7015 0.0711
-5.750 -0.1423 0.06729 0.06146 -0.0530 0.6961 0.0719
-5.500 -0.1261 0.06451 0.05861 -0.0539 0.6897 0.0729
-5.250 -0.1090 0.06182 0.05579 -0.0547 0.6844 0.0739
-5.000 -0.0905 0.05918 0.05300 -0.0558 0.6792 0.0749
-4.750 -0.0706 0.05661 0.05031 -0.0569 0.6731 0.0763
-4.500 -0.0389 0.05569 0.04876 -0.0597 0.6680 0.0819
-4.250 -0.0178 0.05315 0.04598 -0.0600 0.6633 0.0820
-3.750 0.0222 0.04322 0.03553 -0.0599 0.6532 0.0590
-3.500 0.0437 0.04102 0.03312 -0.0596 0.6490 0.0587
-3.250 0.0675 0.03864 0.03046 -0.0595 0.6432 0.0594
-3.000 0.0890 0.03766 0.02952 -0.0594 0.6377 0.0611
-2.750 0.1128 0.03623 0.02788 -0.0591 0.6333 0.0625
-2.500 0.1380 0.03430 0.02568 -0.0588 0.6281 0.0625
-2.250 0.1635 0.03245 0.02353 -0.0585 0.6228 0.0624
-2.000 0.1894 0.03082 0.02159 -0.0580 0.6184 0.0627
-1.750 0.2157 0.02940 0.01986 -0.0576 0.6140 0.0635
-1.500 0.2424 0.02830 0.01852 -0.0574 0.6083 0.0660
-1.250 0.2700 0.02709 0.01689 -0.0570 0.6036 0.0687
-1.000 0.2977 0.02592 0.01534 -0.0566 0.5998 0.0699
-0.750 0.3244 0.02513 0.01450 -0.0566 0.5942 0.0715
-0.500 0.3514 0.02456 0.01385 -0.0566 0.5890 0.0737
-0.250 0.3790 0.02405 0.01315 -0.0564 0.5849 0.0781
0.000 0.4071 0.02354 0.01238 -0.0562 0.5801 0.0817
0.250 0.4344 0.02295 0.01179 -0.0563 0.5747 0.0843
0.500 0.4617 0.02260 0.01140 -0.0562 0.5702 0.0892
0.750 0.4898 0.02225 0.01086 -0.0561 0.5665 0.0941
1.000 0.5169 0.02193 0.01060 -0.0562 0.5607 0.0976
1.250 0.5438 0.02169 0.01037 -0.0562 0.5559 0.1035
1.500 0.5719 0.02142 0.01000 -0.0560 0.5520 0.1084
1.750 0.6000 0.02122 0.00986 -0.0562 0.5471 0.1128
2.000 0.6271 0.02116 0.00983 -0.0563 0.5419 0.1199
2.250 0.6542 0.02105 0.00964 -0.0561 0.5377 0.1255
2.500 0.6807 0.02094 0.00948 -0.0559 0.5340 0.1315
2.750 0.7052 0.02109 0.00969 -0.0556 0.5282 0.1407
3.000 0.7306 0.02112 0.00974 -0.0553 0.5237 0.1531
3.500 0.8584 0.01988 0.01027 -0.0707 0.5128 1.0000
3.750 0.8819 0.02013 0.01044 -0.0701 0.5082 1.0000
4.000 0.9062 0.02030 0.01051 -0.0694 0.5044 1.0000
4.250 0.9294 0.02060 0.01077 -0.0688 0.5001 1.0000
4.500 0.9514 0.02097 0.01117 -0.0681 0.4947 1.0000
4.750 0.9748 0.02120 0.01136 -0.0674 0.4905 1.0000
5.000 0.9993 0.02137 0.01145 -0.0668 0.4871 1.0000
5.250 1.0198 0.02187 0.01203 -0.0659 0.4819 1.0000
5.500 1.0415 0.02223 0.01243 -0.0652 0.4770 1.0000
5.750 1.0652 0.02244 0.01261 -0.0645 0.4732 1.0000
6.000 1.0878 0.02276 0.01292 -0.0638 0.4692 1.0000
6.250 1.1068 0.02332 0.01361 -0.0627 0.4640 1.0000
6.500 1.1286 0.02366 0.01398 -0.0619 0.4597 1.0000
6.750 1.1527 0.02386 0.01415 -0.0614 0.4562 1.0000
7.000 1.1710 0.02444 0.01484 -0.0602 0.4513 1.0000
7.250 1.1896 0.02498 0.01549 -0.0591 0.4464 1.0000
7.500 1.2116 0.02531 0.01585 -0.0583 0.4426 1.0000
7.750 1.2367 0.02549 0.01601 -0.0579 0.4395 1.0000
8.000 1.2484 0.02639 0.01713 -0.0560 0.4336 1.0000
8.250 1.2668 0.02691 0.01774 -0.0549 0.4291 1.0000
8.500 1.2899 0.02719 0.01805 -0.0542 0.4257 1.0000
8.750 1.3072 0.02779 0.01876 -0.0530 0.4214 1.0000
9.000 1.3182 0.02869 0.01983 -0.0511 0.4161 1.0000
9.250 1.3375 0.02913 0.02037 -0.0500 0.4120 1.0000
9.500 1.3630 0.02931 0.02058 -0.0497 0.4088 1.0000
9.750 1.3658 0.03059 0.02208 -0.0469 0.4031 1.0000
10.000 1.3784 0.03121 0.02282 -0.0452 0.3975 1.0000
10.250 1.3996 0.03121 0.02286 -0.0441 0.3912 1.0000
10.500 1.3940 0.03249 0.02431 -0.0403 0.3835 1.0000
10.750 1.4070 0.03257 0.02440 -0.0383 0.3753 1.0000
11.000 1.3967 0.03424 0.02622 -0.0346 0.3674 1.0000
11.250 1.4028 0.03504 0.02707 -0.0326 0.3590 1.0000
11.500 1.3962 0.03705 0.02920 -0.0304 0.3508 1.0000
11.750 1.3940 0.03888 0.03113 -0.0287 0.3422 1.0000
12.000 1.3959 0.04042 0.03274 -0.0273 0.3324 1.0000
12.250 1.3786 0.04421 0.03670 -0.0260 0.3240 1.0000
12.500 1.3775 0.04634 0.03887 -0.0250 0.3136 1.0000
12.750 1.3633 0.05017 0.04284 -0.0243 0.3043 1.0000
13.000 1.3487 0.05424 0.04702 -0.0238 0.2945 1.0000
13.250 1.3406 0.05756 0.05039 -0.0233 0.2830 1.0000
13.500 1.3303 0.06125 0.05413 -0.0230 0.2703 1.0000
13.750 1.3126 0.06625 0.05923 -0.0231 0.2585 1.0000
14.000 1.3003 0.07056 0.06356 -0.0233 0.2424 1.0000
14.250 1.2848 0.07554 0.06856 -0.0237 0.2241 1.0000
14.500 1.2738 0.07978 0.07266 -0.0239 0.1959 1.0000
14.750 1.2594 0.08450 0.07713 -0.0242 0.1632 1.0000
15.000 1.2404 0.09013 0.08255 -0.0250 0.1334 1.0000
15.250 1.2194 0.09628 0.08851 -0.0260 0.1062 1.0000
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Polar data table (+)
Polar graphs
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