NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M21 AIRFOIL (m21-il) Reynolds number: 200,000 Max Cl/Cd: 70.14 at α=10.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m21-il-200000.txt Download as CSV file: xf-m21-il-200000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.2386 0.09048 0.08660 -0.0301 0.7585 0.0524
-7.500 -0.2374 0.08771 0.08381 -0.0343 0.7505 0.0539
-7.250 -0.2324 0.08567 0.08157 -0.0459 0.7445 0.0550
-7.000 -0.2224 0.08121 0.07706 -0.0478 0.7383 0.0553
-6.750 -0.2116 0.07718 0.07305 -0.0460 0.7316 0.0558
-6.500 -0.2000 0.07409 0.06988 -0.0451 0.7262 0.0564
-6.250 -0.1862 0.07125 0.06704 -0.0456 0.7192 0.0573
-6.000 -0.1715 0.06857 0.06427 -0.0466 0.7131 0.0588
-5.750 -0.1551 0.06584 0.06142 -0.0484 0.7081 0.0609
-5.500 -0.1210 0.06460 0.05972 -0.0558 0.7017 0.0643
-5.250 -0.1080 0.05984 0.05489 -0.0561 0.6964 0.0649
-5.000 -0.0942 0.05651 0.05154 -0.0555 0.6917 0.0655
-4.750 -0.0771 0.05387 0.04892 -0.0554 0.6852 0.0663
-4.500 -0.0584 0.05161 0.04657 -0.0555 0.6798 0.0678
-4.250 -0.0376 0.04952 0.04431 -0.0557 0.6754 0.0702
-4.000 0.0099 0.03191 0.02651 -0.0573 0.6600 0.0758
-3.750 0.0249 0.02862 0.02328 -0.0572 0.6552 0.0766
-3.500 0.0427 0.02634 0.02099 -0.0569 0.6499 0.0778
-3.250 0.0623 0.02440 0.01893 -0.0565 0.6453 0.0794
-3.000 0.0839 0.02264 0.01703 -0.0563 0.6410 0.0823
-2.750 0.1121 0.02099 0.01490 -0.0563 0.6354 0.0888
-2.500 0.1313 0.01891 0.01284 -0.0561 0.6309 0.0902
-2.250 0.1527 0.01745 0.01126 -0.0557 0.6270 0.0924
-2.000 0.1832 0.01767 0.01101 -0.0550 0.6211 0.1014
-1.750 0.2028 0.01498 0.00830 -0.0550 0.6167 0.1032
-1.500 0.2254 0.01373 0.00700 -0.0548 0.6126 0.1057
-1.250 0.2505 0.01295 0.00604 -0.0544 0.6087 0.1119
-1.000 0.2747 0.01198 0.00497 -0.0542 0.6032 0.1208
-0.750 0.3002 0.01130 0.00407 -0.0539 0.5983 0.1339
-0.500 0.3251 0.01053 0.00322 -0.0537 0.5946 0.1400
-0.250 0.3724 0.02189 0.01312 -0.0539 0.5964 0.0891
0.000 0.3995 0.02107 0.01229 -0.0540 0.5925 0.0927
0.250 0.4280 0.02021 0.01132 -0.0539 0.5868 0.0929
0.500 0.4564 0.01957 0.01054 -0.0538 0.5819 0.0952
0.750 0.4849 0.01927 0.01002 -0.0536 0.5780 0.1002
1.000 0.5130 0.01865 0.00936 -0.0537 0.5732 0.1036
1.250 0.5406 0.01825 0.00901 -0.0538 0.5679 0.1086
1.500 0.5686 0.01801 0.00867 -0.0537 0.5636 0.1142
1.750 0.5962 0.01764 0.00824 -0.0535 0.5597 0.1184
2.000 0.6225 0.01744 0.00818 -0.0535 0.5542 0.1249
2.250 0.6495 0.01728 0.00801 -0.0533 0.5495 0.1298
2.500 0.6759 0.01700 0.00772 -0.0530 0.5457 0.1362
2.750 0.7018 0.01697 0.00777 -0.0528 0.5408 0.1440
3.000 0.7273 0.01688 0.00775 -0.0524 0.5357 0.1516
3.250 0.7538 0.01683 0.00768 -0.0522 0.5316 0.1645
3.500 0.7804 0.01679 0.00765 -0.0519 0.5281 0.1948
3.750 0.9000 0.01561 0.00819 -0.0716 0.5194 1.0000
4.000 0.9253 0.01574 0.00822 -0.0711 0.5153 1.0000
4.250 0.9508 0.01595 0.00831 -0.0707 0.5117 1.0000
4.500 0.9742 0.01620 0.00865 -0.0702 0.5061 1.0000
4.750 0.9987 0.01636 0.00879 -0.0698 0.5014 1.0000
5.000 1.0242 0.01651 0.00884 -0.0694 0.4976 1.0000
5.250 1.0478 0.01681 0.00919 -0.0689 0.4931 1.0000
5.500 1.0712 0.01708 0.00951 -0.0683 0.4880 1.0000
5.750 1.0960 0.01724 0.00964 -0.0678 0.4838 1.0000
6.000 1.1217 0.01745 0.00975 -0.0675 0.4803 1.0000
6.250 1.1430 0.01781 0.01027 -0.0668 0.4748 1.0000
6.500 1.1666 0.01806 0.01056 -0.0662 0.4703 1.0000
6.750 1.1918 0.01825 0.01070 -0.0658 0.4665 1.0000
7.000 1.2149 0.01859 0.01110 -0.0652 0.4622 1.0000
7.250 1.2363 0.01894 0.01156 -0.0644 0.4569 1.0000
7.500 1.2603 0.01918 0.01182 -0.0640 0.4528 1.0000
7.750 1.2862 0.01941 0.01198 -0.0637 0.4493 1.0000
8.000 1.3056 0.01988 0.01264 -0.0627 0.4441 1.0000
8.250 1.3276 0.02013 0.01296 -0.0619 0.4388 1.0000
8.500 1.3541 0.02012 0.01284 -0.0617 0.4335 1.0000
8.750 1.3710 0.02039 0.01332 -0.0601 0.4258 1.0000
9.000 1.3953 0.02030 0.01316 -0.0595 0.4190 1.0000
9.250 1.4128 0.02059 0.01358 -0.0581 0.4116 1.0000
9.500 1.4343 0.02063 0.01364 -0.0571 0.4045 1.0000
9.750 1.4514 0.02087 0.01399 -0.0556 0.3966 1.0000
10.000 1.4704 0.02097 0.01411 -0.0543 0.3886 1.0000
10.250 1.4847 0.02127 0.01453 -0.0524 0.3799 1.0000
10.500 1.5017 0.02141 0.01463 -0.0508 0.3707 1.0000
10.750 1.5098 0.02179 0.01519 -0.0480 0.3597 1.0000
11.000 1.5166 0.02224 0.01572 -0.0451 0.3475 1.0000
11.250 1.5183 0.02282 0.01634 -0.0414 0.3347 1.0000
11.500 1.5137 0.02373 0.01729 -0.0373 0.3201 1.0000
11.750 1.5065 0.02521 0.01877 -0.0339 0.3006 1.0000
12.000 1.4944 0.02749 0.02099 -0.0310 0.2738 1.0000
12.250 1.4758 0.03078 0.02413 -0.0285 0.2403 1.0000
12.500 1.4518 0.03500 0.02815 -0.0265 0.2046 1.0000
12.750 1.4249 0.03984 0.03280 -0.0250 0.1712 1.0000
13.000 1.3975 0.04498 0.03779 -0.0238 0.1412 1.0000
13.250 1.3691 0.05048 0.04313 -0.0229 0.1110 1.0000
13.500 1.3415 0.05625 0.04874 -0.0224 0.0841 1.0000
13.750 1.3174 0.06189 0.05428 -0.0223 0.0650 1.0000
14.000 1.2991 0.06708 0.05942 -0.0223 0.0541 1.0000
14.250 1.2844 0.07198 0.06434 -0.0225 0.0484 1.0000
14.500 1.2753 0.07631 0.06876 -0.0228 0.0448 1.0000
14.750 1.2654 0.08083 0.07334 -0.0232 0.0423 1.0000
15.000 1.2540 0.08562 0.07820 -0.0237 0.0406 1.0000
15.250 1.2464 0.08997 0.08266 -0.0243 0.0392 1.0000
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Polar data table (+)
Polar graphs
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