NACA M9 AIRFOIL (m9-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M9 AIRFOIL (m9-il) Reynolds number: 200,000 Max Cl/Cd: 63.08 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m9-il-200000.txt Download as CSV file: xf-m9-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M9 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2305 0.09596 0.09248 -0.0098 0.7893 0.0499
-7.750 -0.2412 0.09526 0.09178 -0.0132 0.7760 0.0508
-7.500 -0.2427 0.09344 0.08993 -0.0175 0.7647 0.0510
-7.250 -0.2293 0.08850 0.08492 -0.0148 0.7546 0.0517
-7.000 -0.2138 0.08534 0.08173 -0.0133 0.7427 0.0526
-6.750 -0.2024 0.08292 0.07926 -0.0134 0.7317 0.0540
-6.500 -0.1923 0.08058 0.07685 -0.0147 0.7213 0.0564
-6.250 -0.1799 0.07853 0.07474 -0.0263 0.7111 0.0600
-6.000 -0.1693 0.07406 0.07021 -0.0270 0.7026 0.0608
-5.750 -0.1556 0.07126 0.06741 -0.0248 0.6914 0.0618
-5.500 -0.1397 0.06884 0.06494 -0.0247 0.6813 0.0634
-5.250 -0.1221 0.06629 0.06230 -0.0264 0.6716 0.0658
-5.000 -0.0865 0.06135 0.05717 -0.0399 0.6626 0.0712
-4.750 -0.0767 0.05896 0.05475 -0.0362 0.6535 0.0726
-4.500 -0.0587 0.05711 0.05289 -0.0353 0.6421 0.0753
-4.000 -0.0003 0.04958 0.04504 -0.0451 0.6248 0.0848
-3.750 0.0184 0.04802 0.04342 -0.0438 0.6164 0.0872
-3.500 0.0625 0.03487 0.02963 -0.0553 0.6112 0.0721
-3.250 0.0926 0.02398 0.01746 -0.0600 0.6060 0.0701
-3.000 0.1181 0.02140 0.01438 -0.0600 0.5982 0.0727
-2.750 0.1448 0.02091 0.01387 -0.0596 0.5893 0.0764
-2.500 0.1723 0.01970 0.01227 -0.0592 0.5815 0.0797
-2.250 0.2003 0.01889 0.01102 -0.0588 0.5734 0.0835
-2.000 0.2270 0.01790 0.00992 -0.0584 0.5668 0.0875
-1.750 0.2549 0.01737 0.00926 -0.0580 0.5592 0.0920
-1.500 0.2831 0.01715 0.00874 -0.0575 0.5530 0.0969
-1.250 0.3104 0.01626 0.00786 -0.0572 0.5454 0.1030
-1.000 0.3383 0.01593 0.00739 -0.0568 0.5382 0.1100
-0.750 0.3657 0.01542 0.00687 -0.0564 0.5311 0.1182
-0.500 0.3936 0.01516 0.00652 -0.0559 0.5239 0.1273
-0.250 0.4208 0.01490 0.00624 -0.0554 0.5183 0.1404
0.000 0.4486 0.01473 0.00620 -0.0551 0.5112 0.1573
0.250 0.4762 0.01539 0.00707 -0.0544 0.5052 0.1767
0.500 0.5039 0.01617 0.00784 -0.0538 0.4999 0.1984
0.750 0.5319 0.01681 0.00841 -0.0532 0.4928 0.2259
1.000 0.5590 0.01734 0.00878 -0.0526 0.4867 0.2467
1.250 0.5858 0.01755 0.00904 -0.0521 0.4807 0.2620
1.500 0.6126 0.01764 0.00912 -0.0517 0.4747 0.2787
1.750 0.6394 0.01769 0.00908 -0.0513 0.4699 0.2938
2.000 0.6662 0.01766 0.00905 -0.0510 0.4646 0.3064
2.250 0.6931 0.01755 0.00896 -0.0507 0.4585 0.3173
2.500 0.7206 0.01774 0.00898 -0.0503 0.4531 0.3256
2.750 0.7472 0.01749 0.00880 -0.0501 0.4472 0.3331
3.000 0.7744 0.01756 0.00883 -0.0497 0.4407 0.3392
3.250 0.8014 0.01744 0.00865 -0.0494 0.4355 0.3451
3.500 0.8283 0.01744 0.00871 -0.0492 0.4290 0.3505
3.750 0.8555 0.01744 0.00869 -0.0489 0.4224 0.3514
4.000 0.8826 0.01756 0.00872 -0.0485 0.4164 0.3544
4.250 0.9092 0.01748 0.00872 -0.0483 0.4084 0.3579
4.500 0.9360 0.01746 0.00862 -0.0479 0.4015 0.3598
4.750 0.9626 0.01747 0.00872 -0.0476 0.3921 0.3587
5.000 0.9891 0.01756 0.00872 -0.0472 0.3839 0.3571
5.250 1.0154 0.01762 0.00885 -0.0468 0.3732 0.3558
5.500 1.0413 0.01765 0.00890 -0.0464 0.3625 0.3551
5.750 1.0669 0.01772 0.00892 -0.0460 0.3514 0.3546
6.000 1.0921 0.01782 0.00904 -0.0456 0.3382 0.3544
6.250 1.1171 0.01800 0.00924 -0.0451 0.3249 0.3541
6.500 1.1417 0.01824 0.00949 -0.0447 0.3129 0.3533
6.750 1.1659 0.01856 0.00974 -0.0442 0.3029 0.3525
7.000 1.1901 0.01887 0.01007 -0.0437 0.2932 0.3518
7.250 1.2139 0.01925 0.01045 -0.0432 0.2852 0.3512
7.500 1.2376 0.01962 0.01083 -0.0427 0.2780 0.3507
7.750 1.2609 0.02006 0.01124 -0.0421 0.2716 0.3502
8.000 1.2842 0.02044 0.01168 -0.0416 0.2650 0.3499
8.500 1.3302 0.02137 0.01266 -0.0406 0.2544 0.3495
8.750 1.3530 0.02184 0.01318 -0.0400 0.2502 0.3494
9.000 1.3756 0.02238 0.01372 -0.0394 0.2466 0.3495
9.250 1.3984 0.02302 0.01434 -0.0389 0.2434 0.3496
9.500 1.4201 0.02356 0.01502 -0.0382 0.2405 0.3499
9.750 1.4417 0.02413 0.01569 -0.0375 0.2375 0.3502
10.000 1.4630 0.02470 0.01632 -0.0369 0.2342 0.3508
10.250 1.4848 0.02535 0.01694 -0.0363 0.2308 0.3519
10.500 1.5059 0.02610 0.01775 -0.0357 0.2276 0.3533
10.750 1.5238 0.02673 0.01856 -0.0347 0.2246 0.3544
11.000 1.5414 0.02735 0.01927 -0.0338 0.2206 0.3554
11.250 1.5597 0.02798 0.01990 -0.0329 0.2163 0.3565
11.500 1.5782 0.02882 0.02076 -0.0321 0.2123 0.3578
11.750 1.5898 0.02954 0.02170 -0.0306 0.2091 0.3590
12.000 1.6015 0.03026 0.02254 -0.0291 0.2053 0.3605
12.250 1.6124 0.03095 0.02326 -0.0274 0.2015 0.3623
12.500 1.6253 0.03185 0.02415 -0.0261 0.1972 0.3647
12.750 1.6225 0.03296 0.02548 -0.0236 0.1941 0.3661
13.000 1.6254 0.03416 0.02683 -0.0221 0.1909 0.3683
13.500 1.6415 0.03644 0.02919 -0.0201 0.1837 0.3760
13.750 1.6399 0.03828 0.03124 -0.0193 0.1806 0.3818
14.000 1.6373 0.04037 0.03354 -0.0188 0.1772 0.3904
14.250 1.6542 0.04224 0.03631 -0.0229 0.1721 1.0000
14.500 1.6584 0.04400 0.03802 -0.0225 0.1681 1.0000
14.750 1.6510 0.04709 0.04138 -0.0228 0.1648 1.0000
15.000 1.6456 0.05013 0.04461 -0.0233 0.1607 1.0000
15.250 1.6411 0.05315 0.04771 -0.0239 0.1566 1.0000
15.500 1.6371 0.05620 0.05083 -0.0244 0.1529 1.0000
15.750 1.6263 0.06040 0.05527 -0.0256 0.1479 1.0000
16.000 1.6185 0.06436 0.05933 -0.0268 0.1432 1.0000
16.250 1.6100 0.06857 0.06364 -0.0282 0.1386 1.0000
16.500 1.5993 0.07333 0.06858 -0.0298 0.1332 1.0000
16.750 1.5860 0.07850 0.07380 -0.0318 0.1275 1.0000
17.000 1.5711 0.08415 0.07962 -0.0339 0.1203 1.0000
17.250 1.5533 0.09029 0.08584 -0.0363 0.1138 1.0000
17.500 1.5348 0.09673 0.09238 -0.0388 0.1067 1.0000
17.750 1.5131 0.10380 0.09952 -0.0418 0.1001 1.0000
18.000 1.4890 0.11147 0.10725 -0.0451 0.0942 1.0000
18.250 1.4655 0.11927 0.11513 -0.0487 0.0899 1.0000
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