NACA M9 AIRFOIL (m9-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M9 AIRFOIL (m9-il) Reynolds number: 100,000 Max Cl/Cd: 49.48 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m9-il-100000-n5.txt Download as CSV file: xf-m9-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.2570 0.11729 0.11285 -0.0110 0.8474 0.0478
-9.250 -0.2518 0.11470 0.11019 -0.0124 0.8260 0.0478
-9.000 -0.2478 0.11204 0.10748 -0.0140 0.8090 0.0479
-8.750 -0.2278 0.10632 0.10170 -0.0122 0.7950 0.0486
-8.500 -0.2179 0.10323 0.09856 -0.0124 0.7819 0.0491
-8.250 -0.2100 0.10038 0.09565 -0.0132 0.7700 0.0494
-8.000 -0.2041 0.09761 0.09287 -0.0147 0.7585 0.0491
-7.750 -0.1917 0.09512 0.09032 -0.0144 0.7477 0.0518
-7.500 -0.1852 0.09281 0.08799 -0.0156 0.7374 0.0549
-6.750 -0.1754 0.08156 0.07667 -0.0240 0.7114 0.0394
-6.500 -0.1641 0.07843 0.07349 -0.0255 0.7027 0.0391
-6.250 -0.1514 0.07521 0.07024 -0.0277 0.6933 0.0388
-6.000 -0.1377 0.07149 0.06645 -0.0313 0.6855 0.0391
-5.750 -0.1213 0.06937 0.06432 -0.0323 0.6753 0.0414
-5.500 -0.1041 0.06668 0.06154 -0.0343 0.6671 0.0429
-5.250 -0.0842 0.06293 0.05774 -0.0381 0.6575 0.0432
-5.000 -0.0628 0.05902 0.05368 -0.0419 0.6500 0.0437
-4.750 -0.0367 0.05435 0.04888 -0.0474 0.6409 0.0463
-4.500 -0.0072 0.04772 0.04198 -0.0543 0.6346 0.0483
-4.250 0.0158 0.04652 0.04073 -0.0546 0.6246 0.0499
-4.000 0.0420 0.04419 0.03821 -0.0563 0.6170 0.0534
-3.750 0.0802 0.03309 0.02634 -0.0656 0.6105 0.0584
-3.500 0.1037 0.03407 0.02736 -0.0641 0.6014 0.0607
-3.250 0.1353 0.02813 0.02068 -0.0677 0.5939 0.0653
-3.000 0.1645 0.02455 0.01623 -0.0690 0.5866 0.0701
-2.750 0.1911 0.02373 0.01526 -0.0687 0.5790 0.0728
-2.500 0.2186 0.02284 0.01409 -0.0685 0.5714 0.0771
-2.250 0.2467 0.02180 0.01251 -0.0682 0.5651 0.0817
-2.000 0.2736 0.02122 0.01193 -0.0679 0.5571 0.0849
-1.750 0.3008 0.02076 0.01128 -0.0674 0.5507 0.0900
-1.500 0.3289 0.02029 0.01053 -0.0672 0.5437 0.0960
-1.250 0.3559 0.01991 0.01015 -0.0668 0.5373 0.1011
-1.000 0.3835 0.01960 0.00963 -0.0663 0.5324 0.1090
-0.750 0.4110 0.01934 0.00938 -0.0661 0.5262 0.1168
-0.500 0.4390 0.01914 0.00906 -0.0658 0.5204 0.1264
-0.250 0.4663 0.01902 0.00895 -0.0653 0.5154 0.1388
0.000 0.4935 0.01918 0.00924 -0.0649 0.5082 0.1516
0.250 0.5196 0.01972 0.00990 -0.0641 0.5014 0.1654
0.500 0.5458 0.02028 0.01048 -0.0634 0.4954 0.1799
0.750 0.5743 0.02054 0.01050 -0.0630 0.4880 0.2182
1.250 0.6277 0.02111 0.01078 -0.0619 0.4746 0.2512
1.500 0.6543 0.02119 0.01076 -0.0615 0.4677 0.2626
1.750 0.6810 0.02127 0.01067 -0.0610 0.4619 0.2726
2.000 0.7074 0.02129 0.01073 -0.0607 0.4544 0.2819
2.250 0.7339 0.02127 0.01064 -0.0603 0.4484 0.2904
2.500 0.7606 0.02137 0.01069 -0.0600 0.4425 0.2977
2.750 0.7872 0.02139 0.01072 -0.0597 0.4359 0.3035
3.000 0.8138 0.02143 0.01068 -0.0593 0.4303 0.3108
3.250 0.8403 0.02159 0.01084 -0.0591 0.4241 0.3161
3.500 0.8665 0.02160 0.01090 -0.0588 0.4176 0.3199
3.750 0.8932 0.02166 0.01087 -0.0584 0.4121 0.3210
4.000 0.9200 0.02183 0.01112 -0.0583 0.4042 0.3214
4.250 0.9472 0.02192 0.01116 -0.0581 0.3972 0.3219
4.500 0.9741 0.02209 0.01136 -0.0579 0.3894 0.3222
4.750 1.0007 0.02223 0.01150 -0.0577 0.3816 0.3221
5.000 1.0269 0.02241 0.01169 -0.0575 0.3739 0.3217
5.250 1.0527 0.02259 0.01191 -0.0572 0.3654 0.3215
5.500 1.0781 0.02280 0.01215 -0.0568 0.3571 0.3213
5.750 1.1029 0.02301 0.01239 -0.0564 0.3479 0.3211
6.000 1.1272 0.02328 0.01271 -0.0559 0.3380 0.3210
6.250 1.1509 0.02351 0.01293 -0.0553 0.3288 0.3210
6.500 1.1739 0.02386 0.01336 -0.0548 0.3179 0.3210
6.750 1.1965 0.02421 0.01371 -0.0541 0.3086 0.3210
7.000 1.2183 0.02462 0.01414 -0.0534 0.2993 0.3212
7.250 1.2398 0.02508 0.01464 -0.0527 0.2913 0.3217
7.500 1.2607 0.02556 0.01515 -0.0519 0.2839 0.3224
7.750 1.2811 0.02609 0.01570 -0.0511 0.2775 0.3231
8.000 1.3010 0.02668 0.01635 -0.0502 0.2714 0.3239
8.250 1.3204 0.02725 0.01691 -0.0493 0.2666 0.3246
8.500 1.3396 0.02792 0.01765 -0.0484 0.2617 0.3252
8.750 1.3579 0.02861 0.01841 -0.0475 0.2569 0.3256
9.000 1.3756 0.02928 0.01910 -0.0465 0.2525 0.3260
9.250 1.3928 0.03003 0.01987 -0.0455 0.2483 0.3263
9.500 1.4085 0.03086 0.02082 -0.0444 0.2437 0.3269
9.750 1.4235 0.03168 0.02169 -0.0432 0.2397 0.3274
10.000 1.4395 0.03246 0.02246 -0.0421 0.2364 0.3280
10.250 1.4542 0.03336 0.02345 -0.0410 0.2333 0.3287
10.500 1.4649 0.03438 0.02462 -0.0395 0.2305 0.3292
10.750 1.4756 0.03545 0.02581 -0.0381 0.2277 0.3300
11.000 1.4872 0.03654 0.02699 -0.0369 0.2253 0.3308
11.250 1.4996 0.03760 0.02813 -0.0358 0.2227 0.3317
11.500 1.5161 0.03848 0.02904 -0.0350 0.2202 0.3328
11.750 1.5241 0.03990 0.03060 -0.0339 0.2175 0.3339
12.000 1.5270 0.04170 0.03260 -0.0329 0.2150 0.3349
12.250 1.5306 0.04354 0.03462 -0.0322 0.2121 0.3366
12.500 1.5367 0.04525 0.03647 -0.0315 0.2095 0.3382
12.750 1.5455 0.04679 0.03812 -0.0309 0.2071 0.3409
13.000 1.5597 0.04789 0.03928 -0.0303 0.2044 0.3441
13.250 1.5626 0.04998 0.04151 -0.0298 0.2017 0.3465
13.500 1.5498 0.05359 0.04539 -0.0299 0.1990 0.3478
14.000 1.5339 0.06024 0.05237 -0.0305 0.1916 0.3519
14.250 1.5426 0.06154 0.05363 -0.0303 0.1862 0.3572
14.500 1.5170 0.06750 0.05989 -0.0319 0.1837 0.3571
14.750 1.4897 0.07415 0.06678 -0.0341 0.1811 0.3569
15.000 1.4608 0.08150 0.07435 -0.0369 0.1782 0.3564
15.500 1.4524 0.08891 0.08193 -0.0393 0.1699 0.3651
15.750 1.3787 0.10544 0.09873 -0.0466 0.1696 0.3544
16.000 1.1859 0.14987 0.14317 -0.0694 0.1530 0.3351
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