NACA M9 AIRFOIL (m9-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M9 AIRFOIL (m9-il) Reynolds number: 500,000 Max Cl/Cd: 81.16 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m9-il-500000-n5.txt Download as CSV file: xf-m9-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.2619 0.08945 0.08612 -0.0105 0.6357 0.0111
-8.250 -0.2696 0.08407 0.08076 -0.0126 0.6309 0.0115
-7.750 -0.3601 0.03647 0.03250 -0.0566 0.6367 0.0142
-7.500 -0.3419 0.02626 0.02159 -0.0644 0.6313 0.0147
-7.250 -0.3173 0.02273 0.01774 -0.0664 0.6228 0.0153
-7.000 -0.2916 0.02051 0.01526 -0.0674 0.6143 0.0161
-6.750 -0.2652 0.01897 0.01349 -0.0677 0.6038 0.0169
-6.500 -0.2385 0.01773 0.01202 -0.0678 0.5928 0.0179
-6.250 -0.2117 0.01666 0.01075 -0.0678 0.5811 0.0189
-6.000 -0.1845 0.01583 0.00978 -0.0677 0.5690 0.0203
-5.750 -0.1570 0.01520 0.00898 -0.0675 0.5580 0.0219
-5.500 -0.1296 0.01460 0.00824 -0.0673 0.5475 0.0235
-5.250 -0.1018 0.01414 0.00768 -0.0671 0.5371 0.0252
-5.000 -0.0740 0.01378 0.00716 -0.0668 0.5278 0.0270
-4.750 -0.0463 0.01340 0.00669 -0.0665 0.5180 0.0287
-4.500 -0.0183 0.01315 0.00634 -0.0662 0.5092 0.0307
-4.250 0.0096 0.01292 0.00599 -0.0659 0.5004 0.0325
-4.000 0.0375 0.01264 0.00563 -0.0655 0.4931 0.0343
-3.750 0.0656 0.01244 0.00537 -0.0652 0.4858 0.0363
-3.500 0.0937 0.01229 0.00511 -0.0649 0.4794 0.0383
-3.250 0.1219 0.01211 0.00486 -0.0646 0.4734 0.0399
-3.000 0.1498 0.01189 0.00460 -0.0643 0.4672 0.0423
-2.750 0.1780 0.01176 0.00440 -0.0640 0.4611 0.0445
-2.500 0.2061 0.01163 0.00419 -0.0637 0.4542 0.0465
-2.250 0.2342 0.01152 0.00399 -0.0633 0.4472 0.0484
-2.000 0.2623 0.01134 0.00379 -0.0630 0.4407 0.0514
-1.750 0.2904 0.01124 0.00364 -0.0627 0.4338 0.0540
-1.500 0.3185 0.01118 0.00350 -0.0624 0.4286 0.0567
-1.250 0.3468 0.01106 0.00336 -0.0621 0.4245 0.0604
-1.000 0.3750 0.01097 0.00327 -0.0618 0.4197 0.0650
-0.750 0.4030 0.01094 0.00318 -0.0615 0.4146 0.0690
-0.500 0.4311 0.01087 0.00311 -0.0612 0.4102 0.0750
-0.250 0.4593 0.01083 0.00306 -0.0610 0.4057 0.0808
0.000 0.4873 0.01078 0.00301 -0.0607 0.4008 0.0877
0.250 0.5152 0.01080 0.00298 -0.0604 0.3962 0.0941
0.500 0.5432 0.01077 0.00295 -0.0601 0.3919 0.1004
0.750 0.5712 0.01076 0.00294 -0.0599 0.3870 0.1074
1.000 0.5991 0.01079 0.00295 -0.0596 0.3819 0.1134
1.250 0.6268 0.01080 0.00295 -0.0593 0.3769 0.1233
1.500 0.6545 0.01073 0.00294 -0.0590 0.3722 0.1411
1.750 0.6835 0.01141 0.00379 -0.0588 0.3670 0.2034
2.000 0.7111 0.01141 0.00369 -0.0586 0.3620 0.2163
2.250 0.7392 0.01164 0.00391 -0.0583 0.3576 0.2210
2.500 0.7670 0.01181 0.00405 -0.0580 0.3521 0.2255
2.750 0.7945 0.01196 0.00413 -0.0578 0.3459 0.2285
3.000 0.8221 0.01208 0.00427 -0.0575 0.3395 0.2301
3.250 0.8494 0.01223 0.00440 -0.0573 0.3314 0.2319
3.500 0.8768 0.01235 0.00450 -0.0570 0.3240 0.2339
3.750 0.9039 0.01248 0.00458 -0.0568 0.3148 0.2357
4.000 0.9309 0.01263 0.00468 -0.0565 0.3049 0.2374
4.250 0.9576 0.01278 0.00477 -0.0563 0.2923 0.2383
4.500 0.9838 0.01296 0.00489 -0.0560 0.2758 0.2389
4.750 1.0094 0.01323 0.00508 -0.0557 0.2561 0.2395
5.000 1.0347 0.01354 0.00532 -0.0554 0.2405 0.2406
5.250 1.0601 0.01383 0.00557 -0.0550 0.2297 0.2417
5.500 1.0852 0.01416 0.00585 -0.0547 0.2202 0.2430
5.750 1.1107 0.01440 0.00609 -0.0543 0.2137 0.2444
6.000 1.1359 0.01466 0.00633 -0.0540 0.2081 0.2450
6.250 1.1607 0.01494 0.00660 -0.0536 0.2030 0.2451
6.500 1.1863 0.01510 0.00680 -0.0532 0.1998 0.2452
6.750 1.2115 0.01530 0.00702 -0.0529 0.1963 0.2453
7.000 1.2361 0.01554 0.00728 -0.0525 0.1928 0.2454
7.250 1.2604 0.01581 0.00757 -0.0520 0.1899 0.2455
7.500 1.2843 0.01610 0.00788 -0.0515 0.1870 0.2456
7.750 1.3090 0.01629 0.00813 -0.0511 0.1854 0.2458
8.000 1.3333 0.01651 0.00841 -0.0507 0.1835 0.2459
8.250 1.3572 0.01675 0.00870 -0.0502 0.1814 0.2461
8.500 1.3806 0.01701 0.00902 -0.0497 0.1792 0.2463
8.750 1.4034 0.01732 0.00938 -0.0491 0.1770 0.2465
9.000 1.4255 0.01766 0.00976 -0.0484 0.1748 0.2467
9.250 1.4465 0.01806 0.01019 -0.0477 0.1719 0.2468
9.500 1.4686 0.01836 0.01055 -0.0470 0.1699 0.2471
9.750 1.4904 0.01865 0.01092 -0.0463 0.1666 0.2473
10.000 1.5104 0.01904 0.01134 -0.0455 0.1617 0.2476
10.250 1.5280 0.01957 0.01188 -0.0444 0.1572 0.2480
10.500 1.5471 0.01996 0.01233 -0.0434 0.1530 0.2484
10.750 1.5634 0.02047 0.01287 -0.0422 0.1467 0.2488
11.000 1.5743 0.02109 0.01352 -0.0402 0.1417 0.2490
11.250 1.5834 0.02173 0.01421 -0.0380 0.1356 0.2493
11.500 1.5880 0.02277 0.01525 -0.0358 0.1290 0.2493
11.750 1.5920 0.02404 0.01654 -0.0341 0.1188 0.2497
12.000 1.5799 0.02675 0.01910 -0.0321 0.0946 0.2498
12.250 1.5620 0.03035 0.02260 -0.0307 0.0732 0.2499
12.500 1.5556 0.03315 0.02541 -0.0298 0.0660 0.2498
12.750 1.5497 0.03603 0.02834 -0.0292 0.0601 0.2501
13.000 1.5506 0.03836 0.03075 -0.0288 0.0579 0.2502
13.250 1.5473 0.04122 0.03367 -0.0286 0.0549 0.2503
13.500 1.5410 0.04449 0.03701 -0.0286 0.0517 0.2505
13.750 1.5388 0.04743 0.04004 -0.0287 0.0498 0.2506
14.000 1.5378 0.05030 0.04301 -0.0289 0.0477 0.2510
14.250 1.5316 0.05388 0.04666 -0.0294 0.0440 0.2513
14.500 1.5229 0.05792 0.05076 -0.0301 0.0406 0.2513
14.750 1.5186 0.06157 0.05449 -0.0309 0.0372 0.2515
15.000 1.4928 0.06830 0.06118 -0.0328 0.0214 0.2517
15.250 1.4767 0.07391 0.06684 -0.0345 0.0180 0.2516
15.500 1.4628 0.07938 0.07241 -0.0363 0.0157 0.2518
15.750 1.4512 0.08467 0.07781 -0.0381 0.0144 0.2517
16.000 1.4397 0.09008 0.08334 -0.0400 0.0134 0.2519
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