NACA M15 AIRFOIL (m15-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M15 AIRFOIL (m15-il) Reynolds number: 200,000 Max Cl/Cd: 72.32 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m15-il-200000-n5.txt Download as CSV file: xf-m15-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M15 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.3893 0.09027 0.08673 -0.0373 1.0000 0.0387
-9.000 -0.3822 0.08760 0.08409 -0.0373 1.0000 0.0378
-8.750 -0.3823 0.08378 0.08031 -0.0393 1.0000 0.0370
-8.500 -0.3807 0.07828 0.07482 -0.0454 0.9699 0.0361
-8.250 -0.3800 0.07206 0.06853 -0.0523 0.9400 0.0355
-8.000 -0.3767 0.06675 0.06310 -0.0566 0.9169 0.0358
-7.750 -0.3738 0.06160 0.05779 -0.0595 0.8971 0.0360
-7.500 -0.3706 0.05617 0.05215 -0.0614 0.8797 0.0358
-7.250 -0.3668 0.05033 0.04602 -0.0626 0.8640 0.0353
-7.000 -0.3615 0.04426 0.03956 -0.0629 0.8497 0.0350
-6.750 -0.3536 0.03873 0.03357 -0.0623 0.8367 0.0350
-6.500 -0.3400 0.03508 0.02952 -0.0614 0.8248 0.0357
-6.250 -0.3252 0.03154 0.02551 -0.0602 0.8143 0.0363
-6.000 -0.3080 0.02838 0.02185 -0.0591 0.8037 0.0365
-5.750 -0.2880 0.02598 0.01900 -0.0580 0.7943 0.0367
-5.500 -0.2658 0.02413 0.01673 -0.0572 0.7848 0.0371
-5.250 -0.2421 0.02263 0.01488 -0.0565 0.7757 0.0375
-5.000 -0.2175 0.02141 0.01333 -0.0559 0.7669 0.0379
-4.750 -0.1920 0.02042 0.01207 -0.0554 0.7567 0.0385
-4.500 -0.1661 0.01971 0.01106 -0.0549 0.7467 0.0392
-4.250 -0.1403 0.01875 0.00996 -0.0545 0.7364 0.0399
-4.000 -0.1140 0.01801 0.00912 -0.0543 0.7276 0.0404
-3.750 -0.0874 0.01738 0.00839 -0.0540 0.7203 0.0410
-3.500 -0.0606 0.01683 0.00776 -0.0537 0.7135 0.0416
-3.250 -0.0337 0.01633 0.00720 -0.0535 0.7065 0.0424
-3.000 -0.0071 0.01589 0.00666 -0.0532 0.7006 0.0432
-2.750 0.0199 0.01547 0.00621 -0.0530 0.6938 0.0442
-2.500 0.0466 0.01511 0.00578 -0.0527 0.6878 0.0452
-2.250 0.0735 0.01482 0.00542 -0.0525 0.6819 0.0467
-2.000 0.1001 0.01446 0.00508 -0.0523 0.6754 0.0488
-1.750 0.1269 0.01422 0.00480 -0.0520 0.6698 0.0511
-1.500 0.1539 0.01400 0.00455 -0.0519 0.6639 0.0536
-1.250 0.1810 0.01379 0.00431 -0.0516 0.6576 0.0568
-1.000 0.2078 0.01359 0.00409 -0.0514 0.6521 0.0629
-0.750 0.2348 0.01338 0.00396 -0.0513 0.6455 0.0760
-0.500 0.2616 0.01318 0.00385 -0.0511 0.6396 0.1053
-0.250 0.2880 0.01297 0.00376 -0.0509 0.6338 0.1463
0.000 0.3142 0.01271 0.00372 -0.0507 0.6269 0.2079
0.250 0.3371 0.01208 0.00363 -0.0500 0.6209 0.3796
0.750 0.4594 0.01117 0.00424 -0.0624 0.6056 0.9902
1.000 0.5113 0.01123 0.00421 -0.0676 0.5972 1.0000
1.250 0.5365 0.01128 0.00417 -0.0672 0.5899 1.0000
1.500 0.5617 0.01132 0.00416 -0.0667 0.5823 1.0000
1.750 0.5867 0.01138 0.00414 -0.0662 0.5741 1.0000
2.000 0.6117 0.01143 0.00416 -0.0657 0.5658 1.0000
2.250 0.6365 0.01150 0.00417 -0.0652 0.5574 1.0000
2.500 0.6613 0.01158 0.00422 -0.0646 0.5484 1.0000
2.750 0.6857 0.01167 0.00424 -0.0640 0.5396 1.0000
3.000 0.7103 0.01175 0.00432 -0.0635 0.5296 1.0000
3.250 0.7346 0.01186 0.00438 -0.0628 0.5204 1.0000
3.500 0.7587 0.01198 0.00446 -0.0621 0.5105 1.0000
3.750 0.7828 0.01211 0.00457 -0.0615 0.5008 1.0000
4.000 0.8066 0.01226 0.00467 -0.0608 0.4920 1.0000
4.250 0.8304 0.01242 0.00481 -0.0601 0.4826 1.0000
4.500 0.8539 0.01260 0.00496 -0.0593 0.4742 1.0000
4.750 0.8774 0.01279 0.00512 -0.0586 0.4657 1.0000
5.000 0.9005 0.01299 0.00531 -0.0578 0.4560 1.0000
5.250 0.9231 0.01324 0.00549 -0.0568 0.4457 1.0000
5.500 0.9459 0.01347 0.00571 -0.0560 0.4353 1.0000
5.750 0.9688 0.01371 0.00596 -0.0552 0.4274 1.0000
6.000 0.9917 0.01396 0.00621 -0.0544 0.4200 1.0000
6.250 1.0144 0.01422 0.00648 -0.0535 0.4131 1.0000
6.500 1.0373 0.01447 0.00678 -0.0527 0.4062 1.0000
6.750 1.0597 0.01477 0.00707 -0.0519 0.3999 1.0000
7.000 1.0826 0.01501 0.00740 -0.0511 0.3924 1.0000
7.250 1.1048 0.01532 0.00772 -0.0502 0.3852 1.0000
7.500 1.1274 0.01559 0.00808 -0.0494 0.3781 1.0000
7.750 1.1487 0.01591 0.00842 -0.0484 0.3689 1.0000
8.000 1.1695 0.01621 0.00878 -0.0473 0.3559 1.0000
8.250 1.1892 0.01655 0.00914 -0.0461 0.3409 1.0000
8.500 1.2078 0.01693 0.00953 -0.0447 0.3238 1.0000
8.750 1.2247 0.01740 0.00996 -0.0431 0.3013 1.0000
9.000 1.2386 0.01801 0.01048 -0.0411 0.2735 1.0000
9.250 1.2505 0.01875 0.01110 -0.0389 0.2456 1.0000
9.500 1.2606 0.01959 0.01185 -0.0365 0.2193 1.0000
9.750 1.2679 0.02052 0.01268 -0.0338 0.1952 1.0000
10.250 1.2752 0.02272 0.01472 -0.0276 0.1543 1.0000
10.500 1.2783 0.02407 0.01600 -0.0250 0.1332 1.0000
10.750 1.2791 0.02571 0.01753 -0.0225 0.1071 1.0000
11.250 1.2751 0.02988 0.02148 -0.0183 0.0658 1.0000
11.500 1.2757 0.03197 0.02357 -0.0169 0.0559 1.0000
11.750 1.2767 0.03415 0.02577 -0.0156 0.0491 1.0000
12.000 1.2789 0.03632 0.02800 -0.0147 0.0436 1.0000
12.250 1.2803 0.03863 0.03037 -0.0138 0.0395 1.0000
12.500 1.2813 0.04106 0.03287 -0.0131 0.0361 1.0000
12.750 1.2808 0.04370 0.03558 -0.0126 0.0336 1.0000
13.000 1.2816 0.04626 0.03826 -0.0121 0.0315 1.0000
13.250 1.2804 0.04909 0.04118 -0.0118 0.0299 1.0000
13.500 1.2763 0.05230 0.04446 -0.0116 0.0286 1.0000
13.750 1.2749 0.05530 0.04757 -0.0115 0.0273 1.0000
14.000 1.2732 0.05843 0.05082 -0.0115 0.0262 1.0000
14.250 1.2709 0.06171 0.05420 -0.0117 0.0251 1.0000
14.500 1.2672 0.06524 0.05782 -0.0120 0.0242 1.0000
14.750 1.2617 0.06911 0.06176 -0.0125 0.0235 1.0000
15.000 1.2582 0.07280 0.06557 -0.0131 0.0228 1.0000
15.250 1.2556 0.07640 0.06930 -0.0136 0.0221 1.0000
15.500 1.2527 0.08012 0.07312 -0.0143 0.0215 1.0000
15.750 1.2499 0.08387 0.07698 -0.0150 0.0210 1.0000
16.000 1.2471 0.08766 0.08086 -0.0159 0.0205 1.0000
16.250 1.2442 0.09149 0.08477 -0.0168 0.0200 1.0000
16.500 1.2411 0.09536 0.08871 -0.0177 0.0195 1.0000
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