NACA M15 AIRFOIL (m15-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M15 AIRFOIL (m15-il) Reynolds number: 500,000 Max Cl/Cd: 101.57 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m15-il-500000.txt Download as CSV file: xf-m15-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3788 0.08442 0.08225 -0.0356 1.0000 0.0344
-8.500 -0.3782 0.08046 0.07834 -0.0389 0.9966 0.0352
-8.250 -0.3863 0.06748 0.06515 -0.0589 0.9458 0.0378
-8.000 -0.3971 0.05984 0.05731 -0.0610 0.9195 0.0382
-7.750 -0.3897 0.05772 0.05511 -0.0599 0.8984 0.0385
-7.500 -0.3796 0.05569 0.05298 -0.0592 0.8791 0.0389
-7.250 -0.3676 0.05361 0.05078 -0.0588 0.8614 0.0393
-7.000 -0.3540 0.05145 0.04849 -0.0587 0.8450 0.0400
-6.750 -0.3254 0.02907 0.02551 -0.0601 0.8015 0.0453
-6.500 -0.3115 0.02637 0.02275 -0.0596 0.7894 0.0458
-6.250 -0.2945 0.02459 0.02088 -0.0590 0.7775 0.0463
-6.000 -0.2766 0.02290 0.01910 -0.0586 0.7662 0.0469
-5.750 -0.2584 0.02114 0.01719 -0.0580 0.7563 0.0479
-5.500 -0.2402 0.01899 0.01484 -0.0574 0.7472 0.0500
-5.250 -0.2285 0.01481 0.01005 -0.0559 0.7405 0.0545
-5.000 -0.2316 0.01896 0.01254 -0.0526 0.7485 0.0386
-4.750 -0.2058 0.01787 0.01121 -0.0521 0.7404 0.0384
-4.500 -0.1795 0.01678 0.00993 -0.0518 0.7325 0.0384
-4.250 -0.1531 0.01572 0.00871 -0.0515 0.7256 0.0388
-4.000 -0.1261 0.01496 0.00791 -0.0514 0.7185 0.0395
-3.750 -0.0990 0.01437 0.00720 -0.0512 0.7123 0.0400
-3.500 -0.0715 0.01377 0.00653 -0.0511 0.7060 0.0405
-3.250 -0.0441 0.01325 0.00595 -0.0509 0.6998 0.0410
-3.000 -0.0168 0.01283 0.00545 -0.0507 0.6944 0.0416
-2.750 0.0107 0.01241 0.00500 -0.0506 0.6887 0.0423
-2.500 0.0381 0.01205 0.00459 -0.0504 0.6831 0.0430
-2.250 0.0654 0.01177 0.00424 -0.0502 0.6778 0.0438
-2.000 0.0932 0.01153 0.00399 -0.0502 0.6721 0.0448
-1.750 0.1197 0.01107 0.00352 -0.0499 0.6669 0.0467
-1.500 0.1470 0.01087 0.00328 -0.0497 0.6620 0.0485
-1.250 0.1748 0.01066 0.00309 -0.0497 0.6564 0.0507
-1.000 0.2027 0.01052 0.00291 -0.0496 0.6507 0.0531
-0.750 0.2300 0.01033 0.00272 -0.0494 0.6456 0.0592
-0.500 0.2576 0.01008 0.00258 -0.0493 0.6400 0.0798
-0.250 0.2843 0.00978 0.00251 -0.0492 0.6343 0.1466
0.000 0.3107 0.00949 0.00248 -0.0491 0.6288 0.2363
0.250 0.3215 0.00760 0.00242 -0.0461 0.6232 0.7801
0.500 0.3492 0.00751 0.00268 -0.0450 0.6173 0.9410
0.750 0.3873 0.00772 0.00283 -0.0468 0.6110 0.9699
1.000 0.4593 0.00800 0.00304 -0.0560 0.6024 0.9859
1.250 0.5272 0.00819 0.00313 -0.0646 0.5942 0.9969
1.500 0.5675 0.00821 0.00310 -0.0674 0.5857 1.0000
1.750 0.5936 0.00823 0.00307 -0.0671 0.5778 1.0000
2.000 0.6197 0.00824 0.00304 -0.0669 0.5690 1.0000
2.250 0.6457 0.00826 0.00302 -0.0666 0.5600 1.0000
2.500 0.6715 0.00831 0.00300 -0.0662 0.5503 1.0000
2.750 0.6973 0.00834 0.00301 -0.0659 0.5392 1.0000
3.000 0.7229 0.00839 0.00303 -0.0656 0.5285 1.0000
3.250 0.7480 0.00848 0.00305 -0.0651 0.5172 1.0000
3.500 0.7729 0.00857 0.00310 -0.0646 0.5054 1.0000
3.750 0.7977 0.00867 0.00316 -0.0641 0.4934 1.0000
4.000 0.8220 0.00880 0.00324 -0.0635 0.4814 1.0000
4.250 0.8461 0.00897 0.00334 -0.0629 0.4698 1.0000
4.500 0.8701 0.00912 0.00345 -0.0623 0.4593 1.0000
4.750 0.8943 0.00927 0.00358 -0.0616 0.4506 1.0000
5.000 0.9179 0.00946 0.00373 -0.0609 0.4425 1.0000
5.250 0.9420 0.00961 0.00388 -0.0603 0.4356 1.0000
5.500 0.9656 0.00979 0.00405 -0.0596 0.4287 1.0000
5.750 0.9891 0.00998 0.00424 -0.0589 0.4223 1.0000
6.000 1.0128 0.01014 0.00441 -0.0582 0.4152 1.0000
6.250 1.0356 0.01038 0.00462 -0.0573 0.4088 1.0000
6.500 1.0596 0.01051 0.00481 -0.0567 0.4027 1.0000
6.750 1.0820 0.01076 0.00502 -0.0558 0.3944 1.0000
7.000 1.1050 0.01091 0.00520 -0.0550 0.3824 1.0000
7.250 1.1274 0.01110 0.00539 -0.0541 0.3693 1.0000
7.500 1.1496 0.01132 0.00560 -0.0532 0.3569 1.0000
7.750 1.1715 0.01157 0.00583 -0.0522 0.3447 1.0000
8.000 1.1931 0.01183 0.00607 -0.0512 0.3296 1.0000
8.250 1.2144 0.01212 0.00634 -0.0502 0.3134 1.0000
8.500 1.2352 0.01246 0.00665 -0.0492 0.2951 1.0000
8.750 1.2538 0.01294 0.00703 -0.0478 0.2719 1.0000
9.000 1.2707 0.01355 0.00751 -0.0462 0.2430 1.0000
9.250 1.2850 0.01431 0.00811 -0.0443 0.2110 1.0000
9.500 1.2975 0.01517 0.00881 -0.0422 0.1806 1.0000
9.750 1.3083 0.01610 0.00958 -0.0398 0.1526 1.0000
10.000 1.3142 0.01723 0.01051 -0.0368 0.1172 1.0000
10.250 1.3081 0.01877 0.01177 -0.0320 0.0761 1.0000
10.500 1.3065 0.02029 0.01315 -0.0284 0.0542 1.0000
10.750 1.3099 0.02169 0.01450 -0.0257 0.0437 1.0000
11.000 1.3157 0.02305 0.01589 -0.0237 0.0376 1.0000
11.250 1.3202 0.02461 0.01748 -0.0218 0.0337 1.0000
11.500 1.3282 0.02602 0.01895 -0.0205 0.0311 1.0000
11.750 1.3308 0.02795 0.02091 -0.0190 0.0290 1.0000
12.000 1.3371 0.02967 0.02271 -0.0179 0.0276 1.0000
12.250 1.3437 0.03141 0.02454 -0.0170 0.0262 1.0000
12.500 1.3479 0.03345 0.02663 -0.0162 0.0250 1.0000
12.750 1.3458 0.03616 0.02941 -0.0154 0.0240 1.0000
13.000 1.3464 0.03868 0.03202 -0.0147 0.0233 1.0000
13.250 1.3495 0.04099 0.03443 -0.0142 0.0227 1.0000
13.500 1.3511 0.04350 0.03703 -0.0138 0.0220 1.0000
13.750 1.3515 0.04616 0.03977 -0.0135 0.0214 1.0000
14.000 1.3506 0.04901 0.04270 -0.0132 0.0209 1.0000
14.250 1.3481 0.05210 0.04587 -0.0131 0.0205 1.0000
14.500 1.3425 0.05561 0.04944 -0.0130 0.0200 1.0000
14.750 1.3331 0.05961 0.05352 -0.0129 0.0196 1.0000
15.000 1.3324 0.06273 0.05673 -0.0131 0.0193 1.0000
15.250 1.3331 0.06574 0.05984 -0.0133 0.0189 1.0000
15.500 1.3332 0.06888 0.06307 -0.0136 0.0185 1.0000
15.750 1.3325 0.07214 0.06642 -0.0140 0.0181 1.0000
16.000 1.3318 0.07545 0.06982 -0.0145 0.0177 1.0000
16.250 1.3307 0.07883 0.07326 -0.0150 0.0174 1.0000
16.500 1.3296 0.08223 0.07673 -0.0156 0.0171 1.0000
16.750 1.3284 0.08563 0.08019 -0.0162 0.0168 1.0000
17.000 1.3268 0.08904 0.08364 -0.0167 0.0165 1.0000
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