NACA M15 AIRFOIL (m15-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M15 AIRFOIL (m15-il) Reynolds number: 100,000 Max Cl/Cd: 53.08 at α=7.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m15-il-100000-n5.txt Download as CSV file: xf-m15-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3768 0.09778 0.09291 -0.0396 1.0000 0.0781
-8.750 -0.3919 0.09417 0.08939 -0.0442 1.0000 0.0784
-8.500 -0.4047 0.09102 0.08629 -0.0459 1.0000 0.0786
-8.000 -0.3994 0.07608 0.07125 -0.0511 0.9860 0.0559
-7.750 -0.3787 0.07282 0.06799 -0.0526 0.9750 0.0548
-7.500 -0.3620 0.06824 0.06332 -0.0565 0.9610 0.0536
-7.250 -0.3470 0.06316 0.05808 -0.0608 0.9461 0.0527
-7.000 -0.3327 0.05828 0.05298 -0.0641 0.9313 0.0529
-6.750 -0.3186 0.05375 0.04819 -0.0661 0.9171 0.0529
-6.500 -0.3041 0.04946 0.04362 -0.0671 0.9039 0.0521
-6.250 -0.2894 0.04521 0.03901 -0.0675 0.8917 0.0513
-6.000 -0.2741 0.04120 0.03456 -0.0672 0.8802 0.0507
-5.750 -0.2578 0.03783 0.03076 -0.0664 0.8682 0.0504
-5.500 -0.2391 0.03498 0.02749 -0.0656 0.8580 0.0505
-5.250 -0.2181 0.03290 0.02505 -0.0649 0.8486 0.0515
-5.000 -0.1967 0.03095 0.02272 -0.0641 0.8384 0.0527
-4.750 -0.1733 0.02903 0.02035 -0.0633 0.8306 0.0534
-4.500 -0.1495 0.02735 0.01831 -0.0626 0.8208 0.0537
-4.250 -0.1242 0.02586 0.01646 -0.0620 0.8130 0.0541
-4.000 -0.0988 0.02461 0.01490 -0.0613 0.8029 0.0547
-3.750 -0.0729 0.02351 0.01351 -0.0607 0.7932 0.0554
-3.500 -0.0466 0.02255 0.01232 -0.0601 0.7838 0.0563
-3.250 -0.0209 0.02180 0.01156 -0.0597 0.7742 0.0582
-3.000 0.0055 0.02122 0.01088 -0.0593 0.7667 0.0605
-2.750 0.0320 0.02062 0.01020 -0.0590 0.7583 0.0627
-2.500 0.0586 0.01999 0.00946 -0.0585 0.7513 0.0645
-2.250 0.0849 0.01947 0.00884 -0.0580 0.7435 0.0665
-2.000 0.1105 0.01892 0.00832 -0.0575 0.7363 0.0693
-1.750 0.1364 0.01854 0.00791 -0.0570 0.7292 0.0733
-1.500 0.1623 0.01820 0.00754 -0.0565 0.7217 0.0803
-1.250 0.1884 0.01788 0.00720 -0.0560 0.7152 0.0914
-1.000 0.2141 0.01753 0.00693 -0.0556 0.7074 0.1097
-0.750 0.2404 0.01715 0.00665 -0.0551 0.7013 0.1474
-0.500 0.2661 0.01674 0.00658 -0.0549 0.6932 0.2291
-0.250 0.3870 0.01498 0.00684 -0.0722 0.6860 1.0000
0.000 0.4114 0.01505 0.00677 -0.0716 0.6778 1.0000
0.250 0.4359 0.01510 0.00666 -0.0709 0.6705 1.0000
0.500 0.4602 0.01519 0.00663 -0.0703 0.6626 1.0000
0.750 0.4846 0.01527 0.00657 -0.0696 0.6550 1.0000
1.000 0.5087 0.01537 0.00658 -0.0689 0.6470 1.0000
1.250 0.5330 0.01546 0.00657 -0.0682 0.6391 1.0000
1.500 0.5570 0.01558 0.00661 -0.0675 0.6311 1.0000
1.750 0.5811 0.01569 0.00664 -0.0668 0.6229 1.0000
2.000 0.6050 0.01583 0.00672 -0.0661 0.6146 1.0000
2.250 0.6290 0.01596 0.00678 -0.0653 0.6063 1.0000
2.500 0.6527 0.01611 0.00690 -0.0646 0.5978 1.0000
2.750 0.6766 0.01625 0.00698 -0.0638 0.5895 1.0000
3.000 0.7001 0.01642 0.00715 -0.0630 0.5805 1.0000
3.250 0.7240 0.01656 0.00723 -0.0623 0.5725 1.0000
3.500 0.7472 0.01676 0.00746 -0.0615 0.5632 1.0000
3.750 0.7712 0.01691 0.00755 -0.0607 0.5557 1.0000
4.000 0.7940 0.01714 0.00781 -0.0599 0.5461 1.0000
4.250 0.8179 0.01730 0.00794 -0.0591 0.5386 1.0000
4.500 0.8406 0.01755 0.00823 -0.0582 0.5293 1.0000
4.750 0.8641 0.01775 0.00841 -0.0574 0.5216 1.0000
5.000 0.8869 0.01799 0.00870 -0.0566 0.5129 1.0000
5.250 0.9100 0.01822 0.00894 -0.0557 0.5050 1.0000
5.500 0.9329 0.01848 0.00922 -0.0549 0.4968 1.0000
5.750 0.9556 0.01874 0.00951 -0.0540 0.4883 1.0000
6.000 0.9784 0.01895 0.00971 -0.0531 0.4791 1.0000
6.250 1.0000 0.01924 0.01005 -0.0521 0.4684 1.0000
6.500 1.0224 0.01951 0.01032 -0.0511 0.4592 1.0000
6.750 1.0446 0.01984 0.01071 -0.0502 0.4504 1.0000
7.000 1.0668 0.02020 0.01111 -0.0493 0.4422 1.0000
7.250 1.0890 0.02055 0.01152 -0.0484 0.4340 1.0000
7.500 1.1107 0.02095 0.01201 -0.0475 0.4258 1.0000
7.750 1.1327 0.02134 0.01245 -0.0466 0.4177 1.0000
8.000 1.1535 0.02178 0.01301 -0.0456 0.4090 1.0000
8.250 1.1753 0.02219 0.01345 -0.0447 0.4007 1.0000
8.500 1.1946 0.02268 0.01412 -0.0435 0.3910 1.0000
8.750 1.2137 0.02310 0.01461 -0.0422 0.3800 1.0000
9.000 1.2300 0.02353 0.01510 -0.0405 0.3656 1.0000
9.250 1.2432 0.02400 0.01566 -0.0384 0.3478 1.0000
9.500 1.2545 0.02454 0.01628 -0.0360 0.3285 1.0000
9.750 1.2642 0.02517 0.01693 -0.0335 0.3089 1.0000
10.000 1.2719 0.02592 0.01770 -0.0309 0.2887 1.0000
10.250 1.2752 0.02681 0.01859 -0.0277 0.2671 1.0000
10.500 1.2754 0.02796 0.01968 -0.0244 0.2451 1.0000
10.750 1.2748 0.02941 0.02106 -0.0215 0.2215 1.0000
11.000 1.2724 0.03118 0.02276 -0.0190 0.1990 1.0000
11.250 1.2684 0.03329 0.02481 -0.0168 0.1775 1.0000
11.500 1.2635 0.03571 0.02717 -0.0150 0.1570 1.0000
11.750 1.2579 0.03839 0.02981 -0.0136 0.1356 1.0000
12.000 1.2503 0.04145 0.03282 -0.0124 0.1137 1.0000
12.250 1.2402 0.04492 0.03620 -0.0116 0.0949 1.0000
12.500 1.2297 0.04861 0.03983 -0.0110 0.0819 1.0000
12.750 1.2203 0.05232 0.04354 -0.0106 0.0726 1.0000
13.000 1.2115 0.05609 0.04735 -0.0104 0.0662 1.0000
13.250 1.2020 0.06004 0.05134 -0.0104 0.0613 1.0000
13.500 1.1937 0.06399 0.05536 -0.0105 0.0573 1.0000
13.750 1.1867 0.06791 0.05938 -0.0108 0.0536 1.0000
14.000 1.1778 0.07220 0.06374 -0.0113 0.0509 1.0000
14.250 1.1718 0.07619 0.06781 -0.0118 0.0482 1.0000
14.500 1.1678 0.07998 0.07172 -0.0124 0.0457 1.0000
14.750 1.1634 0.08387 0.07570 -0.0130 0.0437 1.0000
15.000 1.1585 0.08783 0.07970 -0.0137 0.0420 1.0000
15.250 1.1575 0.09120 0.08316 -0.0141 0.0402 1.0000
15.500 1.1577 0.09449 0.08657 -0.0146 0.0383 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M15 AIRFOIL (m15-il)