NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M25 AIRFOIL (m25-il) Reynolds number: 1,000,000 Max Cl/Cd: 107.76 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m25-il-1000000-n5.txt Download as CSV file: xf-m25-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M25 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2127 0.08630 0.08353 0.0050 0.5221 0.0052
-7.250 -0.2082 0.08304 0.08026 0.0039 0.5195 0.0052
-7.000 -0.2039 0.07981 0.07702 0.0026 0.5171 0.0052
-6.750 -0.2003 0.07662 0.07383 0.0013 0.5147 0.0052
-6.500 -0.1918 0.07347 0.07068 -0.0005 0.5124 0.0046
-6.000 -0.2333 0.08098 0.07797 -0.0015 0.5147 0.0035
-5.750 -0.2131 0.07731 0.07427 -0.0050 0.5122 0.0034
-5.500 -0.1917 0.07398 0.07091 -0.0084 0.5095 0.0035
-5.250 -0.1695 0.07119 0.06808 -0.0113 0.5063 0.0037
-5.000 -0.1460 0.06821 0.06506 -0.0144 0.5034 0.0038
-4.750 -0.1211 0.06508 0.06187 -0.0175 0.5005 0.0041
-4.500 -0.0949 0.06173 0.05846 -0.0207 0.4979 0.0043
-4.250 -0.0670 0.05808 0.05474 -0.0240 0.4953 0.0048
-4.000 -0.0370 0.05391 0.05049 -0.0275 0.4926 0.0052
-3.750 -0.0070 0.05027 0.04676 -0.0305 0.4897 0.0054
-3.500 0.0207 0.04805 0.04446 -0.0325 0.4866 0.0057
-3.250 0.0493 0.04568 0.04200 -0.0345 0.4837 0.0062
-3.000 0.0800 0.04260 0.03883 -0.0366 0.4812 0.0068
-2.750 0.1117 0.03924 0.03534 -0.0386 0.4785 0.0080
-2.500 0.1399 0.03763 0.03366 -0.0398 0.4754 0.0085
-2.250 0.1695 0.03554 0.03146 -0.0411 0.4723 0.0093
-2.000 0.2029 0.03192 0.02764 -0.0425 0.4695 0.0111
-1.750 0.2302 0.03107 0.02674 -0.0432 0.4668 0.0116
-1.500 0.2590 0.02968 0.02527 -0.0439 0.4641 0.0124
-1.250 0.2947 0.02589 0.02121 -0.0445 0.4615 0.0153
-1.000 0.3217 0.02513 0.02038 -0.0451 0.4582 0.0157
-0.750 0.3494 0.02430 0.01947 -0.0456 0.4549 0.0161
-0.500 0.3779 0.02329 0.01838 -0.0460 0.4522 0.0166
-0.250 0.4069 0.02215 0.01715 -0.0464 0.4493 0.0176
0.000 0.4369 0.02076 0.01561 -0.0466 0.4464 0.0189
0.250 0.4674 0.01927 0.01395 -0.0466 0.4434 0.0198
0.500 0.4971 0.01797 0.01249 -0.0466 0.4403 0.0202
1.000 0.5561 0.01475 0.00890 -0.0463 0.4346 0.0212
1.250 0.5843 0.01355 0.00754 -0.0465 0.4314 0.0219
1.500 0.6124 0.01319 0.00711 -0.0468 0.4282 0.0225
1.750 0.6406 0.01283 0.00669 -0.0470 0.4254 0.0233
2.000 0.6689 0.01229 0.00606 -0.0472 0.4226 0.0243
2.250 0.6973 0.01167 0.00534 -0.0473 0.4191 0.0250
2.500 0.7256 0.01117 0.00473 -0.0475 0.4155 0.0253
2.750 0.7538 0.01080 0.00427 -0.0477 0.4119 0.0256
3.000 0.7820 0.01052 0.00396 -0.0479 0.4092 0.0261
3.250 0.8100 0.01019 0.00357 -0.0480 0.4061 0.0257
3.500 0.8377 0.00993 0.00327 -0.0482 0.4027 0.0253
3.750 0.8653 0.00974 0.00305 -0.0483 0.3994 0.0250
4.000 0.8928 0.00961 0.00289 -0.0485 0.3962 0.0248
4.250 0.9204 0.00950 0.00278 -0.0486 0.3930 0.0246
4.500 0.9480 0.00944 0.00272 -0.0489 0.3884 0.0245
4.750 0.9756 0.00946 0.00272 -0.0491 0.3827 0.0245
5.000 1.0033 0.00950 0.00273 -0.0494 0.3724 0.0246
5.250 1.0309 0.00961 0.00280 -0.0498 0.3596 0.0251
5.500 1.0582 0.00982 0.00294 -0.0501 0.3424 0.0264
5.750 1.0853 0.01015 0.00317 -0.0506 0.3200 0.0281
6.000 1.1118 0.01063 0.00352 -0.0511 0.2889 0.0290
6.250 1.1264 0.01405 0.00582 -0.0521 0.0796 0.0304
6.500 1.1491 0.01504 0.00663 -0.0524 0.0185 0.0320
6.750 1.1746 0.01537 0.00699 -0.0526 0.0146 0.0350
7.000 1.1996 0.01579 0.00744 -0.0527 0.0108 0.0375
7.250 1.2244 0.01617 0.00792 -0.0529 0.0095 0.0705
7.500 1.2486 0.01655 0.00847 -0.0530 0.0084 0.1893
8.000 1.3255 0.01682 0.01013 -0.0604 0.0059 1.0000
8.250 1.3450 0.01743 0.01079 -0.0599 0.0053 1.0000
8.500 1.3635 0.01812 0.01153 -0.0592 0.0049 1.0000
8.750 1.3803 0.01889 0.01234 -0.0585 0.0045 1.0000
9.000 1.3919 0.02010 0.01362 -0.0575 0.0040 1.0000
9.250 1.4032 0.02146 0.01506 -0.0572 0.0038 1.0000
9.500 1.4068 0.02317 0.01684 -0.0561 0.0036 1.0000
9.750 1.4108 0.02513 0.01888 -0.0554 0.0034 1.0000
10.000 1.4145 0.02735 0.02119 -0.0550 0.0033 1.0000
10.250 1.4170 0.02978 0.02372 -0.0547 0.0031 1.0000
10.500 1.4192 0.03225 0.02627 -0.0544 0.0030 1.0000
10.750 1.4202 0.03487 0.02896 -0.0540 0.0029 1.0000
11.000 1.4195 0.03767 0.03184 -0.0537 0.0028 1.0000
11.250 1.4186 0.04047 0.03471 -0.0533 0.0027 1.0000
11.500 1.4160 0.04350 0.03782 -0.0530 0.0026 1.0000
11.750 1.4116 0.04673 0.04113 -0.0526 0.0026 1.0000
12.000 1.4050 0.05026 0.04474 -0.0523 0.0025 1.0000
12.250 1.3953 0.05428 0.04886 -0.0520 0.0024 1.0000
12.500 1.3847 0.05854 0.05322 -0.0519 0.0023 1.0000
12.750 1.3822 0.06183 0.05659 -0.0518 0.0023 1.0000
13.000 1.3794 0.06519 0.06003 -0.0517 0.0023 1.0000
13.250 1.3763 0.06864 0.06357 -0.0517 0.0022 1.0000
13.500 1.3736 0.07198 0.06699 -0.0516 0.0022 1.0000
13.750 1.3712 0.07531 0.07040 -0.0515 0.0021 1.0000
14.000 1.3698 0.07844 0.07363 -0.0513 0.0021 1.0000
14.250 1.3690 0.08144 0.07672 -0.0510 0.0020 1.0000
14.500 1.3700 0.08415 0.07949 -0.0506 0.0019 1.0000
14.750 1.3727 0.08662 0.08204 -0.0502 0.0018 1.0000
15.000 1.3764 0.08898 0.08448 -0.0499 0.0017 1.0000
15.250 1.3805 0.09136 0.08694 -0.0497 0.0016 1.0000
15.500 1.3844 0.09377 0.08942 -0.0495 0.0016 1.0000
15.750 1.3878 0.09630 0.09202 -0.0494 0.0015 1.0000
16.000 1.3911 0.09878 0.09459 -0.0493 0.0015 1.0000
16.250 1.3944 0.10123 0.09714 -0.0490 0.0014 1.0000
16.500 1.3969 0.10393 0.09992 -0.0490 0.0014 1.0000
16.750 1.3988 0.10669 0.10279 -0.0490 0.0014 1.0000
17.000 1.3992 0.10969 0.10590 -0.0491 0.0013 1.0000
17.250 1.3979 0.11286 0.10923 -0.0490 0.0013 1.0000
17.500 1.3813 0.11725 0.11403 -0.0448 0.0012 1.0000
17.750 1.3670 0.12312 0.12014 -0.0449 0.0012 1.0000
18.000 1.3504 0.12993 0.12713 -0.0466 0.0011 1.0000
18.250 1.3318 0.13751 0.13491 -0.0493 0.0011 1.0000
18.500 1.3136 0.14535 0.14293 -0.0528 0.0011 1.0000
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