NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M25 AIRFOIL (m25-il) Reynolds number: 1,000,000 Max Cl/Cd: 111.87 at α=6° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m25-il-1000000.txt Download as CSV file: xf-m25-il-1000000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3555 0.10934 0.10665 0.0257 0.5779 0.0075
-7.750 -0.3479 0.10638 0.10370 0.0236 0.5747 0.0076
-7.500 -0.3411 0.10351 0.10083 0.0214 0.5712 0.0076
-7.250 -0.3295 0.10040 0.09768 0.0181 0.5677 0.0076
-7.000 -0.3158 0.09717 0.09442 0.0149 0.5640 0.0076
-6.750 -0.3033 0.09320 0.09045 0.0120 0.5611 0.0077
-6.500 -0.2918 0.08958 0.08682 0.0102 0.5576 0.0078
-6.250 -0.2762 0.08653 0.08373 0.0076 0.5541 0.0080
-6.000 -0.2587 0.08352 0.08069 0.0047 0.5506 0.0081
-5.750 -0.2397 0.08048 0.07762 0.0018 0.5475 0.0084
-5.500 -0.2194 0.07739 0.07450 -0.0013 0.5441 0.0087
-5.250 -0.1978 0.07425 0.07132 -0.0044 0.5407 0.0092
-5.000 -0.1733 0.07093 0.06793 -0.0077 0.5375 0.0102
-4.750 -0.1424 0.06761 0.06453 -0.0121 0.5344 0.0104
-4.500 -0.1130 0.06417 0.06103 -0.0160 0.5315 0.0105
-4.250 -0.0938 0.06003 0.05686 -0.0179 0.5284 0.0108
-4.000 -0.0705 0.05755 0.05433 -0.0199 0.5250 0.0111
-3.750 -0.0445 0.05514 0.05184 -0.0220 0.5216 0.0117
-3.500 -0.0166 0.05263 0.04926 -0.0243 0.5187 0.0131
-3.250 0.0195 0.05007 0.04660 -0.0272 0.5157 0.0138
-3.000 0.0528 0.04727 0.04369 -0.0296 0.5125 0.0139
-2.750 0.0852 0.04435 0.04064 -0.0316 0.5095 0.0140
2.250 0.2346 0.01459 0.01111 0.0320 0.4510 0.0303
2.500 0.6912 0.01199 0.00586 -0.0399 0.4457 0.0326
2.750 0.7193 0.01170 0.00556 -0.0402 0.4423 0.0344
3.000 0.7476 0.01138 0.00518 -0.0404 0.4390 0.0371
3.250 0.7762 0.01166 0.00541 -0.0406 0.4353 0.0404
3.500 0.8039 0.01069 0.00437 -0.0408 0.4325 0.0448
3.750 0.8321 0.01060 0.00429 -0.0411 0.4293 0.0498
4.000 0.8600 0.01039 0.00409 -0.0414 0.4254 0.0583
4.250 0.8870 0.00954 0.00306 -0.0407 0.4212 0.0339
4.500 0.9147 0.00941 0.00292 -0.0408 0.4162 0.0319
4.750 0.9422 0.00933 0.00279 -0.0409 0.4082 0.0320
5.000 0.9698 0.00924 0.00271 -0.0411 0.4030 0.0341
5.250 0.9976 0.00927 0.00273 -0.0414 0.3943 0.0358
5.500 1.0255 0.00933 0.00278 -0.0417 0.3840 0.0369
5.750 1.0532 0.00945 0.00287 -0.0421 0.3734 0.0384
6.000 1.0807 0.00966 0.00301 -0.0425 0.3539 0.0399
6.250 1.1070 0.01031 0.00340 -0.0431 0.3049 0.0419
6.500 1.1187 0.01468 0.00638 -0.0445 0.0238 0.0745
6.750 1.1520 0.01388 0.00718 -0.0466 0.0173 0.9926
7.000 1.2172 0.01463 0.00797 -0.0559 0.0146 1.0000
7.250 1.2396 0.01520 0.00860 -0.0556 0.0132 1.0000
7.500 1.2598 0.01608 0.00957 -0.0552 0.0119 1.0000
7.750 1.2768 0.01726 0.01089 -0.0545 0.0111 1.0000
8.000 1.2967 0.01786 0.01152 -0.0539 0.0106 1.0000
8.250 1.3148 0.01859 0.01229 -0.0532 0.0099 1.0000
8.500 1.3295 0.01959 0.01335 -0.0523 0.0093 1.0000
8.750 1.3403 0.02089 0.01475 -0.0514 0.0089 1.0000
9.000 1.3452 0.02290 0.01685 -0.0511 0.0086 1.0000
9.250 1.3405 0.02510 0.01911 -0.0492 0.0084 1.0000
9.500 1.3398 0.02763 0.02172 -0.0485 0.0082 1.0000
9.750 1.3386 0.03041 0.02458 -0.0481 0.0080 1.0000
10.000 1.3343 0.03358 0.02782 -0.0477 0.0078 1.0000
10.250 1.3258 0.03717 0.03150 -0.0472 0.0076 1.0000
10.500 1.3136 0.04110 0.03551 -0.0464 0.0074 1.0000
10.750 1.2992 0.04505 0.03954 -0.0453 0.0073 1.0000
11.000 1.2912 0.04808 0.04262 -0.0437 0.0071 1.0000
11.250 1.2979 0.05011 0.04471 -0.0435 0.0069 1.0000
11.500 1.3023 0.05221 0.04688 -0.0428 0.0068 1.0000
11.750 1.3081 0.05419 0.04891 -0.0421 0.0066 1.0000
12.000 1.3142 0.05608 0.05085 -0.0413 0.0063 1.0000
12.250 1.3213 0.05774 0.05256 -0.0403 0.0061 1.0000
12.500 1.3308 0.05874 0.05359 -0.0383 0.0060 1.0000
12.750 1.3443 0.05917 0.05404 -0.0358 0.0059 1.0000
13.000 1.3603 0.05946 0.05439 -0.0332 0.0058 1.0000
13.250 1.3766 0.06002 0.05502 -0.0307 0.0056 1.0000
13.500 1.3904 0.06114 0.05622 -0.0287 0.0055 1.0000
13.750 1.4605 0.06565 0.06108 -0.0191 0.0070 1.0000
14.000 1.4513 0.06890 0.06449 -0.0186 0.0070 1.0000
14.250 1.4407 0.07213 0.06787 -0.0184 0.0070 1.0000
14.500 1.4287 0.07491 0.07080 -0.0187 0.0069 1.0000
14.750 1.4192 0.07714 0.07316 -0.0195 0.0065 1.0000
15.000 1.4086 0.08092 0.07710 -0.0200 0.0062 1.0000
15.250 1.3979 0.08498 0.08131 -0.0206 0.0060 1.0000
15.500 1.3868 0.08941 0.08589 -0.0214 0.0058 1.0000
15.750 1.3754 0.09406 0.09068 -0.0226 0.0057 1.0000
16.000 1.3634 0.09891 0.09567 -0.0239 0.0056 1.0000
16.250 1.3508 0.10407 0.10095 -0.0256 0.0055 1.0000
16.500 1.3378 0.10950 0.10652 -0.0275 0.0054 1.0000
16.750 1.3242 0.11522 0.11237 -0.0298 0.0054 1.0000
17.000 1.3105 0.12111 0.11839 -0.0324 0.0053 1.0000
17.250 1.2965 0.12735 0.12476 -0.0353 0.0053 1.0000
17.500 1.2818 0.13403 0.13156 -0.0386 0.0052 1.0000
17.750 1.2675 0.14091 0.13857 -0.0422 0.0052 1.0000
18.000 1.2528 0.14823 0.14601 -0.0463 0.0052 1.0000
18.250 1.2373 0.15609 0.15399 -0.0509 0.0052 1.0000
18.500 1.2217 0.16444 0.16246 -0.0559 0.0053 1.0000
18.750 1.2052 0.17360 0.17175 -0.0615 0.0053 1.0000
19.000 1.1884 0.18365 0.18192 -0.0679 0.0054 1.0000
19.250 1.1715 0.19483 0.19321 -0.0749 0.0055 1.0000
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Polar data table (+)
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