NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M25 AIRFOIL (m25-il) Reynolds number: 200,000 Max Cl/Cd: 78.36 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m25-il-200000.txt Download as CSV file: xf-m25-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3409 0.11989 0.11620 0.0197 0.6957 0.0179
-8.250 -0.3328 0.11783 0.11416 0.0169 0.6903 0.0179
-8.000 -0.3258 0.11584 0.11217 0.0139 0.6854 0.0180
-7.750 -0.3179 0.11395 0.11023 0.0107 0.6815 0.0180
-7.250 -0.2916 0.10817 0.10441 0.0037 0.6716 0.0181
-7.000 -0.2865 0.10115 0.09736 0.0069 0.6678 0.0183
-6.750 -0.2765 0.09712 0.09331 0.0061 0.6638 0.0186
-6.500 -0.2638 0.09370 0.08989 0.0040 0.6589 0.0189
-6.250 -0.2497 0.09055 0.08669 0.0017 0.6548 0.0193
-6.000 -0.2342 0.08751 0.08356 -0.0008 0.6514 0.0198
-5.750 -0.2157 0.08437 0.08043 -0.0039 0.6463 0.0204
-5.500 -0.1955 0.08140 0.07740 -0.0071 0.6419 0.0216
-5.250 -0.1663 0.07948 0.07535 -0.0118 0.6382 0.0228
-5.000 -0.1307 0.07831 0.07404 -0.0180 0.6340 0.0231
-4.750 -0.0978 0.07651 0.07213 -0.0230 0.6292 0.0232
-4.250 -0.0590 0.06816 0.06360 -0.0260 0.6224 0.0236
-4.000 -0.0462 0.06356 0.05904 -0.0262 0.6180 0.0240
-3.750 -0.0257 0.06031 0.05575 -0.0276 0.6138 0.0247
-3.500 -0.0013 0.05755 0.05287 -0.0294 0.6102 0.0256
-3.250 0.0257 0.05509 0.05025 -0.0313 0.6069 0.0275
-3.000 0.0658 0.05394 0.04896 -0.0347 0.6018 0.0299
-2.750 0.1108 0.05383 0.04856 -0.0380 0.5977 0.0305
-2.500 0.1345 0.04985 0.04444 -0.0391 0.5946 0.0310
-2.250 0.1494 0.04572 0.04031 -0.0393 0.5911 0.0318
-2.000 0.1731 0.04324 0.03778 -0.0402 0.5865 0.0335
-1.750 0.2014 0.04135 0.03575 -0.0412 0.5827 0.0364
-1.500 0.2439 0.04161 0.03564 -0.0426 0.5795 0.0401
-1.250 0.2747 0.03923 0.03310 -0.0436 0.5755 0.0411
-1.000 0.2941 0.03607 0.02996 -0.0441 0.5711 0.0426
-0.750 0.3194 0.03436 0.02812 -0.0445 0.5675 0.0459
-0.500 0.3595 0.03517 0.02852 -0.0447 0.5645 0.0523
-0.250 0.3852 0.03233 0.02563 -0.0455 0.5602 0.0537
0.000 0.4087 0.03036 0.02365 -0.0460 0.5559 0.0561
0.250 0.4366 0.02923 0.02235 -0.0462 0.5524 0.0610
0.500 0.4700 0.02870 0.02145 -0.0460 0.5496 0.0666
0.750 0.4946 0.02713 0.01997 -0.0468 0.5450 0.0710
1.000 0.5285 0.02717 0.01971 -0.0467 0.5409 0.0791
1.250 0.5526 0.02533 0.01785 -0.0470 0.5376 0.0821
1.500 0.5837 0.02541 0.01759 -0.0467 0.5345 0.0922
1.750 0.6094 0.02385 0.01616 -0.0474 0.5297 0.0962
2.000 0.6384 0.02330 0.01546 -0.0475 0.5258 0.1074
2.250 0.6663 0.02273 0.01473 -0.0474 0.5228 0.1205
2.500 0.6933 0.02205 0.01396 -0.0475 0.5196 0.1355
2.750 0.7209 0.02164 0.01359 -0.0480 0.5146 0.1624
3.000 0.7479 0.02104 0.01296 -0.0481 0.5108 0.1903
3.250 0.7748 0.02077 0.01258 -0.0481 0.5078 0.2297
3.500 0.8019 0.01993 0.01179 -0.0483 0.5044 0.2525
3.750 0.8397 0.01918 0.01051 -0.0470 0.5004 0.0713
4.000 0.8674 0.01881 0.01002 -0.0468 0.4967 0.0667
4.250 0.8941 0.01849 0.00964 -0.0465 0.4937 0.0666
4.500 0.9204 0.01834 0.00952 -0.0464 0.4904 0.0687
4.750 0.9470 0.01845 0.00976 -0.0466 0.4855 0.0765
5.000 0.9729 0.01837 0.00968 -0.0463 0.4820 0.0796
5.250 0.9998 0.01835 0.00964 -0.0462 0.4792 0.0875
5.500 1.0268 0.01842 0.00986 -0.0463 0.4762 0.1985
5.750 1.0966 0.01755 0.01023 -0.0561 0.4685 1.0000
6.000 1.1219 0.01696 0.00953 -0.0553 0.4580 1.0000
6.250 1.1474 0.01655 0.00901 -0.0546 0.4494 1.0000
6.500 1.1727 0.01626 0.00878 -0.0543 0.4373 1.0000
6.750 1.1979 0.01591 0.00842 -0.0539 0.4221 1.0000
7.000 1.2229 0.01567 0.00822 -0.0537 0.3975 1.0000
7.250 1.2460 0.01590 0.00826 -0.0536 0.3499 1.0000
7.500 1.2598 0.01793 0.00958 -0.0536 0.2355 1.0000
7.750 1.2441 0.02322 0.01380 -0.0525 0.0449 1.0000
8.000 1.2542 0.02474 0.01534 -0.0514 0.0360 1.0000
8.250 1.2632 0.02631 0.01709 -0.0507 0.0338 1.0000
8.500 1.2634 0.02844 0.01938 -0.0495 0.0319 1.0000
8.750 1.2598 0.03092 0.02201 -0.0481 0.0302 1.0000
9.000 1.2560 0.03376 0.02500 -0.0473 0.0289 1.0000
9.250 1.2516 0.03683 0.02824 -0.0467 0.0283 1.0000
9.500 1.2452 0.04022 0.03178 -0.0462 0.0278 1.0000
9.750 1.2370 0.04387 0.03557 -0.0457 0.0275 1.0000
10.000 1.2280 0.04764 0.03951 -0.0454 0.0273 1.0000
10.250 1.2185 0.05149 0.04347 -0.0450 0.0271 1.0000
10.500 1.2097 0.05523 0.04730 -0.0445 0.0269 1.0000
10.750 1.2035 0.05866 0.05080 -0.0440 0.0268 1.0000
11.000 1.2008 0.06162 0.05381 -0.0432 0.0267 1.0000
11.250 1.2024 0.06393 0.05614 -0.0420 0.0267 1.0000
11.500 1.2089 0.06519 0.05737 -0.0398 0.0265 1.0000
11.750 1.2176 0.06697 0.05923 -0.0391 0.0258 1.0000
12.000 1.2303 0.06776 0.06004 -0.0371 0.0253 1.0000
12.250 1.2843 0.06311 0.05524 -0.0292 0.0269 1.0000
12.500 1.2924 0.06497 0.05722 -0.0290 0.0278 1.0000
12.750 1.1305 0.06581 0.05865 -0.0164 0.0265 1.0000
13.000 1.1404 0.06761 0.06056 -0.0156 0.0272 1.0000
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Polar data table (+)
Polar graphs
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