NACA M27 AIRFOIL (m27-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M27 AIRFOIL (m27-il) Reynolds number: 500,000 Max Cl/Cd: 105.04 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m27-il-500000-n5.txt Download as CSV file: xf-m27-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M27 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.1815 0.09863 0.09487 -0.0150 0.5487 0.0158
-8.000 -0.1749 0.09554 0.09178 -0.0169 0.5459 0.0160
-7.500 -0.1643 0.08921 0.08544 -0.0219 0.5407 0.0176
-7.000 -0.1438 0.08434 0.08053 -0.0255 0.5352 0.0188
-6.750 -0.1302 0.08107 0.07724 -0.0291 0.5325 0.0203
-6.250 -0.0990 0.07421 0.07027 -0.0362 0.5267 0.0205
-6.000 -0.0792 0.07054 0.06649 -0.0405 0.5242 0.0206
-5.750 -0.0635 0.06725 0.06314 -0.0418 0.5216 0.0207
-5.500 -0.0482 0.06484 0.06072 -0.0421 0.5190 0.0210
-5.250 -0.0298 0.06257 0.05843 -0.0433 0.5162 0.0213
-5.000 -0.0096 0.06030 0.05611 -0.0447 0.5131 0.0218
-4.750 0.0120 0.05796 0.05369 -0.0463 0.5100 0.0226
-4.500 0.0448 0.05505 0.05059 -0.0500 0.5073 0.0246
-4.250 0.0699 0.05219 0.04760 -0.0516 0.5047 0.0247
-4.000 0.0944 0.04942 0.04473 -0.0528 0.5021 0.0248
-3.750 0.1188 0.04683 0.04203 -0.0537 0.4995 0.0248
-3.500 0.1409 0.04386 0.03897 -0.0543 0.4969 0.0250
-3.250 0.1604 0.04210 0.03716 -0.0544 0.4939 0.0253
-3.000 0.1829 0.04066 0.03566 -0.0548 0.4910 0.0259
-2.750 0.2075 0.03913 0.03405 -0.0553 0.4883 0.0267
-2.500 0.2336 0.03733 0.03215 -0.0557 0.4855 0.0276
-2.250 0.2684 0.03518 0.02976 -0.0562 0.4828 0.0298
-2.000 0.2949 0.03323 0.02767 -0.0563 0.4801 0.0299
-1.500 0.3430 0.02962 0.02382 -0.0563 0.4749 0.0293
-1.250 0.3695 0.02790 0.02197 -0.0562 0.4725 0.0292
-1.000 0.3953 0.02632 0.02028 -0.0562 0.4697 0.0280
-0.750 0.4227 0.02413 0.01789 -0.0557 0.4669 0.0267
-0.500 0.4505 0.02197 0.01547 -0.0552 0.4643 0.0264
-0.250 0.4782 0.02007 0.01334 -0.0547 0.4618 0.0266
0.000 0.5059 0.01745 0.01034 -0.0539 0.4595 0.0267
0.250 0.5343 0.01582 0.00840 -0.0538 0.4572 0.0275
0.500 0.5633 0.01467 0.00702 -0.0538 0.4546 0.0276
0.750 0.5921 0.01403 0.00622 -0.0540 0.4516 0.0278
1.000 0.6207 0.01360 0.00567 -0.0542 0.4487 0.0282
1.250 0.6489 0.01329 0.00527 -0.0544 0.4460 0.0286
1.500 0.6767 0.01308 0.00497 -0.0545 0.4434 0.0290
1.750 0.7045 0.01289 0.00475 -0.0547 0.4411 0.0294
2.000 0.7322 0.01274 0.00459 -0.0548 0.4384 0.0298
2.250 0.7595 0.01260 0.00444 -0.0548 0.4355 0.0301
2.500 0.7860 0.01242 0.00426 -0.0548 0.4326 0.0312
2.750 0.8128 0.01238 0.00422 -0.0548 0.4300 0.0324
3.000 0.8395 0.01237 0.00419 -0.0548 0.4275 0.0332
3.250 0.8666 0.01234 0.00417 -0.0549 0.4253 0.0341
3.500 0.8938 0.01233 0.00418 -0.0550 0.4227 0.0351
3.750 0.9209 0.01235 0.00420 -0.0551 0.4198 0.0362
4.000 0.9477 0.01238 0.00424 -0.0552 0.4169 0.0386
4.250 0.9745 0.01245 0.00431 -0.0553 0.4142 0.0426
4.500 1.0008 0.01252 0.00445 -0.0554 0.4118 0.0719
4.750 1.0279 0.01257 0.00459 -0.0556 0.4097 0.0942
5.000 1.0549 0.01265 0.00472 -0.0558 0.4072 0.1097
5.250 1.0816 0.01272 0.00487 -0.0560 0.4044 0.1412
5.500 1.1075 0.01277 0.00506 -0.0561 0.4016 0.2238
6.000 1.1817 0.01216 0.00586 -0.0609 0.3962 0.9836
6.250 1.2296 0.01239 0.00612 -0.0658 0.3937 0.9869
6.500 1.2784 0.01265 0.00640 -0.0709 0.3908 0.9927
6.750 1.3291 0.01289 0.00664 -0.0766 0.3861 0.9963
7.000 1.3719 0.01314 0.00686 -0.0806 0.3789 0.9991
7.250 1.4002 0.01333 0.00707 -0.0814 0.3735 0.9997
7.500 1.4250 0.01360 0.00733 -0.0815 0.3680 1.0000
7.750 1.4469 0.01385 0.00759 -0.0810 0.3634 1.0000
8.000 1.4685 0.01409 0.00786 -0.0804 0.3576 1.0000
8.250 1.4876 0.01448 0.00821 -0.0795 0.3504 1.0000
8.500 1.5067 0.01482 0.00857 -0.0785 0.3403 1.0000
8.750 1.5236 0.01526 0.00900 -0.0773 0.3312 1.0000
9.000 1.5365 0.01586 0.00955 -0.0755 0.3187 1.0000
9.250 1.5441 0.01664 0.01027 -0.0729 0.3025 1.0000
9.500 1.5426 0.01768 0.01125 -0.0690 0.2847 1.0000
9.750 1.5235 0.01897 0.01252 -0.0624 0.2749 1.0000
10.000 1.5074 0.02079 0.01431 -0.0575 0.2602 1.0000
10.250 1.4919 0.02323 0.01668 -0.0536 0.2411 1.0000
10.500 1.4684 0.02674 0.02009 -0.0501 0.2160 1.0000
10.750 1.4400 0.03113 0.02435 -0.0471 0.1878 1.0000
11.000 1.4117 0.03583 0.02896 -0.0446 0.1622 1.0000
11.250 1.3845 0.04066 0.03371 -0.0425 0.1383 1.0000
11.500 1.3573 0.04579 0.03878 -0.0408 0.1137 1.0000
11.750 1.3310 0.05115 0.04406 -0.0396 0.0873 1.0000
12.000 1.3057 0.05663 0.04948 -0.0386 0.0608 1.0000
12.250 1.2821 0.06214 0.05493 -0.0378 0.0329 1.0000
12.500 1.2702 0.06643 0.05921 -0.0374 0.0201 1.0000
12.750 1.2684 0.06960 0.06243 -0.0372 0.0178 1.0000
13.000 1.2691 0.07247 0.06537 -0.0371 0.0165 1.0000
13.250 1.2695 0.07542 0.06839 -0.0370 0.0157 1.0000
13.500 1.2694 0.07845 0.07148 -0.0369 0.0148 1.0000
13.750 1.2681 0.08169 0.07479 -0.0369 0.0141 1.0000
14.000 1.2666 0.08497 0.07814 -0.0369 0.0134 1.0000
14.250 1.2669 0.08803 0.08128 -0.0370 0.0130 1.0000
14.500 1.2679 0.09102 0.08434 -0.0371 0.0127 1.0000
14.750 1.2681 0.09415 0.08755 -0.0373 0.0123 1.0000
15.000 1.2678 0.09738 0.09085 -0.0375 0.0119 1.0000
15.250 1.2679 0.10055 0.09409 -0.0378 0.0115 1.0000
15.500 1.2663 0.10397 0.09758 -0.0381 0.0112 1.0000
15.750 1.2657 0.10728 0.10095 -0.0385 0.0108 1.0000
16.000 1.2623 0.11103 0.10478 -0.0390 0.0105 1.0000
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Polar data table (+)
Polar graphs
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