NACA M27 AIRFOIL (m27-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
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Airfoil: NACA M27 AIRFOIL (m27-il) Reynolds number: 1,000,000 Max Cl/Cd: 135.9 at α=7.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m27-il-1000000.txt Download as CSV file: xf-m27-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.1823 0.09298 0.08993 -0.0155 0.5507 0.0159
-7.500 -0.1758 0.09016 0.08713 -0.0177 0.5483 0.0163
-7.250 -0.1700 0.08750 0.08447 -0.0195 0.5455 0.0168
-7.000 -0.1580 0.08417 0.08110 -0.0230 0.5428 0.0178
-6.750 -0.1427 0.08039 0.07728 -0.0288 0.5402 0.0181
-6.500 -0.1265 0.07675 0.07357 -0.0331 0.5372 0.0181
-6.250 -0.1133 0.07267 0.06945 -0.0351 0.5345 0.0182
-6.000 -0.0997 0.06979 0.06656 -0.0355 0.5318 0.0184
-5.750 -0.0831 0.06727 0.06402 -0.0367 0.5288 0.0185
-5.500 -0.0648 0.06485 0.06155 -0.0382 0.5257 0.0187
-5.250 -0.0451 0.06244 0.05909 -0.0399 0.5226 0.0190
-5.000 -0.0242 0.06000 0.05658 -0.0416 0.5194 0.0195
-4.750 -0.0020 0.05744 0.05398 -0.0434 0.5171 0.0205
-4.500 0.0330 0.05406 0.05044 -0.0473 0.5146 0.0216
-4.250 0.0576 0.05121 0.04750 -0.0487 0.5117 0.0217
-4.000 0.0819 0.04847 0.04465 -0.0498 0.5087 0.0217
-3.750 0.0986 0.04510 0.04121 -0.0502 0.5056 0.0220
-3.500 0.1203 0.04331 0.03938 -0.0506 0.5028 0.0223
-3.250 0.1437 0.04161 0.03763 -0.0511 0.5005 0.0226
-3.000 0.1684 0.03985 0.03581 -0.0517 0.4977 0.0230
-2.750 0.1941 0.03801 0.03388 -0.0523 0.4948 0.0238
-2.000 0.2826 0.03161 0.02702 -0.0530 0.4863 0.0259
-1.750 0.3039 0.02879 0.02409 -0.0528 0.4841 0.0264
-1.500 0.3290 0.02766 0.02292 -0.0530 0.4814 0.0267
-1.250 0.3551 0.02652 0.02169 -0.0531 0.4785 0.0272
-1.000 0.3818 0.02531 0.02037 -0.0532 0.4754 0.0279
-0.750 0.4140 0.02360 0.01844 -0.0526 0.4722 0.0307
-0.500 0.4390 0.02044 0.01504 -0.0517 0.4704 0.0314
-0.250 0.4656 0.01966 0.01422 -0.0519 0.4679 0.0318
0.000 0.4928 0.01894 0.01343 -0.0520 0.4652 0.0323
0.250 0.5204 0.01812 0.01251 -0.0520 0.4624 0.0332
0.500 0.5506 0.01711 0.01126 -0.0516 0.4595 0.0366
0.750 0.5763 0.01509 0.00902 -0.0512 0.4565 0.0381
1.000 0.6044 0.01464 0.00854 -0.0515 0.4543 0.0390
1.250 0.6329 0.01407 0.00790 -0.0516 0.4517 0.0408
1.500 0.6614 0.01327 0.00692 -0.0516 0.4491 0.0457
1.750 0.6897 0.01294 0.00655 -0.0519 0.4463 0.0471
2.000 0.7187 0.01066 0.00378 -0.0510 0.4436 0.0317
2.250 0.7460 0.01049 0.00361 -0.0510 0.4408 0.0324
2.500 0.7736 0.01032 0.00345 -0.0511 0.4385 0.0330
2.750 0.8012 0.01020 0.00333 -0.0512 0.4359 0.0337
3.000 0.8287 0.01012 0.00325 -0.0513 0.4332 0.0347
3.250 0.8561 0.01010 0.00321 -0.0514 0.4305 0.0357
3.500 0.8835 0.01013 0.00322 -0.0516 0.4274 0.0366
3.750 0.9109 0.01010 0.00319 -0.0517 0.4250 0.0384
4.000 0.9389 0.01008 0.00320 -0.0519 0.4227 0.0415
4.250 0.9667 0.01007 0.00323 -0.0522 0.4201 0.0507
4.500 0.9942 0.01008 0.00334 -0.0524 0.4174 0.0919
4.750 1.0218 0.01016 0.00344 -0.0527 0.4147 0.1101
5.000 1.0488 0.01027 0.00359 -0.0529 0.4114 0.1400
5.250 1.0724 0.00989 0.00380 -0.0526 0.4095 0.5160
5.500 1.0859 0.00898 0.00404 -0.0495 0.4076 0.9699
5.750 1.1319 0.00928 0.00434 -0.0537 0.4046 0.9837
6.000 1.1982 0.00962 0.00465 -0.0627 0.4011 0.9881
6.250 1.2407 0.00992 0.00489 -0.0665 0.3962 0.9913
6.500 1.2844 0.01005 0.00504 -0.0704 0.3922 0.9942
6.750 1.3403 0.01018 0.00516 -0.0771 0.3875 0.9968
7.000 1.3831 0.01041 0.00535 -0.0811 0.3808 0.9990
7.250 1.4184 0.01049 0.00545 -0.0832 0.3770 1.0000
7.500 1.4433 0.01062 0.00560 -0.0831 0.3721 1.0000
7.750 1.4670 0.01087 0.00581 -0.0830 0.3657 1.0000
8.000 1.4918 0.01098 0.00594 -0.0828 0.3604 1.0000
8.250 1.5150 0.01122 0.00616 -0.0826 0.3531 1.0000
8.500 1.5380 0.01145 0.00639 -0.0823 0.3442 1.0000
8.750 1.5589 0.01185 0.00672 -0.0817 0.3291 1.0000
9.000 1.5753 0.01258 0.00729 -0.0806 0.3033 1.0000
9.250 1.5790 0.01409 0.00848 -0.0778 0.2547 1.0000
9.500 1.5577 0.01651 0.01054 -0.0713 0.1981 1.0000
9.750 1.5280 0.01811 0.01210 -0.0630 0.1830 1.0000
10.000 1.4988 0.02026 0.01423 -0.0560 0.1684 1.0000
10.250 1.4685 0.02356 0.01744 -0.0507 0.1438 1.0000
10.500 1.4405 0.02746 0.02124 -0.0468 0.1180 1.0000
10.750 1.4129 0.03187 0.02557 -0.0439 0.0915 1.0000
11.000 1.3855 0.03661 0.03024 -0.0416 0.0642 1.0000
11.250 1.3598 0.04142 0.03500 -0.0397 0.0401 1.0000
11.500 1.3420 0.04575 0.03931 -0.0385 0.0212 1.0000
11.750 1.3395 0.04867 0.04227 -0.0379 0.0182 1.0000
12.000 1.3405 0.05129 0.04495 -0.0374 0.0173 1.0000
12.250 1.3408 0.05397 0.04769 -0.0370 0.0165 1.0000
12.500 1.3402 0.05683 0.05061 -0.0366 0.0159 1.0000
12.750 1.3379 0.05994 0.05378 -0.0363 0.0152 1.0000
13.000 1.3347 0.06315 0.05707 -0.0359 0.0145 1.0000
13.250 1.3346 0.06606 0.06004 -0.0357 0.0141 1.0000
13.500 1.3348 0.06897 0.06300 -0.0356 0.0138 1.0000
13.750 1.3348 0.07192 0.06602 -0.0354 0.0134 1.0000
14.000 1.3330 0.07512 0.06929 -0.0354 0.0131 1.0000
14.250 1.3322 0.07821 0.07244 -0.0353 0.0128 1.0000
14.500 1.3294 0.08163 0.07592 -0.0354 0.0125 1.0000
14.750 1.3276 0.08492 0.07927 -0.0354 0.0122 1.0000
15.000 1.3224 0.08869 0.08311 -0.0356 0.0119 1.0000
15.250 1.3169 0.09253 0.08702 -0.0357 0.0116 1.0000
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Polar data table (+)
Polar graphs
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