NACA M14 AIRFOIL (m14-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: NACA M14 AIRFOIL (m14-il) Reynolds number: 200,000 Max Cl/Cd: 76.76 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m14-il-200000.txt Download as CSV file: xf-m14-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M14 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4039 0.09415 0.09076 -0.0149 1.0000 0.0371
-7.750 -0.4001 0.09089 0.08754 -0.0170 1.0000 0.0382
-7.500 -0.3977 0.08765 0.08433 -0.0197 1.0000 0.0395
-7.250 -0.3895 0.08408 0.08078 -0.0279 1.0000 0.0413
-7.000 -0.3740 0.08063 0.07724 -0.0375 1.0000 0.0421
-6.750 -0.3705 0.07480 0.07143 -0.0385 1.0000 0.0428
-6.500 -0.3653 0.07163 0.06834 -0.0356 1.0000 0.0439
-6.250 -0.3579 0.06899 0.06573 -0.0347 1.0000 0.0451
-6.000 -0.3533 0.06650 0.06325 -0.0342 1.0000 0.0463
-5.750 -0.3579 0.06472 0.06151 -0.0321 1.0000 0.0474
-5.500 -0.3033 0.06054 0.05685 -0.0450 0.9926 0.0537
-5.250 -0.2785 0.05364 0.04993 -0.0495 0.9862 0.0552
-5.000 -0.2498 0.05057 0.04690 -0.0521 0.9804 0.0580
-4.500 -0.1715 0.04232 0.03808 -0.0626 0.9630 0.0690
-4.250 -0.1400 0.03965 0.03538 -0.0651 0.9541 0.0724
-4.000 -0.1026 0.03675 0.03194 -0.0681 0.9418 0.0824
-3.750 -0.0793 0.03428 0.02956 -0.0686 0.9280 0.0855
-3.500 -0.0496 0.03274 0.02751 -0.0689 0.9134 0.0969
-3.250 -0.0291 0.03026 0.02507 -0.0685 0.8988 0.0996
-3.000 -0.0033 0.02918 0.02357 -0.0679 0.8844 0.1119
-2.750 0.0180 0.02706 0.02149 -0.0673 0.8704 0.1153
-2.500 0.0427 0.02579 0.01994 -0.0667 0.8569 0.1282
-2.250 0.0669 0.02467 0.01862 -0.0661 0.8439 0.1427
-2.000 0.0908 0.02346 0.01730 -0.0655 0.8314 0.1581
-1.750 0.1314 0.01855 0.01091 -0.0631 0.8227 0.0677
-1.500 0.1581 0.01725 0.00945 -0.0626 0.8103 0.0653
-1.250 0.1851 0.01614 0.00815 -0.0620 0.7985 0.0640
-1.000 0.2120 0.01533 0.00716 -0.0614 0.7872 0.0636
-0.750 0.2388 0.01477 0.00645 -0.0608 0.7766 0.0645
-0.500 0.2658 0.01439 0.00597 -0.0603 0.7652 0.0662
-0.250 0.2926 0.01390 0.00543 -0.0599 0.7539 0.0666
0.000 0.3190 0.01339 0.00490 -0.0594 0.7435 0.0675
0.250 0.3451 0.01294 0.00442 -0.0589 0.7340 0.0699
0.500 0.3720 0.01269 0.00418 -0.0586 0.7229 0.0741
0.750 0.3991 0.01255 0.00398 -0.0583 0.7128 0.0811
1.000 0.4253 0.01210 0.00381 -0.0579 0.7040 0.1598
1.250 0.4722 0.01008 0.00376 -0.0617 0.6931 1.0000
1.500 0.4983 0.01020 0.00376 -0.0613 0.6832 1.0000
1.750 0.5243 0.01034 0.00374 -0.0608 0.6744 1.0000
2.000 0.5505 0.01046 0.00380 -0.0604 0.6639 1.0000
2.250 0.5767 0.01059 0.00386 -0.0600 0.6542 1.0000
2.500 0.6029 0.01074 0.00391 -0.0596 0.6455 1.0000
2.750 0.6293 0.01087 0.00403 -0.0593 0.6350 1.0000
3.000 0.6556 0.01102 0.00413 -0.0590 0.6255 1.0000
3.250 0.6820 0.01118 0.00422 -0.0586 0.6165 1.0000
3.500 0.7085 0.01133 0.00440 -0.0584 0.6060 1.0000
3.750 0.7349 0.01151 0.00455 -0.0581 0.5963 1.0000
4.000 0.7613 0.01168 0.00468 -0.0577 0.5871 1.0000
4.250 0.7877 0.01186 0.00491 -0.0575 0.5763 1.0000
4.500 0.8141 0.01205 0.00512 -0.0572 0.5662 1.0000
4.750 0.8406 0.01226 0.00530 -0.0569 0.5567 1.0000
5.000 0.8668 0.01246 0.00555 -0.0567 0.5457 1.0000
5.250 0.8929 0.01265 0.00582 -0.0563 0.5340 1.0000
5.500 0.9185 0.01279 0.00597 -0.0559 0.5192 1.0000
5.750 0.9433 0.01286 0.00600 -0.0552 0.5001 1.0000
6.000 0.9673 0.01288 0.00603 -0.0543 0.4734 1.0000
6.250 0.9913 0.01301 0.00617 -0.0536 0.4465 1.0000
6.500 1.0155 0.01323 0.00640 -0.0530 0.4215 1.0000
6.750 1.0381 0.01355 0.00665 -0.0521 0.3839 1.0000
7.000 1.0579 0.01419 0.00706 -0.0509 0.3182 1.0000
7.250 1.0540 0.01773 0.00899 -0.0474 0.0760 1.0000
7.500 1.0668 0.01946 0.01059 -0.0453 0.0497 1.0000
7.750 1.0800 0.02095 0.01214 -0.0434 0.0430 1.0000
8.000 1.0961 0.02205 0.01337 -0.0417 0.0398 1.0000
8.250 1.1099 0.02331 0.01471 -0.0398 0.0373 1.0000
8.500 1.1206 0.02484 0.01625 -0.0377 0.0350 1.0000
8.750 1.1296 0.02674 0.01817 -0.0352 0.0330 1.0000
9.000 1.1454 0.02795 0.01951 -0.0336 0.0317 1.0000
9.250 1.1616 0.02946 0.02109 -0.0320 0.0307 1.0000
9.500 1.1801 0.03113 0.02284 -0.0307 0.0298 1.0000
9.750 1.2008 0.03301 0.02483 -0.0297 0.0291 1.0000
10.000 1.2217 0.03498 0.02692 -0.0289 0.0283 1.0000
10.250 1.2397 0.03686 0.02891 -0.0279 0.0271 1.0000
10.500 1.2611 0.03992 0.03205 -0.0277 0.0260 1.0000
10.750 1.2790 0.04291 0.03530 -0.0266 0.0259 1.0000
11.000 1.2916 0.04554 0.03823 -0.0249 0.0262 1.0000
11.250 1.2939 0.04887 0.04216 -0.0217 0.0276 1.0000
11.500 1.2851 0.05423 0.04813 -0.0180 0.0302 1.0000
11.750 1.2775 0.05911 0.05330 -0.0154 0.0319 1.0000
12.000 1.1660 0.05133 0.04585 -0.0027 0.0303 1.0000
12.250 1.1475 0.05598 0.05074 -0.0007 0.0311 1.0000
12.500 1.1280 0.06099 0.05596 0.0007 0.0319 1.0000
12.750 1.0850 0.06850 0.06404 0.0020 0.0385 1.0000
13.000 1.0556 0.07449 0.07024 0.0017 0.0394 1.0000
13.250 1.0274 0.08030 0.07621 0.0007 0.0395 1.0000
13.500 0.9983 0.08623 0.08230 -0.0009 0.0393 1.0000
13.750 0.9692 0.09220 0.08841 -0.0032 0.0388 1.0000
14.000 0.9404 0.09835 0.09469 -0.0061 0.0382 1.0000
14.250 0.9118 0.10436 0.10080 -0.0092 0.0375 1.0000
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Polar data table (+)
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