NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=200,000 Ncrit=9
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 200,000 Max Cl/Cd: 75.29 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m17-il-200000.txt Download as CSV file: xf-m17-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M17 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3428 0.08771 0.08483 -0.0083 1.0000 0.0369
-8.000 -0.3387 0.08403 0.08119 -0.0096 1.0000 0.0378
-7.750 -0.4468 0.09189 0.08888 -0.0078 1.0000 0.0353
-7.500 -0.4367 0.08820 0.08522 -0.0071 1.0000 0.0362
-7.250 -0.4263 0.08462 0.08166 -0.0091 1.0000 0.0371
-7.000 -0.4143 0.08095 0.07801 -0.0122 1.0000 0.0383
-6.750 -0.3998 0.07714 0.07421 -0.0161 1.0000 0.0398
-6.500 -0.3770 0.07287 0.06989 -0.0225 0.9689 0.0423
-6.250 -0.3368 0.06998 0.06658 -0.0336 0.9262 0.0449
-6.000 -0.3336 0.06393 0.06049 -0.0331 0.9021 0.0459
-5.750 -0.3272 0.06111 0.05762 -0.0307 0.8825 0.0471
-5.500 -0.3143 0.05850 0.05491 -0.0300 0.8650 0.0488
-5.250 -0.2975 0.05574 0.05200 -0.0304 0.8499 0.0512
-5.000 -0.2597 0.05453 0.05018 -0.0338 0.8364 0.0573
-4.750 -0.2484 0.04878 0.04435 -0.0339 0.8249 0.0586
-4.500 -0.2338 0.04616 0.04173 -0.0332 0.8127 0.0604
-4.250 -0.2140 0.04405 0.03950 -0.0330 0.8015 0.0641
-4.000 -0.1820 0.04179 0.03658 -0.0338 0.7924 0.0717
-3.750 -0.1647 0.03853 0.03336 -0.0336 0.7821 0.0735
-3.500 -0.1433 0.03655 0.03131 -0.0333 0.7723 0.0766
-3.000 -0.0937 0.03246 0.02671 -0.0331 0.7547 0.0896
-2.750 -0.0650 0.03163 0.02535 -0.0326 0.7466 0.1010
-2.500 -0.0430 0.02899 0.02275 -0.0325 0.7382 0.1038
-2.250 -0.0155 0.02806 0.02145 -0.0322 0.7302 0.1161
-2.000 0.0083 0.02605 0.01944 -0.0320 0.7225 0.1201
-1.750 0.0346 0.02483 0.01802 -0.0319 0.7146 0.1336
-1.500 0.0602 0.02372 0.01676 -0.0316 0.7073 0.1489
-1.250 0.1004 0.01933 0.01115 -0.0295 0.7015 0.0692
-1.000 0.1284 0.01792 0.00953 -0.0292 0.6942 0.0670
-0.750 0.1566 0.01693 0.00833 -0.0288 0.6874 0.0660
-0.500 0.1848 0.01617 0.00742 -0.0286 0.6800 0.0661
-0.250 0.2125 0.01573 0.00687 -0.0283 0.6738 0.0686
0.000 0.2404 0.01523 0.00632 -0.0281 0.6664 0.0697
0.250 0.2673 0.01479 0.00578 -0.0277 0.6608 0.0706
0.500 0.2941 0.01421 0.00528 -0.0275 0.6531 0.0730
0.750 0.3206 0.01390 0.00497 -0.0271 0.6474 0.0777
1.250 0.3741 0.01333 0.00457 -0.0265 0.6343 0.1243
1.500 0.4587 0.01139 0.00468 -0.0385 0.6267 1.0000
1.750 0.4849 0.01150 0.00470 -0.0381 0.6203 1.0000
2.000 0.5108 0.01164 0.00473 -0.0377 0.6147 1.0000
2.250 0.5370 0.01176 0.00483 -0.0375 0.6078 1.0000
2.500 0.5628 0.01191 0.00487 -0.0370 0.6026 1.0000
2.750 0.5890 0.01206 0.00505 -0.0368 0.5959 1.0000
3.000 0.6149 0.01220 0.00514 -0.0364 0.5900 1.0000
3.250 0.6407 0.01238 0.00529 -0.0361 0.5845 1.0000
3.500 0.6667 0.01253 0.00547 -0.0358 0.5777 1.0000
3.750 0.6924 0.01269 0.00557 -0.0353 0.5726 1.0000
4.000 0.7181 0.01281 0.00576 -0.0350 0.5640 1.0000
4.250 0.7435 0.01287 0.00576 -0.0344 0.5559 1.0000
4.500 0.7691 0.01293 0.00586 -0.0340 0.5466 1.0000
4.750 0.7948 0.01308 0.00605 -0.0336 0.5401 1.0000
5.000 0.8205 0.01323 0.00627 -0.0332 0.5327 1.0000
5.250 0.8462 0.01339 0.00644 -0.0328 0.5263 1.0000
5.500 0.8718 0.01355 0.00673 -0.0325 0.5186 1.0000
5.750 0.8973 0.01368 0.00690 -0.0320 0.5108 1.0000
6.000 0.9223 0.01360 0.00684 -0.0313 0.4969 1.0000
6.250 0.9467 0.01343 0.00672 -0.0305 0.4765 1.0000
6.500 0.9712 0.01339 0.00673 -0.0298 0.4578 1.0000
6.750 0.9955 0.01341 0.00681 -0.0291 0.4335 1.0000
7.000 1.0187 0.01353 0.00690 -0.0283 0.3968 1.0000
7.250 1.0372 0.01428 0.00724 -0.0272 0.2958 1.0000
7.500 1.0330 0.01829 0.00959 -0.0246 0.0654 1.0000
7.750 1.0464 0.01991 0.01114 -0.0230 0.0461 1.0000
8.000 1.0621 0.02112 0.01249 -0.0215 0.0412 1.0000
8.250 1.0763 0.02236 0.01384 -0.0200 0.0381 1.0000
8.500 1.0863 0.02389 0.01544 -0.0181 0.0358 1.0000
8.750 1.0897 0.02595 0.01753 -0.0155 0.0338 1.0000
9.000 1.1031 0.02705 0.01873 -0.0139 0.0323 1.0000
9.250 1.1134 0.02847 0.02024 -0.0119 0.0311 1.0000
9.500 1.1236 0.02994 0.02177 -0.0098 0.0302 1.0000
9.750 1.1368 0.03149 0.02337 -0.0080 0.0294 1.0000
10.000 1.1544 0.03316 0.02510 -0.0066 0.0287 1.0000
10.250 1.1740 0.03496 0.02697 -0.0056 0.0279 1.0000
10.500 1.1967 0.03724 0.02927 -0.0053 0.0265 1.0000
10.750 1.2261 0.04132 0.03352 -0.0057 0.0257 1.0000
11.000 1.2434 0.04410 0.03653 -0.0047 0.0258 1.0000
11.250 1.2558 0.04637 0.03906 -0.0033 0.0260 1.0000
11.500 1.2624 0.04833 0.04133 -0.0011 0.0266 1.0000
11.750 1.2489 0.05188 0.04555 0.0030 0.0289 1.0000
12.000 1.1514 0.04416 0.03809 0.0121 0.0279 1.0000
12.250 1.1306 0.04893 0.04326 0.0141 0.0292 1.0000
12.500 1.1132 0.05401 0.04864 0.0154 0.0304 1.0000
12.750 1.0948 0.05913 0.05399 0.0162 0.0313 1.0000
13.000 1.0750 0.06451 0.05957 0.0166 0.0320 1.0000
13.250 1.0547 0.07039 0.06561 0.0166 0.0328 1.0000
13.500 1.0152 0.07732 0.07304 0.0155 0.0381 1.0000
13.750 0.9869 0.08337 0.07925 0.0139 0.0382 1.0000
14.000 0.9592 0.08941 0.08545 0.0118 0.0381 1.0000
14.250 0.9326 0.09541 0.09157 0.0093 0.0378 1.0000
14.500 0.9055 0.10154 0.09782 0.0064 0.0374 1.0000
14.750 0.8782 0.10712 0.10349 0.0037 0.0370 1.0000
15.000 0.8492 0.11289 0.10934 0.0004 0.0366 1.0000
15.250 0.8163 0.12052 0.11707 -0.0044 0.0364 1.0000
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Polar data table (+)
Polar graphs
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