NACA M9 AIRFOIL (m9-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M9 AIRFOIL (m9-il) Reynolds number: 500,000 Max Cl/Cd: 80.37 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m9-il-500000.txt Download as CSV file: xf-m9-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M9 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.2421 0.08898 0.08592 -0.0073 0.6866 0.0276
-7.500 -0.2353 0.08654 0.08350 -0.0081 0.6784 0.0283
-7.250 -0.2331 0.08399 0.08093 -0.0090 0.6711 0.0299
-6.500 -0.2720 0.02335 0.01877 -0.0598 0.6695 0.0258
-6.250 -0.2472 0.02000 0.01501 -0.0610 0.6608 0.0270
-6.000 -0.2200 0.01961 0.01457 -0.0608 0.6518 0.0280
-5.750 -0.1932 0.01847 0.01322 -0.0607 0.6413 0.0294
-5.500 -0.1665 0.01725 0.01167 -0.0605 0.6313 0.0307
-5.250 -0.1409 0.01603 0.01023 -0.0601 0.6204 0.0323
-5.000 -0.1133 0.01570 0.00985 -0.0598 0.6087 0.0337
-4.750 -0.0856 0.01529 0.00927 -0.0594 0.5970 0.0357
-4.500 -0.0589 0.01449 0.00823 -0.0590 0.5854 0.0377
-4.250 -0.0313 0.01419 0.00789 -0.0586 0.5732 0.0394
-4.000 -0.0036 0.01389 0.00746 -0.0582 0.5617 0.0415
-3.750 0.0244 0.01369 0.00707 -0.0578 0.5507 0.0434
-3.500 0.0513 0.01307 0.00638 -0.0573 0.5402 0.0460
-3.250 0.0792 0.01287 0.00610 -0.0570 0.5310 0.0483
-3.000 0.1073 0.01272 0.00582 -0.0566 0.5222 0.0506
-2.750 0.1347 0.01221 0.00521 -0.0561 0.5141 0.0533
-2.500 0.1626 0.01206 0.00501 -0.0558 0.5055 0.0560
-2.250 0.1907 0.01192 0.00478 -0.0554 0.4976 0.0589
-2.000 0.2185 0.01160 0.00438 -0.0550 0.4900 0.0619
-1.750 0.2464 0.01149 0.00422 -0.0546 0.4835 0.0657
-1.500 0.2747 0.01136 0.00407 -0.0543 0.4769 0.0695
-1.250 0.3026 0.01116 0.00380 -0.0539 0.4708 0.0740
-1.000 0.3307 0.01111 0.00373 -0.0536 0.4651 0.0793
-0.750 0.3589 0.01099 0.00358 -0.0533 0.4591 0.0846
-0.500 0.3868 0.01090 0.00345 -0.0529 0.4534 0.0915
-0.250 0.4149 0.01085 0.00336 -0.0526 0.4486 0.0980
0.000 0.4429 0.01069 0.00324 -0.0523 0.4440 0.1074
0.250 0.4707 0.01061 0.00313 -0.0520 0.4394 0.1163
0.500 0.4983 0.01059 0.00307 -0.0516 0.4343 0.1266
0.750 0.5263 0.01048 0.00305 -0.0513 0.4291 0.1430
1.000 0.5555 0.01110 0.00386 -0.0511 0.4239 0.2091
1.250 0.5837 0.01165 0.00430 -0.0507 0.4188 0.2231
1.500 0.6119 0.01210 0.00471 -0.0503 0.4145 0.2326
1.750 0.6398 0.01234 0.00503 -0.0500 0.4099 0.2378
2.000 0.6678 0.01268 0.00528 -0.0496 0.4053 0.2451
2.250 0.6948 0.01276 0.00540 -0.0493 0.4008 0.2510
2.500 0.7231 0.01296 0.00558 -0.0491 0.3964 0.2581
2.750 0.7503 0.01283 0.00553 -0.0489 0.3913 0.2633
3.000 0.7778 0.01301 0.00563 -0.0486 0.3859 0.2685
3.250 0.8054 0.01290 0.00554 -0.0484 0.3806 0.2722
3.500 0.8327 0.01287 0.00555 -0.0482 0.3747 0.2769
3.750 0.8600 0.01297 0.00559 -0.0479 0.3688 0.2800
4.000 0.8879 0.01304 0.00568 -0.0477 0.3627 0.2820
4.250 0.9149 0.01291 0.00554 -0.0475 0.3555 0.2855
4.500 0.9421 0.01289 0.00555 -0.0473 0.3481 0.2880
4.750 0.9692 0.01292 0.00557 -0.0471 0.3388 0.2894
5.000 0.9963 0.01298 0.00563 -0.0468 0.3283 0.2913
5.250 1.0230 0.01308 0.00569 -0.0466 0.3151 0.2920
5.500 1.0492 0.01327 0.00582 -0.0463 0.2975 0.2938
5.750 1.0746 0.01355 0.00598 -0.0460 0.2756 0.2942
6.000 1.0995 0.01384 0.00617 -0.0457 0.2565 0.2931
6.250 1.1242 0.01416 0.00641 -0.0453 0.2431 0.2920
6.500 1.1491 0.01444 0.00665 -0.0449 0.2344 0.2910
6.750 1.1737 0.01476 0.00694 -0.0445 0.2276 0.2909
7.000 1.1988 0.01498 0.00719 -0.0441 0.2230 0.2902
7.250 1.2232 0.01528 0.00749 -0.0437 0.2187 0.2892
7.500 1.2465 0.01567 0.00786 -0.0431 0.2144 0.2883
7.750 1.2713 0.01590 0.00814 -0.0427 0.2118 0.2875
8.000 1.2958 0.01613 0.00844 -0.0423 0.2091 0.2869
8.250 1.3197 0.01642 0.00876 -0.0418 0.2062 0.2863
8.500 1.3427 0.01678 0.00913 -0.0412 0.2032 0.2859
8.750 1.3644 0.01723 0.00959 -0.0405 0.2000 0.2855
9.000 1.3866 0.01762 0.01003 -0.0398 0.1974 0.2852
9.250 1.4102 0.01783 0.01034 -0.0393 0.1951 0.2850
9.500 1.4332 0.01808 0.01065 -0.0388 0.1918 0.2848
9.750 1.4545 0.01843 0.01103 -0.0380 0.1878 0.2847
10.000 1.4732 0.01897 0.01158 -0.0370 0.1839 0.2845
10.250 1.4943 0.01929 0.01199 -0.0363 0.1815 0.2845
10.500 1.5157 0.01955 0.01234 -0.0356 0.1791 0.2846
10.750 1.5358 0.01987 0.01275 -0.0347 0.1763 0.2846
11.000 1.5538 0.02029 0.01322 -0.0336 0.1732 0.2847
11.250 1.5678 0.02090 0.01384 -0.0321 0.1694 0.2849
11.500 1.5823 0.02130 0.01434 -0.0305 0.1664 0.2852
11.750 1.5956 0.02173 0.01488 -0.0288 0.1636 0.2854
12.000 1.6059 0.02238 0.01560 -0.0270 0.1600 0.2857
12.250 1.6132 0.02333 0.01658 -0.0253 0.1566 0.2864
12.500 1.6218 0.02434 0.01766 -0.0240 0.1527 0.2870
12.750 1.6334 0.02525 0.01865 -0.0231 0.1470 0.2877
13.000 1.6382 0.02676 0.02019 -0.0222 0.1417 0.2882
13.250 1.6468 0.02806 0.02157 -0.0214 0.1355 0.2888
13.500 1.6483 0.03005 0.02355 -0.0207 0.1271 0.2887
13.750 1.6459 0.03245 0.02594 -0.0201 0.1145 0.2893
14.000 1.6316 0.03611 0.02951 -0.0196 0.0956 0.2892
14.250 1.6069 0.04106 0.03437 -0.0195 0.0786 0.2889
14.500 1.5906 0.04540 0.03873 -0.0197 0.0707 0.2892
14.750 1.5781 0.04952 0.04289 -0.0201 0.0659 0.2894
15.000 1.5645 0.05392 0.04735 -0.0209 0.0618 0.2895
15.250 1.5546 0.05812 0.05163 -0.0217 0.0580 0.2896
15.500 1.5444 0.06251 0.05609 -0.0228 0.0554 0.2897
15.750 1.5300 0.06763 0.06128 -0.0243 0.0519 0.2901
16.000 1.5229 0.07185 0.06558 -0.0255 0.0477 0.2903
16.250 1.5134 0.07653 0.07036 -0.0270 0.0454 0.2905
16.500 1.5017 0.08166 0.07555 -0.0288 0.0409 0.2909
16.750 1.4820 0.08821 0.08213 -0.0314 0.0311 0.2911
17.000 1.4621 0.09502 0.08897 -0.0341 0.0243 0.2907
|
Polar data table (+)
Polar graphs
<< Back to NACA M9 AIRFOIL (m9-il)