NACA M17 AIRFOIL (m17-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M17 AIRFOIL (m17-il) Reynolds number: 100,000 Max Cl/Cd: 55.47 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m17-il-100000-n5.txt Download as CSV file: xf-m17-il-100000-n5.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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4727 0.12798 0.12337 0.0090 1.0000 0.0384
-9.750 -0.4712 0.12534 0.12078 0.0057 1.0000 0.0386
-9.500 -0.4683 0.12222 0.11771 0.0027 1.0000 0.0388
-9.250 -0.4647 0.11884 0.11436 -0.0003 1.0000 0.0389
-9.000 -0.4606 0.11528 0.11085 -0.0033 1.0000 0.0389
-8.750 -0.4565 0.11159 0.10721 -0.0065 1.0000 0.0390
-8.500 -0.4523 0.10786 0.10351 -0.0098 1.0000 0.0391
-8.250 -0.4365 0.10059 0.09626 -0.0033 1.0000 0.0409
-8.000 -0.4275 0.09723 0.09292 -0.0034 1.0000 0.0432
-7.750 -0.4220 0.09395 0.08969 -0.0056 1.0000 0.0456
-7.500 -0.4171 0.09060 0.08639 -0.0092 1.0000 0.0482
-7.250 -0.4072 0.08758 0.08336 -0.0183 1.0000 0.0511
-7.000 -0.3925 0.08414 0.07984 -0.0249 1.0000 0.0516
-6.750 -0.3736 0.08000 0.07559 -0.0299 0.9669 0.0518
-6.500 -0.3604 0.07378 0.06939 -0.0306 0.9338 0.0528
-6.250 -0.3482 0.07044 0.06604 -0.0289 0.9079 0.0556
-6.000 -0.3320 0.06750 0.06296 -0.0306 0.8848 0.0609
-5.750 -0.3011 0.06554 0.06039 -0.0375 0.8647 0.0649
-5.500 -0.2919 0.06023 0.05511 -0.0368 0.8500 0.0658
-5.250 -0.2787 0.05678 0.05160 -0.0361 0.8361 0.0671
-5.000 -0.2628 0.05414 0.04886 -0.0356 0.8228 0.0705
-4.750 -0.2313 0.05236 0.04642 -0.0383 0.8106 0.0794
-4.500 -0.2084 0.04457 0.03820 -0.0383 0.8016 0.0515
-4.250 -0.1892 0.04152 0.03497 -0.0380 0.7911 0.0506
-4.000 -0.1678 0.03879 0.03200 -0.0380 0.7804 0.0538
-3.500 -0.1169 0.03319 0.02545 -0.0371 0.7623 0.0513
-3.250 -0.0924 0.03083 0.02275 -0.0367 0.7530 0.0510
-3.000 -0.0675 0.02876 0.02026 -0.0361 0.7451 0.0509
-2.750 -0.0408 0.02697 0.01804 -0.0357 0.7358 0.0522
-2.500 -0.0160 0.02554 0.01643 -0.0355 0.7279 0.0541
-2.250 0.0112 0.02406 0.01459 -0.0351 0.7194 0.0540
-2.000 0.0387 0.02277 0.01294 -0.0347 0.7119 0.0537
-1.750 0.0662 0.02167 0.01157 -0.0344 0.7040 0.0538
-1.500 0.0938 0.02073 0.01042 -0.0341 0.6966 0.0541
-1.250 0.1213 0.01992 0.00945 -0.0339 0.6891 0.0547
-1.000 0.1487 0.01923 0.00863 -0.0336 0.6820 0.0557
-0.750 0.1760 0.01864 0.00795 -0.0334 0.6747 0.0570
-0.500 0.2031 0.01813 0.00736 -0.0331 0.6681 0.0586
-0.250 0.2301 0.01775 0.00691 -0.0328 0.6609 0.0620
0.000 0.2566 0.01741 0.00651 -0.0324 0.6548 0.0667
0.250 0.2838 0.01712 0.00623 -0.0323 0.6475 0.0709
0.500 0.3111 0.01690 0.00588 -0.0319 0.6419 0.0754
0.750 0.3391 0.01671 0.00569 -0.0320 0.6346 0.0834
1.000 0.3662 0.01650 0.00549 -0.0317 0.6287 0.1096
1.500 0.4690 0.01455 0.00556 -0.0416 0.6150 1.0000
1.750 0.4947 0.01470 0.00559 -0.0412 0.6091 1.0000
2.000 0.5204 0.01487 0.00568 -0.0408 0.6025 1.0000
2.250 0.5458 0.01502 0.00572 -0.0403 0.5975 1.0000
2.500 0.5717 0.01522 0.00592 -0.0400 0.5906 1.0000
2.750 0.5971 0.01539 0.00604 -0.0396 0.5849 1.0000
3.000 0.6227 0.01559 0.00622 -0.0392 0.5791 1.0000
3.250 0.6482 0.01580 0.00644 -0.0389 0.5728 1.0000
3.500 0.6736 0.01598 0.00659 -0.0384 0.5679 1.0000
3.750 0.6991 0.01624 0.00692 -0.0381 0.5608 1.0000
4.000 0.7245 0.01644 0.00713 -0.0377 0.5552 1.0000
4.250 0.7498 0.01669 0.00745 -0.0373 0.5489 1.0000
4.500 0.7751 0.01692 0.00775 -0.0369 0.5423 1.0000
4.750 0.8002 0.01712 0.00799 -0.0364 0.5355 1.0000
5.000 0.8251 0.01726 0.00821 -0.0359 0.5258 1.0000
5.250 0.8498 0.01742 0.00846 -0.0353 0.5150 1.0000
5.500 0.8747 0.01757 0.00868 -0.0347 0.5060 1.0000
5.750 0.8996 0.01779 0.00903 -0.0343 0.4973 1.0000
6.000 0.9245 0.01804 0.00945 -0.0338 0.4890 1.0000
6.250 0.9494 0.01825 0.00977 -0.0333 0.4809 1.0000
6.500 0.9739 0.01854 0.01027 -0.0328 0.4711 1.0000
6.750 0.9981 0.01869 0.01058 -0.0321 0.4571 1.0000
7.000 1.0203 0.01860 0.01054 -0.0310 0.4251 1.0000
7.250 1.0407 0.01876 0.01063 -0.0298 0.3774 1.0000
7.500 1.0581 0.01944 0.01103 -0.0284 0.3014 1.0000
7.750 1.0654 0.02141 0.01224 -0.0265 0.1922 1.0000
8.000 1.0633 0.02452 0.01438 -0.0242 0.0672 1.0000
8.250 1.0696 0.02654 0.01621 -0.0221 0.0438 1.0000
8.500 1.0784 0.02818 0.01795 -0.0202 0.0373 1.0000
8.750 1.0871 0.02970 0.01963 -0.0183 0.0340 1.0000
9.000 1.0941 0.03124 0.02132 -0.0164 0.0309 1.0000
9.250 1.0952 0.03301 0.02319 -0.0141 0.0287 1.0000
9.500 1.0928 0.03518 0.02544 -0.0121 0.0273 1.0000
9.750 1.0956 0.03717 0.02760 -0.0107 0.0264 1.0000
10.000 1.0986 0.03928 0.02985 -0.0094 0.0256 1.0000
10.250 1.1027 0.04139 0.03208 -0.0082 0.0247 1.0000
10.500 1.1084 0.04341 0.03421 -0.0070 0.0236 1.0000
10.750 1.1151 0.04541 0.03629 -0.0059 0.0223 1.0000
11.000 1.1218 0.04745 0.03844 -0.0048 0.0211 1.0000
11.250 1.1298 0.04953 0.04055 -0.0035 0.0200 1.0000
11.500 1.1451 0.05162 0.04268 -0.0014 0.0193 1.0000
11.750 1.1585 0.05392 0.04514 0.0000 0.0189 1.0000
12.000 1.1678 0.05645 0.04793 0.0010 0.0186 1.0000
12.250 1.1732 0.05933 0.05107 0.0019 0.0183 1.0000
12.500 1.1740 0.06250 0.05452 0.0024 0.0180 1.0000
12.750 1.1713 0.06596 0.05826 0.0025 0.0177 1.0000
13.000 1.1656 0.06975 0.06232 0.0023 0.0173 1.0000
13.250 1.1576 0.07388 0.06671 0.0018 0.0170 1.0000
13.500 1.1474 0.07838 0.07146 0.0008 0.0168 1.0000
13.750 1.1352 0.08326 0.07659 -0.0006 0.0166 1.0000
14.000 1.1214 0.08856 0.08212 -0.0025 0.0165 1.0000
14.250 1.1061 0.09429 0.08807 -0.0048 0.0165 1.0000
14.500 1.0897 0.10055 0.09454 -0.0077 0.0165 1.0000
14.750 1.0725 0.10741 0.10160 -0.0113 0.0166 1.0000
15.000 1.0546 0.11495 0.10932 -0.0155 0.0167 1.0000
15.250 1.0361 0.12324 0.11779 -0.0204 0.0169 1.0000
15.500 1.0174 0.13231 0.12700 -0.0259 0.0172 1.0000
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Polar data table (+)
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