NACA 64-110 AIRFOIL (n64110-il) Xfoil prediction polar at RE=100,000 Ncrit=5
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Airfoil: NACA 64-110 AIRFOIL (n64110-il) Reynolds number: 100,000 Max Cl/Cd: 39.05 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-n64110-il-100000-n5.txt Download as CSV file: xf-n64110-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA 64-110 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.6292 0.09027 0.08530 -0.0144 1.0000 0.0283
-10.000 -0.6446 0.08066 0.07572 -0.0218 1.0000 0.0273
-9.750 -0.6673 0.07236 0.06738 -0.0282 1.0000 0.0266
-9.500 -0.6919 0.06623 0.06113 -0.0315 1.0000 0.0261
-9.250 -0.7133 0.06148 0.05619 -0.0318 1.0000 0.0258
-9.000 -0.7249 0.05692 0.05137 -0.0317 1.0000 0.0256
-8.750 -0.7251 0.05353 0.04777 -0.0312 1.0000 0.0260
-8.500 -0.7203 0.05066 0.04472 -0.0307 1.0000 0.0268
-8.250 -0.7136 0.04780 0.04158 -0.0301 1.0000 0.0281
-8.000 -0.7062 0.04449 0.03791 -0.0292 1.0000 0.0292
-7.750 -0.6964 0.04095 0.03396 -0.0282 1.0000 0.0295
-7.500 -0.6837 0.03764 0.03020 -0.0271 1.0000 0.0299
-7.250 -0.6679 0.03470 0.02682 -0.0260 1.0000 0.0304
-7.000 -0.6497 0.03210 0.02380 -0.0249 1.0000 0.0311
-6.750 -0.6296 0.03022 0.02150 -0.0238 1.0000 0.0328
-6.500 -0.6089 0.02838 0.01932 -0.0228 1.0000 0.0343
-6.250 -0.5880 0.02618 0.01696 -0.0220 1.0000 0.0355
-6.000 -0.5668 0.02459 0.01528 -0.0210 1.0000 0.0367
-5.750 -0.5460 0.02326 0.01387 -0.0199 1.0000 0.0382
-5.500 -0.5262 0.02209 0.01264 -0.0186 1.0000 0.0399
-5.250 -0.5072 0.02117 0.01164 -0.0172 1.0000 0.0429
-5.000 -0.4889 0.02041 0.01076 -0.0156 1.0000 0.0458
-4.750 -0.4741 0.01933 0.00973 -0.0138 1.0000 0.0488
-4.500 -0.4584 0.01860 0.00898 -0.0120 1.0000 0.0522
-4.250 -0.4252 0.01777 0.00807 -0.0137 0.9913 0.0601
-4.000 -0.3903 0.01697 0.00723 -0.0157 0.9827 0.0727
-3.750 -0.3556 0.01593 0.00638 -0.0179 0.9749 0.1112
-3.500 -0.3304 0.01367 0.00566 -0.0193 0.9654 0.3986
-3.250 -0.3014 0.01319 0.00574 -0.0194 0.9560 0.5646
-3.000 -0.2713 0.01319 0.00598 -0.0192 0.9473 0.6504
-2.750 -0.2450 0.01350 0.00642 -0.0176 0.9375 0.7187
-2.500 -0.2168 0.01364 0.00653 -0.0169 0.9276 0.7443
-2.250 -0.1847 0.01361 0.00637 -0.0175 0.9191 0.7554
-2.000 -0.1547 0.01356 0.00620 -0.0178 0.9091 0.7660
-1.750 -0.1255 0.01350 0.00604 -0.0181 0.8990 0.7760
-1.500 -0.0958 0.01346 0.00592 -0.0182 0.8902 0.7841
-1.250 -0.0678 0.01341 0.00579 -0.0181 0.8801 0.7936
-1.000 -0.0403 0.01338 0.00571 -0.0179 0.8702 0.8028
-0.750 -0.0120 0.01335 0.00563 -0.0178 0.8616 0.8116
-0.500 0.0148 0.01333 0.00557 -0.0175 0.8517 0.8218
-0.250 0.0418 0.01332 0.00554 -0.0172 0.8423 0.8312
0.000 0.0694 0.01330 0.00550 -0.0169 0.8343 0.8407
0.250 0.0953 0.01331 0.00551 -0.0164 0.8242 0.8515
0.500 0.1221 0.01331 0.00553 -0.0160 0.8158 0.8620
0.750 0.1490 0.01332 0.00555 -0.0156 0.8071 0.8722
1.000 0.1759 0.01334 0.00561 -0.0153 0.7981 0.8833
1.250 0.2036 0.01334 0.00565 -0.0151 0.7907 0.8948
1.500 0.2310 0.01339 0.00576 -0.0150 0.7813 0.9072
1.750 0.2614 0.01340 0.00583 -0.0153 0.7737 0.9181
2.000 0.2925 0.01344 0.00596 -0.0160 0.7649 0.9297
2.250 0.3251 0.01349 0.00611 -0.0170 0.7564 0.9414
2.500 0.3591 0.01350 0.00621 -0.0181 0.7474 0.9528
2.750 0.3943 0.01353 0.00638 -0.0196 0.7358 0.9643
3.000 0.4305 0.01352 0.00651 -0.0213 0.7222 0.9753
3.250 0.4637 0.01320 0.00616 -0.0214 0.6812 0.9870
3.500 0.4916 0.01299 0.00567 -0.0206 0.6033 1.0000
3.750 0.5080 0.01313 0.00570 -0.0183 0.5478 1.0000
4.000 0.5252 0.01345 0.00578 -0.0162 0.4667 1.0000
4.250 0.5325 0.01510 0.00615 -0.0133 0.2295 1.0000
4.500 0.5456 0.01679 0.00702 -0.0119 0.1005 1.0000
4.750 0.5652 0.01779 0.00786 -0.0108 0.0746 1.0000
5.000 0.5860 0.01865 0.00868 -0.0100 0.0615 1.0000
5.250 0.6073 0.01948 0.00963 -0.0091 0.0547 1.0000
5.500 0.6274 0.02048 0.01060 -0.0082 0.0489 1.0000
5.750 0.6489 0.02141 0.01162 -0.0073 0.0443 1.0000
6.000 0.6705 0.02251 0.01278 -0.0065 0.0415 1.0000
6.250 0.6926 0.02375 0.01404 -0.0057 0.0393 1.0000
6.500 0.7153 0.02532 0.01564 -0.0051 0.0374 1.0000
6.750 0.7393 0.02712 0.01754 -0.0046 0.0356 1.0000
7.000 0.7635 0.02848 0.01916 -0.0041 0.0333 1.0000
7.250 0.7873 0.03034 0.02128 -0.0035 0.0318 1.0000
7.500 0.8098 0.03250 0.02373 -0.0028 0.0308 1.0000
7.750 0.8304 0.03488 0.02648 -0.0020 0.0301 1.0000
8.000 0.8487 0.03752 0.02949 -0.0011 0.0295 1.0000
8.250 0.8644 0.04026 0.03259 -0.0001 0.0289 1.0000
8.500 0.8787 0.04266 0.03525 0.0008 0.0278 1.0000
8.750 0.8896 0.04549 0.03825 0.0015 0.0266 1.0000
9.000 0.8932 0.04946 0.04258 0.0028 0.0258 1.0000
9.250 0.8946 0.05294 0.04649 0.0044 0.0255 1.0000
9.500 0.8915 0.05670 0.05062 0.0058 0.0254 1.0000
9.750 0.8839 0.06050 0.05471 0.0072 0.0254 1.0000
10.000 0.8704 0.06411 0.05855 0.0087 0.0254 1.0000
10.250 0.8546 0.06792 0.06253 0.0093 0.0255 1.0000
10.500 0.8372 0.07234 0.06710 0.0084 0.0256 1.0000
10.750 0.8183 0.07756 0.07245 0.0061 0.0256 1.0000
11.000 0.8009 0.08351 0.07851 0.0026 0.0258 1.0000
11.250 0.7861 0.09003 0.08508 -0.0015 0.0260 1.0000
11.500 0.7708 0.09801 0.09318 -0.0075 0.0263 1.0000
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