NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 200,000 Max Cl/Cd: 60.69 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m11-il-200000-n5.txt Download as CSV file: xf-m11-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M11 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.5381 0.08371 0.08012 -0.0206 1.0000 0.0197
-9.000 -0.5479 0.07794 0.07440 -0.0246 1.0000 0.0199
-8.750 -0.5598 0.07285 0.06934 -0.0280 1.0000 0.0199
-8.250 -0.5867 0.05863 0.05496 -0.0330 1.0000 0.0206
-8.000 -0.5812 0.05714 0.05343 -0.0323 1.0000 0.0213
-7.750 -0.5726 0.05546 0.05171 -0.0316 1.0000 0.0219
-7.500 -0.5668 0.05249 0.04862 -0.0310 1.0000 0.0227
-7.250 -0.5641 0.04823 0.04418 -0.0299 1.0000 0.0234
-7.000 -0.5617 0.04365 0.03934 -0.0282 1.0000 0.0240
-6.750 -0.5583 0.03914 0.03449 -0.0258 1.0000 0.0252
-6.500 -0.5537 0.03478 0.02964 -0.0228 1.0000 0.0270
-6.250 -0.5451 0.03184 0.02621 -0.0198 1.0000 0.0282
-6.000 -0.5374 0.02901 0.02294 -0.0167 1.0000 0.0290
-5.750 -0.5275 0.02672 0.02047 -0.0143 1.0000 0.0303
-5.500 -0.5040 0.02550 0.01911 -0.0142 0.9972 0.0317
-5.250 -0.4724 0.02383 0.01716 -0.0155 0.9912 0.0331
-5.000 -0.4397 0.02206 0.01506 -0.0168 0.9856 0.0343
-4.750 -0.4080 0.02080 0.01351 -0.0178 0.9788 0.0361
-4.500 -0.3746 0.01952 0.01195 -0.0190 0.9730 0.0372
-4.250 -0.3418 0.01836 0.01056 -0.0200 0.9661 0.0378
-4.000 -0.3083 0.01739 0.00940 -0.0212 0.9594 0.0384
-3.750 -0.2750 0.01654 0.00842 -0.0223 0.9518 0.0389
-3.500 -0.2426 0.01588 0.00768 -0.0232 0.9428 0.0397
-3.250 -0.2093 0.01477 0.00652 -0.0244 0.9350 0.0410
-3.000 -0.1789 0.01408 0.00579 -0.0249 0.9240 0.0415
-2.750 -0.1485 0.01350 0.00518 -0.0254 0.9126 0.0421
-2.500 -0.1184 0.01301 0.00465 -0.0258 0.9005 0.0429
-2.250 -0.0892 0.01259 0.00419 -0.0259 0.8874 0.0441
-2.000 -0.0608 0.01226 0.00380 -0.0259 0.8736 0.0455
-1.750 -0.0332 0.01199 0.00346 -0.0256 0.8591 0.0473
-1.500 -0.0062 0.01177 0.00316 -0.0252 0.8444 0.0493
-1.250 0.0202 0.01159 0.00292 -0.0247 0.8294 0.0519
-1.000 0.0460 0.01139 0.00271 -0.0241 0.8143 0.0591
-0.750 0.0704 0.01108 0.00254 -0.0233 0.7993 0.1028
-0.500 0.0916 0.01046 0.00241 -0.0221 0.7842 0.2503
-0.250 0.1832 0.00851 0.00271 -0.0351 0.7722 0.9385
0.250 0.2861 0.00895 0.00288 -0.0445 0.7402 0.9946
0.750 0.3471 0.00898 0.00272 -0.0457 0.7083 1.0000
1.000 0.3701 0.00903 0.00269 -0.0447 0.6930 1.0000
1.250 0.3934 0.00909 0.00267 -0.0437 0.6777 1.0000
1.500 0.4167 0.00915 0.00268 -0.0428 0.6627 1.0000
1.750 0.4402 0.00923 0.00270 -0.0418 0.6472 1.0000
2.000 0.4637 0.00931 0.00273 -0.0409 0.6317 1.0000
2.250 0.4873 0.00941 0.00278 -0.0399 0.6163 1.0000
2.500 0.5108 0.00952 0.00285 -0.0390 0.5998 1.0000
2.750 0.5341 0.00964 0.00293 -0.0380 0.5812 1.0000
3.000 0.5571 0.00979 0.00302 -0.0369 0.5587 1.0000
3.250 0.5799 0.00996 0.00313 -0.0359 0.5348 1.0000
3.500 0.6022 0.01017 0.00324 -0.0347 0.5047 1.0000
3.750 0.6240 0.01042 0.00336 -0.0335 0.4680 1.0000
4.000 0.6459 0.01070 0.00352 -0.0323 0.4322 1.0000
4.250 0.6676 0.01100 0.00372 -0.0311 0.3955 1.0000
4.500 0.6890 0.01136 0.00395 -0.0299 0.3567 1.0000
4.750 0.7101 0.01177 0.00422 -0.0287 0.3166 1.0000
5.000 0.7308 0.01224 0.00456 -0.0275 0.2741 1.0000
5.250 0.7502 0.01286 0.00495 -0.0261 0.2172 1.0000
5.500 0.7682 0.01366 0.00545 -0.0246 0.1570 1.0000
5.750 0.7870 0.01439 0.00599 -0.0233 0.1165 1.0000
6.000 0.8063 0.01508 0.00656 -0.0219 0.0904 1.0000
6.250 0.8260 0.01569 0.00714 -0.0206 0.0727 1.0000
6.500 0.8458 0.01630 0.00773 -0.0193 0.0611 1.0000
6.750 0.8649 0.01695 0.00837 -0.0179 0.0521 1.0000
7.000 0.8851 0.01749 0.00899 -0.0166 0.0452 1.0000
7.250 0.9033 0.01822 0.00974 -0.0151 0.0390 1.0000
7.500 0.9223 0.01886 0.01045 -0.0137 0.0331 1.0000
7.750 0.9394 0.01969 0.01134 -0.0120 0.0289 1.0000
8.000 0.9568 0.02047 0.01221 -0.0103 0.0258 1.0000
8.250 0.9714 0.02150 0.01324 -0.0084 0.0228 1.0000
8.500 0.9884 0.02229 0.01418 -0.0067 0.0208 1.0000
8.750 1.0036 0.02326 0.01526 -0.0048 0.0193 1.0000
9.000 1.0182 0.02429 0.01642 -0.0029 0.0181 1.0000
9.250 1.0317 0.02536 0.01757 -0.0010 0.0169 1.0000
9.500 1.0411 0.02697 0.01926 0.0014 0.0159 1.0000
9.750 1.0543 0.02815 0.02061 0.0034 0.0150 1.0000
10.000 1.0663 0.02946 0.02209 0.0055 0.0142 1.0000
10.250 1.0765 0.03097 0.02378 0.0076 0.0136 1.0000
10.500 1.0850 0.03255 0.02554 0.0100 0.0131 1.0000
10.750 1.0901 0.03413 0.02730 0.0127 0.0127 1.0000
11.000 1.0930 0.03586 0.02921 0.0155 0.0124 1.0000
11.250 1.0948 0.03771 0.03123 0.0180 0.0121 1.0000
11.500 1.0947 0.03965 0.03339 0.0202 0.0118 1.0000
11.750 1.0933 0.04178 0.03568 0.0221 0.0116 1.0000
12.000 1.0898 0.04412 0.03815 0.0235 0.0112 1.0000
12.250 1.0832 0.04728 0.04155 0.0245 0.0112 1.0000
12.500 1.0726 0.05088 0.04534 0.0248 0.0109 1.0000
12.750 1.0581 0.05528 0.04998 0.0244 0.0107 1.0000
13.000 1.0450 0.05987 0.05477 0.0232 0.0108 1.0000
13.250 1.0324 0.06465 0.05978 0.0211 0.0107 1.0000
13.500 1.0144 0.07077 0.06613 0.0181 0.0106 1.0000
13.750 0.9984 0.07698 0.07250 0.0146 0.0107 1.0000
14.000 0.9763 0.08493 0.08063 0.0099 0.0107 1.0000
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Polar data table (+)
Polar graphs
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