NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 500,000 Max Cl/Cd: 88.64 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m16-il-500000-n5.txt Download as CSV file: xf-m16-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M16 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5102 0.09806 0.09578 0.0225 0.8176 0.0053
-7.500 -0.5073 0.09448 0.09213 0.0201 0.7983 0.0053
-7.250 -0.4995 0.09043 0.08801 0.0164 0.7818 0.0054
-7.000 -0.4886 0.08650 0.08401 0.0128 0.7669 0.0057
-6.750 -0.4758 0.08217 0.07961 0.0087 0.7536 0.0062
-6.250 -0.4443 0.07246 0.06972 -0.0001 0.7304 0.0065
-6.000 -0.4253 0.06734 0.06451 -0.0044 0.7203 0.0065
-5.750 -0.4041 0.06177 0.05880 -0.0087 0.7111 0.0067
-5.500 -0.3803 0.05536 0.05223 -0.0131 0.7025 0.0070
-5.250 -0.3558 0.05097 0.04766 -0.0160 0.6941 0.0075
-5.000 -0.3307 0.04821 0.04476 -0.0178 0.6852 0.0081
-4.750 -0.3039 0.04409 0.04045 -0.0200 0.6777 0.0093
-4.500 -0.2742 0.03599 0.03194 -0.0225 0.6717 0.0106
-4.250 -0.2472 0.03342 0.02918 -0.0234 0.6642 0.0115
-4.000 -0.2164 0.02110 0.01587 -0.0240 0.6601 0.0171
-3.750 -0.1893 0.02089 0.01554 -0.0242 0.6528 0.0180
-3.500 -0.1621 0.02184 0.01650 -0.0245 0.6451 0.0186
-3.250 -0.1352 0.02242 0.01706 -0.0248 0.6383 0.0198
-3.000 -0.1076 0.02126 0.01570 -0.0249 0.6317 0.0217
-2.750 -0.0793 0.01930 0.01341 -0.0249 0.6259 0.0240
-2.500 -0.0509 0.01752 0.01131 -0.0248 0.6199 0.0256
-2.250 -0.0226 0.01634 0.00984 -0.0248 0.6140 0.0266
-2.000 0.0056 0.01549 0.00877 -0.0248 0.6081 0.0273
-1.750 0.0339 0.01472 0.00780 -0.0248 0.6021 0.0277
-1.500 0.0622 0.01435 0.00729 -0.0249 0.5964 0.0281
-1.250 0.0901 0.01265 0.00532 -0.0250 0.5906 0.0304
-1.000 0.1182 0.01222 0.00481 -0.0251 0.5848 0.0317
-0.750 0.1463 0.01185 0.00438 -0.0252 0.5788 0.0326
-0.500 0.1745 0.01151 0.00397 -0.0252 0.5726 0.0334
-0.250 0.2026 0.01119 0.00358 -0.0253 0.5667 0.0340
0.000 0.2307 0.01088 0.00324 -0.0254 0.5604 0.0344
0.250 0.2587 0.01068 0.00299 -0.0255 0.5543 0.0353
0.500 0.2869 0.01044 0.00273 -0.0256 0.5476 0.0360
1.000 0.3431 0.01000 0.00224 -0.0257 0.5345 0.0354
1.250 0.3713 0.00985 0.00205 -0.0258 0.5277 0.0351
1.500 0.3996 0.00972 0.00191 -0.0260 0.5207 0.0349
1.750 0.4278 0.00964 0.00181 -0.0261 0.5134 0.0348
2.000 0.4562 0.00957 0.00175 -0.0263 0.5061 0.0348
2.500 0.5129 0.00953 0.00168 -0.0266 0.4908 0.0351
2.750 0.5412 0.00954 0.00168 -0.0268 0.4834 0.0358
3.000 0.5695 0.00955 0.00172 -0.0270 0.4752 0.0370
3.250 0.5978 0.00959 0.00176 -0.0272 0.4677 0.0399
3.500 0.6248 0.00759 0.00193 -0.0274 0.4599 1.0000
3.750 0.6525 0.00770 0.00204 -0.0274 0.4522 1.0000
4.000 0.6801 0.00784 0.00215 -0.0275 0.4376 1.0000
4.250 0.7076 0.00806 0.00227 -0.0277 0.4139 1.0000
4.500 0.7350 0.00830 0.00243 -0.0278 0.3857 1.0000
4.750 0.7623 0.00860 0.00264 -0.0280 0.3533 1.0000
5.000 0.7882 0.00948 0.00305 -0.0284 0.2561 1.0000
5.250 0.8098 0.01202 0.00444 -0.0293 0.0210 1.0000
5.500 0.8365 0.01249 0.00495 -0.0293 0.0137 1.0000
5.750 0.8632 0.01288 0.00544 -0.0293 0.0122 1.0000
6.000 0.8896 0.01335 0.00599 -0.0294 0.0106 1.0000
6.250 0.9154 0.01389 0.00657 -0.0294 0.0088 1.0000
6.500 0.9398 0.01486 0.00767 -0.0293 0.0077 1.0000
6.750 0.9646 0.01556 0.00848 -0.0292 0.0072 1.0000
7.000 0.9883 0.01643 0.00946 -0.0290 0.0067 1.0000
7.250 1.0114 0.01733 0.01045 -0.0287 0.0061 1.0000
7.500 1.0350 0.01797 0.01112 -0.0286 0.0055 1.0000
7.750 1.0554 0.01914 0.01234 -0.0281 0.0050 1.0000
8.000 1.0751 0.02040 0.01371 -0.0274 0.0047 1.0000
8.250 1.0943 0.02168 0.01516 -0.0266 0.0045 1.0000
8.500 1.1122 0.02318 0.01679 -0.0256 0.0043 1.0000
8.750 1.1295 0.02486 0.01862 -0.0244 0.0040 1.0000
9.000 1.1465 0.02673 0.02065 -0.0232 0.0038 1.0000
9.250 1.1631 0.02884 0.02295 -0.0219 0.0037 1.0000
9.500 1.1786 0.03134 0.02568 -0.0206 0.0036 1.0000
9.750 1.1919 0.03421 0.02882 -0.0192 0.0036 1.0000
10.000 1.2037 0.03645 0.03127 -0.0180 0.0034 1.0000
10.250 1.2135 0.03830 0.03330 -0.0168 0.0033 1.0000
10.500 1.2201 0.04002 0.03518 -0.0156 0.0032 1.0000
10.750 1.2191 0.04213 0.03746 -0.0136 0.0031 1.0000
11.000 1.2092 0.04542 0.04098 -0.0118 0.0030 1.0000
11.250 1.1997 0.04951 0.04536 -0.0107 0.0029 1.0000
11.500 1.1890 0.05394 0.05006 -0.0103 0.0029 1.0000
11.750 1.1762 0.05862 0.05497 -0.0108 0.0028 1.0000
12.000 1.1618 0.06373 0.06028 -0.0119 0.0028 1.0000
12.250 1.1462 0.06924 0.06599 -0.0138 0.0028 1.0000
12.500 1.1301 0.07503 0.07195 -0.0162 0.0028 1.0000
12.750 1.1133 0.08127 0.07834 -0.0192 0.0028 1.0000
13.000 1.0957 0.08802 0.08524 -0.0227 0.0028 1.0000
13.250 1.0782 0.09517 0.09253 -0.0267 0.0029 1.0000
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Polar data table (+)
Polar graphs
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