NACA M23 AIRFOIL (m23-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M23 AIRFOIL (m23-il) Reynolds number: 500,000 Max Cl/Cd: 106.55 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m23-il-500000.txt Download as CSV file: xf-m23-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M23 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.4010 0.10645 0.10351 0.0229 0.6651 0.0157
-7.750 -0.3959 0.10343 0.10047 0.0195 0.6609 0.0159
-7.500 -0.3925 0.10050 0.09755 0.0158 0.6561 0.0160
-7.250 -0.3822 0.09714 0.09416 0.0111 0.6515 0.0160
-7.000 -0.3693 0.09363 0.09058 0.0066 0.6474 0.0161
-6.750 -0.3617 0.08865 0.08559 0.0069 0.6428 0.0162
-6.500 -0.3521 0.08514 0.08206 0.0070 0.6377 0.0165
-6.250 -0.3389 0.08202 0.07889 0.0052 0.6333 0.0169
-6.000 -0.3232 0.07882 0.07564 0.0028 0.6289 0.0173
-5.750 -0.3058 0.07553 0.07231 0.0003 0.6242 0.0178
-5.500 -0.2866 0.07219 0.06890 -0.0024 0.6198 0.0187
-5.250 -0.2546 0.06908 0.06564 -0.0075 0.6157 0.0201
-5.000 -0.2243 0.06574 0.06216 -0.0117 0.6112 0.0203
-4.750 -0.1986 0.06228 0.05858 -0.0140 0.6068 0.0204
-4.500 -0.1880 0.05734 0.05359 -0.0142 0.6031 0.0208
-4.250 -0.1701 0.05456 0.05077 -0.0146 0.5987 0.0212
-4.000 -0.1479 0.05197 0.04811 -0.0156 0.5939 0.0218
-3.750 -0.1234 0.04942 0.04544 -0.0168 0.5897 0.0229
-3.500 -0.0950 0.04689 0.04277 -0.0181 0.5857 0.0247
-3.250 -0.0534 0.04512 0.04074 -0.0198 0.5814 0.0260
-3.000 -0.0254 0.04240 0.03784 -0.0204 0.5772 0.0260
-2.750 -0.0085 0.03779 0.03310 -0.0207 0.5736 0.0269
-2.500 0.0145 0.03589 0.03114 -0.0210 0.5693 0.0275
-2.250 0.0399 0.03410 0.02926 -0.0213 0.5648 0.0284
-2.000 0.0670 0.03236 0.02737 -0.0215 0.5609 0.0300
-1.750 0.1031 0.03144 0.02617 -0.0214 0.5570 0.0334
-1.500 0.1329 0.02973 0.02428 -0.0212 0.5527 0.0337
-1.250 0.1554 0.02574 0.02009 -0.0212 0.5488 0.0351
-1.000 0.1811 0.02453 0.01877 -0.0214 0.5450 0.0360
-0.750 0.2081 0.02342 0.01755 -0.0215 0.5411 0.0375
-0.500 0.2367 0.02226 0.01627 -0.0215 0.5368 0.0405
-0.250 0.2694 0.02213 0.01588 -0.0210 0.5326 0.0439
0.000 0.2942 0.01922 0.01274 -0.0210 0.5290 0.0466
0.250 0.3217 0.01851 0.01199 -0.0213 0.5251 0.0490
0.500 0.3526 0.01890 0.01222 -0.0211 0.5207 0.0561
0.750 0.3789 0.01671 0.00988 -0.0212 0.5169 0.0599
1.000 0.4067 0.01626 0.00934 -0.0214 0.5130 0.0639
1.250 0.4354 0.01561 0.00858 -0.0215 0.5091 0.0735
1.500 0.4636 0.01517 0.00813 -0.0217 0.5049 0.0790
1.750 0.4920 0.01461 0.00744 -0.0218 0.5011 0.0885
2.000 0.5205 0.01423 0.00701 -0.0219 0.4972 0.0945
2.250 0.5492 0.01363 0.00635 -0.0220 0.4931 0.1028
2.500 0.5796 0.01155 0.00387 -0.0205 0.4898 0.0430
2.750 0.6070 0.01135 0.00362 -0.0204 0.4860 0.0433
3.000 0.6347 0.01123 0.00352 -0.0205 0.4821 0.0442
3.250 0.6621 0.01100 0.00330 -0.0205 0.4779 0.0454
3.500 0.6898 0.01090 0.00320 -0.0205 0.4740 0.0463
3.750 0.7176 0.01091 0.00316 -0.0206 0.4703 0.0485
4.000 0.7458 0.01089 0.00319 -0.0208 0.4663 0.0539
4.250 0.7735 0.01078 0.00330 -0.0210 0.4620 0.1419
4.500 0.8744 0.00963 0.00387 -0.0375 0.4561 1.0000
4.750 0.9009 0.00972 0.00396 -0.0374 0.4522 1.0000
5.000 0.9272 0.00978 0.00404 -0.0373 0.4473 1.0000
5.250 0.9531 0.00989 0.00408 -0.0372 0.4410 1.0000
5.500 0.9793 0.00992 0.00419 -0.0371 0.4347 1.0000
5.750 1.0052 0.01002 0.00424 -0.0369 0.4271 1.0000
6.000 1.0312 0.01006 0.00433 -0.0368 0.4181 1.0000
6.250 1.0570 0.01020 0.00444 -0.0367 0.4106 1.0000
6.500 1.0828 0.01029 0.00457 -0.0366 0.3997 1.0000
6.750 1.1084 0.01043 0.00470 -0.0365 0.3851 1.0000
7.000 1.1337 0.01064 0.00488 -0.0364 0.3647 1.0000
7.250 1.1579 0.01119 0.00522 -0.0364 0.3200 1.0000
7.500 1.1662 0.01516 0.00784 -0.0370 0.0963 1.0000
7.750 1.1778 0.01726 0.00954 -0.0362 0.0223 1.0000
8.000 1.1974 0.01790 0.01027 -0.0354 0.0201 1.0000
8.250 1.2155 0.01863 0.01110 -0.0345 0.0189 1.0000
8.500 1.2312 0.01952 0.01208 -0.0334 0.0178 1.0000
8.750 1.2434 0.02060 0.01326 -0.0320 0.0169 1.0000
9.000 1.2495 0.02206 0.01483 -0.0301 0.0160 1.0000
9.250 1.2429 0.02424 0.01714 -0.0278 0.0154 1.0000
9.500 1.2315 0.02699 0.02002 -0.0255 0.0150 1.0000
9.750 1.2335 0.02914 0.02225 -0.0247 0.0148 1.0000
10.000 1.2353 0.03146 0.02466 -0.0241 0.0145 1.0000
10.250 1.2357 0.03400 0.02729 -0.0235 0.0143 1.0000
10.500 1.2351 0.03668 0.03005 -0.0230 0.0141 1.0000
10.750 1.2339 0.03942 0.03288 -0.0225 0.0139 1.0000
11.000 1.2323 0.04221 0.03575 -0.0220 0.0137 1.0000
11.250 1.2308 0.04498 0.03860 -0.0214 0.0135 1.0000
11.500 1.2296 0.04767 0.04137 -0.0208 0.0133 1.0000
11.750 1.2293 0.05025 0.04401 -0.0201 0.0131 1.0000
12.000 1.2305 0.05260 0.04643 -0.0192 0.0130 1.0000
12.250 1.2339 0.05478 0.04866 -0.0184 0.0127 1.0000
12.500 1.2388 0.05680 0.05072 -0.0177 0.0123 1.0000
12.750 1.2445 0.05870 0.05263 -0.0169 0.0119 1.0000
13.000 1.2514 0.06037 0.05432 -0.0159 0.0115 1.0000
13.250 1.2615 0.06153 0.05550 -0.0143 0.0114 1.0000
13.500 1.2745 0.06232 0.05633 -0.0122 0.0112 1.0000
13.750 1.2894 0.06307 0.05714 -0.0101 0.0112 1.0000
14.000 1.3038 0.06422 0.05841 -0.0082 0.0114 1.0000
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Polar data table (+)
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