NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 1,000,000 Max Cl/Cd: 102.95 at α=4.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m16-il-1000000-n5.txt Download as CSV file: xf-m16-il-1000000-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 = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5403 0.10608 0.10402 0.0297 0.7567 0.0028
-8.250 -0.5382 0.10184 0.09974 0.0276 0.7429 0.0028
-8.000 -0.5364 0.09780 0.09567 0.0251 0.7299 0.0029
-7.750 -0.5334 0.09434 0.09219 0.0226 0.7175 0.0030
-7.500 -0.5244 0.09055 0.08835 0.0188 0.7062 0.0031
-7.250 -0.5136 0.08649 0.08425 0.0148 0.6961 0.0032
-7.000 -0.5007 0.08232 0.08003 0.0108 0.6863 0.0033
-6.750 -0.4862 0.07772 0.07537 0.0065 0.6781 0.0034
-6.250 -0.4513 0.06665 0.06414 -0.0032 0.6638 0.0041
-6.000 -0.4302 0.05571 0.05302 -0.0111 0.6590 0.0046
-5.750 -0.4059 0.04976 0.04690 -0.0149 0.6531 0.0051
-5.500 -0.3810 0.04574 0.04273 -0.0172 0.6463 0.0055
-5.000 -0.3284 0.03104 0.02729 -0.0219 0.6361 0.0071
-4.750 -0.3018 0.02805 0.02412 -0.0224 0.6298 0.0084
-4.500 -0.2751 0.02533 0.02119 -0.0228 0.6243 0.0097
-4.250 -0.2485 0.01972 0.01508 -0.0229 0.6195 0.0118
-4.000 -0.2242 0.01234 0.00662 -0.0227 0.6149 0.0170
-3.750 -0.1962 0.01208 0.00627 -0.0228 0.6094 0.0175
-3.500 -0.1676 0.01249 0.00671 -0.0229 0.6032 0.0184
-3.250 -0.1392 0.01289 0.00714 -0.0231 0.5976 0.0191
-3.000 -0.1110 0.01312 0.00737 -0.0232 0.5920 0.0199
-2.750 -0.0828 0.01288 0.00703 -0.0233 0.5861 0.0207
-2.500 -0.0547 0.01251 0.00657 -0.0234 0.5810 0.0217
-2.250 -0.0265 0.01193 0.00586 -0.0235 0.5754 0.0228
-1.750 0.0300 0.01096 0.00466 -0.0236 0.5645 0.0253
-1.500 0.0584 0.01079 0.00443 -0.0238 0.5585 0.0263
-1.250 0.0867 0.01072 0.00431 -0.0239 0.5529 0.0271
-1.000 0.1151 0.01067 0.00422 -0.0240 0.5466 0.0276
-0.500 0.1717 0.01018 0.00362 -0.0243 0.5344 0.0289
-0.250 0.1996 0.00961 0.00296 -0.0244 0.5281 0.0298
0.000 0.2278 0.00930 0.00262 -0.0245 0.5217 0.0302
0.250 0.2560 0.00908 0.00236 -0.0246 0.5147 0.0303
0.500 0.2843 0.00888 0.00214 -0.0248 0.5082 0.0303
0.750 0.3125 0.00870 0.00194 -0.0249 0.5006 0.0312
1.000 0.3409 0.00857 0.00178 -0.0251 0.4936 0.0312
1.250 0.3693 0.00847 0.00165 -0.0252 0.4861 0.0308
1.500 0.3977 0.00839 0.00156 -0.0254 0.4788 0.0305
1.750 0.4262 0.00834 0.00148 -0.0256 0.4709 0.0302
2.000 0.4546 0.00829 0.00143 -0.0258 0.4638 0.0300
2.250 0.4831 0.00827 0.00139 -0.0260 0.4558 0.0298
2.500 0.5115 0.00825 0.00137 -0.0262 0.4485 0.0296
2.750 0.5400 0.00825 0.00136 -0.0264 0.4410 0.0296
3.000 0.5684 0.00825 0.00137 -0.0267 0.4334 0.0296
3.250 0.5969 0.00829 0.00139 -0.0269 0.4263 0.0298
3.500 0.6253 0.00833 0.00142 -0.0271 0.4168 0.0306
3.750 0.6535 0.00846 0.00150 -0.0274 0.3938 0.0363
4.250 0.7052 0.00685 0.00187 -0.0274 0.3469 1.0000
4.500 0.7327 0.00728 0.00208 -0.0277 0.2995 1.0000
4.750 0.7593 0.00813 0.00251 -0.0282 0.2060 1.0000
5.000 0.7837 0.01004 0.00354 -0.0290 0.0163 1.0000
5.250 0.8112 0.01028 0.00380 -0.0291 0.0125 1.0000
5.500 0.8385 0.01055 0.00411 -0.0293 0.0104 1.0000
5.750 0.8656 0.01096 0.00454 -0.0294 0.0078 1.0000
6.000 0.8927 0.01124 0.00485 -0.0295 0.0071 1.0000
6.250 0.9196 0.01159 0.00523 -0.0297 0.0063 1.0000
6.500 0.9463 0.01197 0.00563 -0.0298 0.0056 1.0000
6.750 0.9721 0.01261 0.00636 -0.0298 0.0049 1.0000
7.000 0.9985 0.01296 0.00675 -0.0299 0.0045 1.0000
7.250 1.0243 0.01345 0.00728 -0.0300 0.0041 1.0000
7.500 1.0498 0.01397 0.00786 -0.0300 0.0037 1.0000
7.750 1.0748 0.01453 0.00846 -0.0300 0.0035 1.0000
8.000 1.0992 0.01517 0.00915 -0.0299 0.0032 1.0000
8.250 1.1208 0.01632 0.01044 -0.0295 0.0030 1.0000
8.500 1.1436 0.01710 0.01131 -0.0293 0.0029 1.0000
8.750 1.1652 0.01805 0.01236 -0.0289 0.0027 1.0000
9.000 1.1851 0.01917 0.01360 -0.0282 0.0026 1.0000
9.250 1.2034 0.02047 0.01503 -0.0274 0.0024 1.0000
9.500 1.2211 0.02176 0.01645 -0.0266 0.0023 1.0000
9.750 1.2392 0.02291 0.01771 -0.0258 0.0022 1.0000
10.000 1.2573 0.02394 0.01883 -0.0251 0.0021 1.0000
10.250 1.2750 0.02492 0.01989 -0.0244 0.0020 1.0000
10.500 1.2918 0.02589 0.02099 -0.0236 0.0019 1.0000
10.750 1.3053 0.02721 0.02243 -0.0226 0.0018 1.0000
11.000 1.3141 0.02892 0.02429 -0.0210 0.0018 1.0000
11.250 1.3147 0.03109 0.02665 -0.0187 0.0017 1.0000
11.500 1.3097 0.03426 0.03008 -0.0166 0.0017 1.0000
11.750 1.2985 0.03862 0.03476 -0.0149 0.0016 1.0000
12.000 1.2924 0.04238 0.03877 -0.0141 0.0016 1.0000
12.250 1.2811 0.04698 0.04364 -0.0136 0.0016 1.0000
12.500 1.2592 0.05336 0.05033 -0.0136 0.0016 1.0000
12.750 1.2350 0.06026 0.05750 -0.0145 0.0016 1.0000
13.000 1.2110 0.06734 0.06480 -0.0162 0.0016 1.0000
13.250 1.1888 0.07438 0.07202 -0.0188 0.0016 1.0000
13.500 1.1674 0.08152 0.07933 -0.0220 0.0016 1.0000
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Polar data table (+)
Polar graphs
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