NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M16 AIRFOIL (m16-il) Reynolds number: 500,000 Max Cl/Cd: 92.65 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m16-il-500000.txt Download as CSV file: xf-m16-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5229 0.11456 0.11261 0.0262 1.0000 0.0139
-8.500 -0.5192 0.11112 0.10918 0.0233 1.0000 0.0139
-8.250 -0.5152 0.10757 0.10565 0.0204 1.0000 0.0140
-8.000 -0.5108 0.10389 0.10200 0.0173 1.0000 0.0140
-7.750 -0.5031 0.09985 0.09797 0.0135 1.0000 0.0140
-7.500 -0.5007 0.09431 0.09246 0.0159 1.0000 0.0145
-7.250 -0.4910 0.09081 0.08897 0.0138 1.0000 0.0148
-7.000 -0.4794 0.08734 0.08551 0.0109 1.0000 0.0153
-6.750 -0.4656 0.08389 0.08201 0.0075 0.9312 0.0161
-6.500 -0.4569 0.08084 0.07880 0.0060 0.8826 0.0171
-6.250 -0.4415 0.07724 0.07505 0.0027 0.8536 0.0181
-6.000 -0.4135 0.07371 0.07133 -0.0038 0.8308 0.0190
-5.750 -0.3903 0.06982 0.06728 -0.0076 0.8123 0.0192
-5.500 -0.3681 0.06579 0.06310 -0.0104 0.7962 0.0193
-5.250 -0.3451 0.06156 0.05872 -0.0129 0.7820 0.0194
-5.000 -0.3212 0.05727 0.05423 -0.0152 0.7693 0.0194
-4.500 -0.2897 0.04861 0.04535 -0.0174 0.7467 0.0211
-4.250 -0.2660 0.04632 0.04295 -0.0184 0.7358 0.0229
-4.000 -0.2284 0.04421 0.04057 -0.0201 0.7264 0.0267
-3.750 -0.1985 0.04128 0.03741 -0.0209 0.7176 0.0271
-3.250 -0.1477 0.03074 0.02625 -0.0226 0.7028 0.0286
-3.000 -0.1241 0.02936 0.02479 -0.0231 0.6941 0.0300
-2.750 -0.0979 0.02782 0.02308 -0.0233 0.6868 0.0324
-2.500 -0.0646 0.02734 0.02234 -0.0229 0.6789 0.0382
-2.250 -0.0370 0.02263 0.01707 -0.0230 0.6731 0.0398
-2.000 -0.0112 0.02093 0.01532 -0.0234 0.6656 0.0419
-1.750 0.0160 0.01992 0.01415 -0.0236 0.6591 0.0445
-1.500 0.0456 0.01937 0.01339 -0.0234 0.6519 0.0506
-1.250 0.0735 0.01730 0.01097 -0.0235 0.6459 0.0548
-1.000 0.1012 0.01655 0.01021 -0.0238 0.6389 0.0581
-0.500 0.1606 0.01362 0.00667 -0.0231 0.6264 0.0486
-0.250 0.1893 0.01247 0.00533 -0.0230 0.6200 0.0457
0.000 0.2177 0.01171 0.00444 -0.0228 0.6138 0.0439
0.250 0.2459 0.01127 0.00396 -0.0228 0.6070 0.0447
0.500 0.2739 0.01082 0.00344 -0.0228 0.6008 0.0440
0.750 0.3020 0.01041 0.00300 -0.0227 0.5939 0.0432
1.000 0.3300 0.01014 0.00268 -0.0227 0.5876 0.0432
1.250 0.3583 0.00988 0.00243 -0.0228 0.5803 0.0440
1.500 0.3864 0.00974 0.00225 -0.0228 0.5739 0.0461
1.750 0.4149 0.00959 0.00212 -0.0230 0.5664 0.0490
2.000 0.4431 0.00955 0.00203 -0.0231 0.5597 0.0516
2.250 0.4717 0.00945 0.00197 -0.0232 0.5519 0.0587
2.500 0.4942 0.00769 0.00204 -0.0229 0.5455 0.7989
2.750 0.5305 0.00730 0.00205 -0.0242 0.5370 1.0000
3.000 0.5579 0.00739 0.00208 -0.0242 0.5295 1.0000
3.250 0.5854 0.00747 0.00213 -0.0242 0.5214 1.0000
3.500 0.6130 0.00755 0.00222 -0.0242 0.5133 1.0000
3.750 0.6406 0.00766 0.00229 -0.0242 0.5048 1.0000
4.000 0.6682 0.00774 0.00234 -0.0242 0.4898 1.0000
4.250 0.6959 0.00783 0.00241 -0.0243 0.4723 1.0000
4.500 0.7236 0.00795 0.00254 -0.0244 0.4573 1.0000
4.750 0.7510 0.00813 0.00264 -0.0245 0.4297 1.0000
5.000 0.7783 0.00840 0.00280 -0.0246 0.3952 1.0000
5.250 0.8018 0.01026 0.00357 -0.0253 0.1771 1.0000
5.500 0.8249 0.01233 0.00491 -0.0258 0.0248 1.0000
5.750 0.8515 0.01288 0.00558 -0.0258 0.0218 1.0000
6.000 0.8777 0.01349 0.00628 -0.0258 0.0200 1.0000
6.250 0.9031 0.01424 0.00713 -0.0257 0.0184 1.0000
6.500 0.9273 0.01522 0.00818 -0.0255 0.0167 1.0000
6.750 0.9463 0.01723 0.01034 -0.0249 0.0152 1.0000
7.000 0.9687 0.01831 0.01151 -0.0244 0.0147 1.0000
7.250 0.9911 0.01936 0.01263 -0.0238 0.0143 1.0000
7.500 1.0123 0.02073 0.01407 -0.0230 0.0139 1.0000
7.750 1.0330 0.02248 0.01591 -0.0219 0.0138 1.0000
8.000 1.0540 0.02501 0.01853 -0.0208 0.0143 1.0000
8.250 1.0754 0.02902 0.02268 -0.0198 0.0149 1.0000
8.500 1.1012 0.02904 0.02277 -0.0193 0.0160 1.0000
8.750 1.1242 0.03300 0.02731 -0.0172 0.0196 1.0000
9.000 1.1426 0.03509 0.02959 -0.0164 0.0182 1.0000
9.250 1.1590 0.03746 0.03211 -0.0157 0.0172 1.0000
9.500 1.1750 0.04036 0.03510 -0.0151 0.0164 1.0000
9.750 1.1643 0.05149 0.04670 -0.0138 0.0150 1.0000
10.000 1.1678 0.05484 0.05034 -0.0123 0.0150 1.0000
10.250 1.1674 0.05806 0.05384 -0.0108 0.0150 1.0000
10.500 1.1628 0.06112 0.05715 -0.0092 0.0149 1.0000
10.750 1.1541 0.06395 0.06021 -0.0075 0.0149 1.0000
11.000 1.1389 0.06653 0.06296 -0.0055 0.0148 1.0000
11.250 1.1229 0.06969 0.06628 -0.0052 0.0148 1.0000
11.500 1.1068 0.07350 0.07025 -0.0060 0.0147 1.0000
11.750 1.0902 0.07803 0.07493 -0.0078 0.0147 1.0000
12.000 1.0725 0.08336 0.08040 -0.0103 0.0147 1.0000
12.250 1.0543 0.08937 0.08654 -0.0134 0.0147 1.0000
12.500 1.0359 0.09588 0.09316 -0.0170 0.0148 1.0000
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Polar data table (+)
Polar graphs
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