NACA M11 AIRFOIL (m11-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M11 AIRFOIL (m11-il) Reynolds number: 500,000 Max Cl/Cd: 75.06 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m11-il-500000-n5.txt Download as CSV file: xf-m11-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.6613 0.06385 0.06165 -0.0315 1.0000 0.0109
-9.250 -0.6814 0.05614 0.05382 -0.0339 1.0000 0.0110
-9.000 -0.6840 0.05135 0.04890 -0.0341 1.0000 0.0113
-8.750 -0.6945 0.04428 0.04157 -0.0333 1.0000 0.0116
-8.500 -0.7685 0.02294 0.01852 -0.0238 1.0000 0.0144
-8.250 -0.7507 0.02206 0.01748 -0.0223 1.0000 0.0151
-8.000 -0.7270 0.02250 0.01797 -0.0215 1.0000 0.0155
-7.750 -0.7045 0.02277 0.01827 -0.0206 1.0000 0.0159
-7.500 -0.6848 0.02250 0.01792 -0.0192 1.0000 0.0165
-7.250 -0.6634 0.02169 0.01696 -0.0183 0.9991 0.0175
-7.000 -0.6333 0.02032 0.01529 -0.0192 0.9952 0.0193
-6.750 -0.6025 0.01963 0.01439 -0.0201 0.9908 0.0202
-6.500 -0.5731 0.01819 0.01273 -0.0210 0.9862 0.0211
-6.250 -0.5429 0.01774 0.01227 -0.0219 0.9807 0.0219
-6.000 -0.5107 0.01730 0.01175 -0.0231 0.9761 0.0228
-5.750 -0.4799 0.01667 0.01101 -0.0239 0.9700 0.0236
-5.500 -0.4476 0.01602 0.01025 -0.0251 0.9642 0.0247
-5.250 -0.4166 0.01536 0.00947 -0.0259 0.9564 0.0257
-5.000 -0.3835 0.01460 0.00857 -0.0271 0.9490 0.0265
-4.750 -0.3527 0.01407 0.00794 -0.0277 0.9382 0.0273
-4.500 -0.3219 0.01362 0.00739 -0.0283 0.9262 0.0279
-4.250 -0.2927 0.01311 0.00678 -0.0286 0.9123 0.0282
-4.000 -0.2648 0.01270 0.00626 -0.0285 0.8964 0.0284
-3.750 -0.2381 0.01230 0.00576 -0.0282 0.8798 0.0286
-3.500 -0.2130 0.01168 0.00504 -0.0275 0.8630 0.0289
-3.250 -0.1886 0.01112 0.00436 -0.0267 0.8464 0.0292
-3.000 -0.1645 0.01065 0.00380 -0.0258 0.8297 0.0298
-2.750 -0.1403 0.01029 0.00335 -0.0250 0.8137 0.0305
-2.500 -0.1157 0.01002 0.00300 -0.0242 0.7983 0.0308
-2.250 -0.0910 0.00977 0.00268 -0.0234 0.7833 0.0313
-2.000 -0.0661 0.00958 0.00241 -0.0227 0.7686 0.0318
-1.750 -0.0410 0.00941 0.00217 -0.0220 0.7541 0.0325
-1.500 -0.0157 0.00927 0.00196 -0.0214 0.7394 0.0334
-1.250 0.0098 0.00916 0.00179 -0.0208 0.7252 0.0345
-1.000 0.0354 0.00908 0.00164 -0.0202 0.7110 0.0357
-0.750 0.0613 0.00901 0.00151 -0.0197 0.6976 0.0369
-0.500 0.0872 0.00896 0.00141 -0.0192 0.6844 0.0381
-0.250 0.1132 0.00893 0.00132 -0.0187 0.6714 0.0394
0.000 0.1392 0.00888 0.00126 -0.0183 0.6580 0.0444
0.250 0.1642 0.00870 0.00121 -0.0176 0.6448 0.0906
0.500 0.1880 0.00842 0.00118 -0.0169 0.6317 0.1847
0.750 0.1864 0.00640 0.00114 -0.0111 0.6212 0.7833
1.250 0.3271 0.00662 0.00167 -0.0293 0.5861 0.9764
1.500 0.3635 0.00675 0.00173 -0.0312 0.5708 0.9821
1.750 0.4038 0.00691 0.00182 -0.0339 0.5545 0.9887
2.000 0.4511 0.00706 0.00189 -0.0382 0.5333 0.9958
2.250 0.4835 0.00717 0.00189 -0.0393 0.5088 0.9980
2.500 0.5154 0.00728 0.00192 -0.0403 0.4836 0.9998
3.000 0.5615 0.00772 0.00204 -0.0386 0.4052 1.0000
3.250 0.5840 0.00797 0.00214 -0.0376 0.3665 1.0000
3.500 0.6071 0.00819 0.00225 -0.0367 0.3378 1.0000
3.750 0.6301 0.00843 0.00239 -0.0358 0.3094 1.0000
4.000 0.6530 0.00870 0.00255 -0.0349 0.2796 1.0000
4.250 0.6752 0.00905 0.00275 -0.0339 0.2391 1.0000
4.500 0.6962 0.00954 0.00301 -0.0328 0.1876 1.0000
4.750 0.7173 0.01004 0.00331 -0.0317 0.1440 1.0000
5.000 0.7385 0.01052 0.00364 -0.0306 0.1079 1.0000
5.250 0.7599 0.01097 0.00396 -0.0295 0.0808 1.0000
5.500 0.7817 0.01137 0.00428 -0.0285 0.0638 1.0000
5.750 0.8039 0.01172 0.00459 -0.0275 0.0537 1.0000
6.000 0.8265 0.01202 0.00492 -0.0265 0.0475 1.0000
6.250 0.8484 0.01240 0.00527 -0.0255 0.0403 1.0000
6.500 0.8708 0.01271 0.00561 -0.0245 0.0358 1.0000
6.750 0.8924 0.01310 0.00599 -0.0234 0.0304 1.0000
7.000 0.9140 0.01348 0.00637 -0.0224 0.0248 1.0000
7.250 0.9352 0.01390 0.00682 -0.0212 0.0206 1.0000
7.500 0.9558 0.01438 0.00730 -0.0200 0.0172 1.0000
7.750 0.9760 0.01489 0.00785 -0.0187 0.0151 1.0000
8.000 0.9964 0.01535 0.00838 -0.0175 0.0137 1.0000
8.250 1.0160 0.01590 0.00895 -0.0161 0.0122 1.0000
8.500 1.0334 0.01665 0.00976 -0.0144 0.0109 1.0000
8.750 1.0523 0.01721 0.01042 -0.0130 0.0104 1.0000
9.000 1.0705 0.01783 0.01113 -0.0115 0.0099 1.0000
9.250 1.0880 0.01848 0.01186 -0.0099 0.0093 1.0000
9.500 1.1046 0.01921 0.01267 -0.0082 0.0089 1.0000
9.750 1.1214 0.01987 0.01340 -0.0066 0.0083 1.0000
10.000 1.1355 0.02074 0.01433 -0.0046 0.0079 1.0000
10.250 1.1466 0.02183 0.01551 -0.0022 0.0075 1.0000
10.500 1.1612 0.02257 0.01639 -0.0003 0.0072 1.0000
10.750 1.1735 0.02347 0.01739 0.0018 0.0069 1.0000
11.000 1.1838 0.02447 0.01852 0.0042 0.0067 1.0000
11.250 1.1905 0.02549 0.01965 0.0071 0.0066 1.0000
11.500 1.1960 0.02646 0.02072 0.0102 0.0063 1.0000
11.750 1.2009 0.02757 0.02194 0.0130 0.0062 1.0000
12.000 1.2068 0.02868 0.02315 0.0153 0.0059 1.0000
12.250 1.2080 0.03032 0.02494 0.0177 0.0059 1.0000
12.500 1.2125 0.03175 0.02646 0.0194 0.0058 1.0000
12.750 1.2149 0.03348 0.02831 0.0208 0.0056 1.0000
13.000 1.2148 0.03561 0.03058 0.0220 0.0056 1.0000
13.250 1.2141 0.03794 0.03306 0.0227 0.0055 1.0000
13.500 1.2087 0.04098 0.03624 0.0231 0.0053 1.0000
13.750 1.2050 0.04404 0.03944 0.0230 0.0053 1.0000
14.000 1.1920 0.04848 0.04407 0.0222 0.0052 1.0000
14.250 1.1844 0.05251 0.04825 0.0210 0.0052 1.0000
14.500 1.1772 0.05671 0.05261 0.0195 0.0051 1.0000
14.750 1.1640 0.06202 0.05810 0.0172 0.0051 1.0000
15.000 1.1556 0.06687 0.06309 0.0149 0.0051 1.0000
15.250 1.1365 0.07369 0.07009 0.0115 0.0051 1.0000
15.500 1.1178 0.08088 0.07744 0.0077 0.0051 1.0000
15.750 1.1028 0.08775 0.08444 0.0040 0.0051 1.0000
16.000 1.0806 0.09639 0.09324 -0.0006 0.0051 1.0000
16.250 1.0592 0.10522 0.10221 -0.0053 0.0051 1.0000
16.500 1.0349 0.11498 0.11210 -0.0104 0.0052 1.0000
16.750 1.0097 0.12542 0.12264 -0.0158 0.0052 1.0000
17.000 0.9873 0.13563 0.13297 -0.0208 0.0052 1.0000
17.250 0.9514 0.15018 0.14766 -0.0279 0.0053 1.0000
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