NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 1,000,000 Max Cl/Cd: 75.27 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m10-il-1000000-n5.txt Download as CSV file: xf-m10-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M10 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6382 0.09655 0.09505 0.0155 1.0000 0.0033
-9.000 -0.6412 0.09138 0.08991 0.0125 1.0000 0.0033
-8.750 -0.7898 0.02400 0.02080 -0.0265 1.0000 0.0030
-8.500 -0.7726 0.01976 0.01607 -0.0261 1.0000 0.0032
-8.250 -0.7500 0.01782 0.01382 -0.0259 1.0000 0.0033
-8.000 -0.7257 0.01651 0.01231 -0.0257 1.0000 0.0035
-7.750 -0.7006 0.01547 0.01112 -0.0255 1.0000 0.0037
-7.500 -0.6751 0.01455 0.01004 -0.0253 1.0000 0.0039
-7.250 -0.6492 0.01375 0.00910 -0.0252 1.0000 0.0042
-7.000 -0.6231 0.01305 0.00827 -0.0250 1.0000 0.0045
-6.750 -0.5966 0.01248 0.00759 -0.0249 1.0000 0.0047
-6.500 -0.5705 0.01164 0.00662 -0.0247 1.0000 0.0053
-6.250 -0.5439 0.01106 0.00597 -0.0246 1.0000 0.0058
-6.000 -0.5171 0.01058 0.00541 -0.0245 1.0000 0.0064
-5.750 -0.4902 0.01015 0.00488 -0.0243 1.0000 0.0070
-5.500 -0.4633 0.00973 0.00439 -0.0242 1.0000 0.0080
-5.250 -0.4365 0.00932 0.00399 -0.0240 1.0000 0.0099
-5.000 -0.4067 0.00903 0.00371 -0.0246 0.9941 0.0134
-4.750 -0.3742 0.00892 0.00365 -0.0257 0.9803 0.0174
-4.500 -0.3425 0.00880 0.00348 -0.0266 0.9620 0.0187
-4.250 -0.3150 0.00862 0.00323 -0.0264 0.9370 0.0200
-4.000 -0.2897 0.00853 0.00307 -0.0257 0.9098 0.0222
-3.750 -0.2638 0.00848 0.00294 -0.0252 0.8847 0.0240
-3.500 -0.2372 0.00844 0.00281 -0.0248 0.8618 0.0263
-3.250 -0.2102 0.00843 0.00270 -0.0246 0.8403 0.0282
-3.000 -0.1829 0.00839 0.00258 -0.0244 0.8196 0.0293
-2.750 -0.1555 0.00834 0.00244 -0.0243 0.8003 0.0299
-2.500 -0.1279 0.00827 0.00228 -0.0242 0.7822 0.0303
-2.250 -0.1003 0.00820 0.00213 -0.0242 0.7642 0.0305
-2.000 -0.0725 0.00811 0.00198 -0.0241 0.7469 0.0306
-1.750 -0.0449 0.00795 0.00172 -0.0241 0.7299 0.0310
-1.500 -0.0171 0.00782 0.00150 -0.0241 0.7136 0.0314
-1.250 0.0109 0.00772 0.00132 -0.0241 0.6973 0.0318
-1.000 0.0389 0.00764 0.00117 -0.0241 0.6820 0.0322
-0.750 0.0669 0.00758 0.00105 -0.0242 0.6671 0.0329
-0.500 0.0950 0.00753 0.00095 -0.0242 0.6521 0.0340
-0.250 0.1231 0.00750 0.00087 -0.0243 0.6376 0.0346
0.000 0.1513 0.00748 0.00081 -0.0243 0.6232 0.0362
0.250 0.1793 0.00747 0.00077 -0.0244 0.6048 0.0399
0.500 0.2071 0.00745 0.00076 -0.0244 0.5769 0.0700
0.750 0.2349 0.00738 0.00076 -0.0245 0.5499 0.1246
1.250 0.2895 0.00669 0.00082 -0.0250 0.5039 0.4563
1.500 0.3051 0.00533 0.00089 -0.0223 0.4756 0.9606
1.750 0.3465 0.00548 0.00093 -0.0255 0.4428 1.0000
2.000 0.3735 0.00564 0.00098 -0.0253 0.4151 1.0000
2.250 0.4006 0.00582 0.00106 -0.0253 0.3854 1.0000
2.500 0.4270 0.00621 0.00118 -0.0252 0.3188 1.0000
2.750 0.4535 0.00661 0.00133 -0.0252 0.2561 1.0000
3.000 0.4800 0.00704 0.00152 -0.0251 0.1942 1.0000
3.250 0.5059 0.00765 0.00178 -0.0251 0.1094 1.0000
3.500 0.5322 0.00817 0.00205 -0.0251 0.0489 1.0000
3.750 0.5594 0.00839 0.00223 -0.0250 0.0389 1.0000
4.000 0.5867 0.00861 0.00243 -0.0250 0.0344 1.0000
4.250 0.6139 0.00883 0.00266 -0.0250 0.0305 1.0000
4.500 0.6413 0.00899 0.00284 -0.0250 0.0293 1.0000
4.750 0.6686 0.00917 0.00304 -0.0250 0.0268 1.0000
5.000 0.6956 0.00943 0.00327 -0.0249 0.0188 1.0000
5.250 0.7225 0.00970 0.00352 -0.0249 0.0146 1.0000
5.500 0.7493 0.01003 0.00384 -0.0248 0.0104 1.0000
5.750 0.7760 0.01031 0.00416 -0.0247 0.0090 1.0000
6.000 0.8024 0.01069 0.00455 -0.0246 0.0073 1.0000
6.250 0.8287 0.01109 0.00500 -0.0245 0.0065 1.0000
6.500 0.8549 0.01147 0.00543 -0.0243 0.0059 1.0000
6.750 0.8809 0.01188 0.00588 -0.0242 0.0054 1.0000
7.000 0.9066 0.01233 0.00637 -0.0240 0.0049 1.0000
7.250 0.9311 0.01303 0.00717 -0.0236 0.0043 1.0000
7.500 0.9565 0.01349 0.00769 -0.0234 0.0041 1.0000
7.750 0.9816 0.01400 0.00827 -0.0231 0.0038 1.0000
8.000 1.0062 0.01459 0.00894 -0.0228 0.0036 1.0000
8.250 1.0304 0.01522 0.00964 -0.0224 0.0033 1.0000
8.500 1.0544 0.01585 0.01035 -0.0221 0.0031 1.0000
8.750 1.0780 0.01654 0.01111 -0.0216 0.0030 1.0000
9.000 1.1007 0.01739 0.01205 -0.0211 0.0028 1.0000
9.250 1.1203 0.01880 0.01364 -0.0202 0.0026 1.0000
9.500 1.1417 0.01982 0.01480 -0.0196 0.0026 1.0000
9.750 1.1627 0.02086 0.01604 -0.0189 0.0025 1.0000
10.000 1.1825 0.02207 0.01742 -0.0180 0.0024 1.0000
10.250 1.2008 0.02348 0.01903 -0.0171 0.0023 1.0000
10.500 1.2173 0.02511 0.02087 -0.0160 0.0022 1.0000
10.750 1.2317 0.02698 0.02299 -0.0147 0.0022 1.0000
11.000 1.2433 0.02916 0.02542 -0.0133 0.0021 1.0000
11.250 1.2512 0.03164 0.02819 -0.0116 0.0020 1.0000
11.500 1.2553 0.03437 0.03119 -0.0098 0.0020 1.0000
11.750 1.2539 0.03740 0.03450 -0.0077 0.0019 1.0000
12.000 1.2416 0.04060 0.03795 -0.0046 0.0019 1.0000
12.250 1.2264 0.04434 0.04191 -0.0031 0.0019 1.0000
12.500 1.2029 0.05006 0.04788 -0.0042 0.0019 1.0000
12.750 1.1687 0.05901 0.05712 -0.0092 0.0019 1.0000
13.000 1.1296 0.07126 0.06962 -0.0181 0.0020 1.0000
13.250 1.0640 0.09480 0.09341 -0.0336 0.0021 1.0000
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Polar data table (+)
Polar graphs
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