NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 50,000 Max Cl/Cd: 34.37 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m10-il-50000.txt Download as CSV file: xf-m10-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5906 0.11194 0.10526 0.0130 1.0000 0.1764
-8.250 -0.5752 0.10664 0.09995 0.0147 1.0000 0.1864
-8.000 -0.5922 0.10530 0.09875 0.0089 1.0000 0.1910
-7.750 -0.5761 0.10008 0.09353 0.0114 1.0000 0.2033
-7.500 -0.5685 0.09586 0.08935 0.0116 1.0000 0.2131
-7.250 -0.5667 0.09218 0.08573 0.0095 1.0000 0.2239
-7.000 -0.5625 0.08848 0.08206 0.0080 1.0000 0.2374
-6.750 -0.5547 0.08469 0.07827 0.0081 1.0000 0.2541
-6.500 -0.5548 0.08178 0.07538 0.0046 1.0000 0.2757
-6.250 -0.5468 0.07813 0.07177 0.0064 1.0000 0.3038
-6.000 -0.5354 0.07423 0.06794 0.0113 1.0000 0.3379
-5.750 -0.5297 0.07107 0.06483 0.0142 1.0000 0.3784
-5.500 -0.5189 0.06790 0.06172 0.0200 1.0000 0.4261
-5.250 -0.2061 0.04887 0.04162 0.0139 1.0000 1.0000
-5.000 -0.1935 0.04624 0.03900 0.0123 1.0000 1.0000
-4.750 -0.2213 0.04653 0.03943 0.0204 1.0000 0.9751
-4.500 -0.2886 0.04879 0.04199 0.0347 1.0000 0.9020
-4.250 -0.3561 0.04955 0.04306 0.0428 1.0000 0.8043
-4.000 -0.3384 0.03848 0.02953 -0.0265 1.0000 0.1922
-3.750 -0.3083 0.03480 0.02535 -0.0269 1.0000 0.1709
-3.500 -0.2776 0.03222 0.02195 -0.0271 1.0000 0.1625
-3.250 -0.2502 0.02998 0.01924 -0.0267 1.0000 0.1630
-3.000 -0.2232 0.02767 0.01668 -0.0263 1.0000 0.1611
-2.750 -0.1957 0.02580 0.01447 -0.0257 1.0000 0.1609
-2.500 -0.1683 0.02427 0.01254 -0.0250 1.0000 0.1627
-2.250 -0.1424 0.02278 0.01108 -0.0244 1.0000 0.1721
-2.000 -0.1158 0.02157 0.00980 -0.0237 1.0000 0.1847
-1.750 -0.0876 0.02041 0.00857 -0.0231 1.0000 0.1969
-1.500 -0.0618 0.01934 0.00751 -0.0222 1.0000 0.2193
-1.250 -0.0113 0.01489 0.00569 -0.0243 1.0000 1.0000
-1.000 0.0095 0.01495 0.00525 -0.0229 1.0000 1.0000
-0.750 0.0294 0.01503 0.00506 -0.0216 1.0000 1.0000
-0.500 0.0487 0.01515 0.00497 -0.0204 1.0000 1.0000
-0.250 0.0676 0.01532 0.00498 -0.0192 1.0000 1.0000
0.000 0.0863 0.01554 0.00504 -0.0181 1.0000 1.0000
0.250 0.1049 0.01581 0.00521 -0.0172 1.0000 1.0000
0.500 0.1233 0.01615 0.00545 -0.0163 1.0000 1.0000
0.750 0.1417 0.01654 0.00577 -0.0157 1.0000 1.0000
1.000 0.1600 0.01700 0.00618 -0.0152 1.0000 1.0000
1.250 0.1782 0.01752 0.00668 -0.0149 1.0000 1.0000
1.500 0.1964 0.01812 0.00725 -0.0147 1.0000 1.0000
1.750 0.2144 0.01878 0.00792 -0.0147 1.0000 1.0000
2.000 0.2322 0.01953 0.00869 -0.0149 1.0000 1.0000
2.250 0.2498 0.02035 0.00955 -0.0152 1.0000 1.0000
2.500 0.3116 0.02132 0.01068 -0.0236 0.9799 1.0000
2.750 0.3756 0.02215 0.01174 -0.0320 0.9569 1.0000
3.000 0.4331 0.02282 0.01272 -0.0387 0.9318 1.0000
3.250 0.4925 0.02332 0.01357 -0.0451 0.9063 1.0000
3.500 0.5470 0.02365 0.01428 -0.0497 0.8799 1.0000
3.750 0.5927 0.02386 0.01492 -0.0519 0.8517 1.0000
4.000 0.6318 0.02366 0.01509 -0.0512 0.8177 1.0000
4.250 0.6527 0.02156 0.01308 -0.0412 0.7487 1.0000
4.500 0.6651 0.02041 0.01189 -0.0327 0.6842 1.0000
4.750 0.6781 0.01973 0.01128 -0.0259 0.6056 1.0000
5.000 0.6749 0.02133 0.01072 -0.0170 0.2556 1.0000
5.250 0.6928 0.02430 0.01281 -0.0154 0.1765 1.0000
5.500 0.7152 0.02633 0.01463 -0.0140 0.1474 1.0000
5.750 0.7417 0.02843 0.01667 -0.0127 0.1330 1.0000
6.000 0.7693 0.03072 0.01904 -0.0117 0.1218 1.0000
6.250 0.7960 0.03335 0.02186 -0.0109 0.1134 1.0000
6.500 0.8229 0.03631 0.02529 -0.0099 0.1110 1.0000
6.750 0.8472 0.03959 0.02909 -0.0089 0.1093 1.0000
7.000 0.8689 0.04290 0.03270 -0.0081 0.1053 1.0000
7.250 0.8879 0.04681 0.03716 -0.0071 0.1051 1.0000
7.500 0.9009 0.05181 0.04304 -0.0061 0.1105 1.0000
7.750 0.9154 0.05701 0.04856 -0.0056 0.1150 1.0000
8.000 0.9150 0.06330 0.05573 -0.0059 0.1255 1.0000
8.250 0.9085 0.07045 0.06343 -0.0076 0.1377 1.0000
8.500 0.8147 0.06590 0.05945 -0.0018 0.1362 1.0000
8.750 0.8012 0.07261 0.06626 -0.0028 0.1452 1.0000
9.000 0.7562 0.07989 0.07354 -0.0069 0.1473 1.0000
9.250 0.7197 0.09094 0.08452 -0.0148 0.1675 1.0000
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Polar data table (+)
Polar graphs
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