NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 500,000 Max Cl/Cd: 64.93 at α=3° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m10-il-500000-n5.txt Download as CSV file: xf-m10-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6095 0.09411 0.09197 0.0122 1.0000 0.0053
-8.500 -0.6086 0.08991 0.08779 0.0095 1.0000 0.0054
-8.250 -0.6084 0.08563 0.08354 0.0063 1.0000 0.0056
-8.000 -0.6061 0.08048 0.07840 0.0010 1.0000 0.0058
-7.750 -0.6003 0.07441 0.07231 -0.0052 1.0000 0.0060
-7.500 -0.5927 0.06752 0.06536 -0.0114 1.0000 0.0062
-7.250 -0.5840 0.05919 0.05691 -0.0177 1.0000 0.0063
-7.000 -0.6117 0.02016 0.01572 -0.0267 1.0000 0.0065
-6.750 -0.5887 0.01762 0.01271 -0.0263 1.0000 0.0069
-6.500 -0.5639 0.01605 0.01081 -0.0260 1.0000 0.0072
-6.250 -0.5388 0.01458 0.00908 -0.0257 1.0000 0.0081
-6.000 -0.5126 0.01386 0.00821 -0.0255 1.0000 0.0090
-5.750 -0.4862 0.01329 0.00753 -0.0254 1.0000 0.0103
-5.500 -0.4600 0.01251 0.00659 -0.0251 1.0000 0.0119
-5.250 -0.4331 0.01224 0.00634 -0.0249 1.0000 0.0145
-5.000 -0.4060 0.01207 0.00613 -0.0248 1.0000 0.0172
-4.750 -0.3790 0.01203 0.00603 -0.0247 1.0000 0.0190
-4.500 -0.3526 0.01172 0.00570 -0.0245 1.0000 0.0216
-4.250 -0.3260 0.01162 0.00558 -0.0243 1.0000 0.0238
-4.000 -0.2979 0.01141 0.00532 -0.0244 0.9976 0.0263
-3.750 -0.2642 0.01111 0.00494 -0.0258 0.9871 0.0284
-3.500 -0.2308 0.01095 0.00471 -0.0271 0.9738 0.0299
-3.250 -0.1981 0.01076 0.00446 -0.0281 0.9569 0.0305
-3.000 -0.1687 0.00990 0.00350 -0.0285 0.9369 0.0327
-2.750 -0.1418 0.00957 0.00309 -0.0282 0.9134 0.0343
-2.500 -0.1160 0.00936 0.00279 -0.0276 0.8899 0.0354
-2.250 -0.0903 0.00919 0.00252 -0.0270 0.8675 0.0363
-2.000 -0.0640 0.00904 0.00227 -0.0265 0.8460 0.0370
-1.750 -0.0375 0.00891 0.00203 -0.0261 0.8259 0.0377
-1.500 -0.0106 0.00880 0.00182 -0.0258 0.8062 0.0383
-1.250 0.0165 0.00872 0.00166 -0.0256 0.7875 0.0396
-1.000 0.0438 0.00865 0.00149 -0.0254 0.7695 0.0406
-0.750 0.0713 0.00859 0.00135 -0.0253 0.7517 0.0412
-0.500 0.0989 0.00854 0.00124 -0.0252 0.7343 0.0424
-0.250 0.1266 0.00848 0.00114 -0.0251 0.7177 0.0501
0.000 0.1541 0.00835 0.00109 -0.0251 0.7016 0.0869
0.250 0.1815 0.00814 0.00107 -0.0251 0.6860 0.1653
0.500 0.2079 0.00748 0.00106 -0.0253 0.6707 0.3999
0.750 0.2230 0.00587 0.00112 -0.0222 0.6569 0.9407
1.000 0.2703 0.00593 0.00111 -0.0266 0.6353 1.0000
1.250 0.2966 0.00606 0.00111 -0.0262 0.6075 1.0000
1.500 0.3231 0.00620 0.00113 -0.0259 0.5793 1.0000
1.750 0.3498 0.00634 0.00117 -0.0257 0.5563 1.0000
2.000 0.3765 0.00650 0.00121 -0.0255 0.5263 1.0000
2.250 0.4033 0.00667 0.00129 -0.0253 0.4983 1.0000
2.500 0.4302 0.00685 0.00137 -0.0252 0.4708 1.0000
2.750 0.4568 0.00710 0.00147 -0.0251 0.4296 1.0000
3.000 0.4831 0.00744 0.00162 -0.0249 0.3765 1.0000
3.250 0.5091 0.00789 0.00180 -0.0248 0.3090 1.0000
3.500 0.5347 0.00847 0.00206 -0.0248 0.2286 1.0000
3.750 0.5594 0.00933 0.00244 -0.0247 0.1195 1.0000
4.000 0.5848 0.01001 0.00285 -0.0246 0.0547 1.0000
4.250 0.6114 0.01035 0.00315 -0.0245 0.0425 1.0000
4.500 0.6382 0.01062 0.00345 -0.0244 0.0373 1.0000
4.750 0.6648 0.01096 0.00380 -0.0242 0.0324 1.0000
5.000 0.6915 0.01124 0.00412 -0.0241 0.0283 1.0000
5.250 0.7181 0.01153 0.00445 -0.0240 0.0224 1.0000
5.500 0.7443 0.01194 0.00487 -0.0239 0.0170 1.0000
5.750 0.7704 0.01235 0.00528 -0.0237 0.0135 1.0000
6.000 0.7951 0.01314 0.00618 -0.0232 0.0105 1.0000
6.250 0.8207 0.01364 0.00676 -0.0230 0.0097 1.0000
6.500 0.8461 0.01412 0.00730 -0.0227 0.0085 1.0000
6.750 0.8710 0.01468 0.00791 -0.0224 0.0076 1.0000
7.000 0.8937 0.01576 0.00910 -0.0218 0.0068 1.0000
7.250 0.9171 0.01666 0.01012 -0.0212 0.0064 1.0000
7.500 0.9400 0.01765 0.01130 -0.0206 0.0061 1.0000
7.750 0.9623 0.01880 0.01260 -0.0198 0.0057 1.0000
8.000 0.9849 0.01979 0.01372 -0.0192 0.0053 1.0000
8.250 1.0081 0.02052 0.01453 -0.0188 0.0049 1.0000
8.500 1.0302 0.02141 0.01551 -0.0184 0.0045 1.0000
8.750 1.0479 0.02344 0.01777 -0.0173 0.0043 1.0000
9.000 1.0668 0.02518 0.01979 -0.0163 0.0041 1.0000
9.250 1.0833 0.02740 0.02234 -0.0150 0.0040 1.0000
9.500 1.0960 0.03031 0.02566 -0.0135 0.0038 1.0000
9.750 1.1020 0.03431 0.03015 -0.0116 0.0037 1.0000
10.000 1.0957 0.04015 0.03658 -0.0090 0.0036 1.0000
10.250 1.0666 0.04891 0.04594 -0.0059 0.0035 1.0000
10.500 1.0355 0.05533 0.05267 -0.0036 0.0035 1.0000
10.750 1.0102 0.06141 0.05896 -0.0051 0.0036 1.0000
11.000 0.9869 0.06879 0.06651 -0.0101 0.0036 1.0000
11.250 0.9653 0.07804 0.07590 -0.0178 0.0037 1.0000
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Polar data table (+)
Polar graphs
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