NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 200,000 Max Cl/Cd: 54.94 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m10-il-200000-n5.txt Download as CSV file: xf-m10-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5816 0.09642 0.09302 0.0099 1.0000 0.0131
-8.250 -0.5820 0.09164 0.08828 0.0069 1.0000 0.0128
-8.000 -0.5838 0.08661 0.08330 0.0031 1.0000 0.0125
-7.750 -0.5817 0.08068 0.07738 -0.0029 1.0000 0.0123
-7.500 -0.5756 0.07455 0.07123 -0.0085 1.0000 0.0122
-7.250 -0.5664 0.06887 0.06550 -0.0132 1.0000 0.0122
-7.000 -0.5543 0.06369 0.06023 -0.0170 1.0000 0.0124
-6.750 -0.5387 0.06037 0.05683 -0.0192 1.0000 0.0133
-6.500 -0.5221 0.05583 0.05217 -0.0219 1.0000 0.0143
-6.250 -0.5047 0.05000 0.04612 -0.0246 1.0000 0.0150
-6.000 -0.4864 0.04348 0.03929 -0.0266 1.0000 0.0153
-5.750 -0.4670 0.03634 0.03168 -0.0277 1.0000 0.0159
-5.500 -0.4476 0.02739 0.02185 -0.0276 1.0000 0.0176
-5.250 -0.4248 0.02340 0.01722 -0.0272 1.0000 0.0199
-5.000 -0.4001 0.02259 0.01629 -0.0271 1.0000 0.0217
-4.750 -0.3749 0.02111 0.01447 -0.0267 1.0000 0.0246
-4.500 -0.3489 0.01960 0.01252 -0.0262 1.0000 0.0290
-4.250 -0.3233 0.01813 0.01068 -0.0257 1.0000 0.0314
-4.000 -0.2984 0.01740 0.00993 -0.0255 1.0000 0.0348
-3.750 -0.2729 0.01654 0.00888 -0.0251 1.0000 0.0373
-3.500 -0.2476 0.01587 0.00803 -0.0245 1.0000 0.0402
-3.250 -0.2226 0.01504 0.00704 -0.0239 1.0000 0.0414
-3.000 -0.1983 0.01419 0.00610 -0.0232 1.0000 0.0426
-2.750 -0.1746 0.01335 0.00525 -0.0225 1.0000 0.0445
-2.500 -0.1416 0.01279 0.00464 -0.0238 0.9933 0.0464
-2.250 -0.1063 0.01231 0.00413 -0.0256 0.9825 0.0482
-2.000 -0.0712 0.01196 0.00376 -0.0273 0.9698 0.0516
-1.750 -0.0367 0.01161 0.00335 -0.0287 0.9554 0.0539
-1.500 -0.0029 0.01129 0.00297 -0.0300 0.9394 0.0551
-1.250 0.0290 0.01095 0.00262 -0.0308 0.9211 0.0589
-1.000 0.0590 0.01073 0.00237 -0.0310 0.9013 0.0671
-0.750 0.0869 0.01045 0.00218 -0.0309 0.8811 0.0985
-0.500 0.1128 0.00998 0.00206 -0.0306 0.8605 0.2095
-0.250 0.1259 0.00783 0.00211 -0.0269 0.8421 0.8762
0.250 0.2052 0.00785 0.00197 -0.0314 0.8037 1.0000
0.500 0.2303 0.00794 0.00192 -0.0307 0.7846 1.0000
0.750 0.2559 0.00803 0.00190 -0.0300 0.7657 1.0000
1.000 0.2816 0.00813 0.00190 -0.0295 0.7474 1.0000
1.250 0.3075 0.00823 0.00192 -0.0289 0.7298 1.0000
1.500 0.3336 0.00835 0.00196 -0.0284 0.7124 1.0000
1.750 0.3598 0.00846 0.00202 -0.0280 0.6948 1.0000
2.000 0.3861 0.00859 0.00209 -0.0276 0.6772 1.0000
2.250 0.4122 0.00873 0.00218 -0.0271 0.6550 1.0000
2.500 0.4379 0.00891 0.00226 -0.0266 0.6247 1.0000
2.750 0.4636 0.00911 0.00234 -0.0261 0.5893 1.0000
3.000 0.4895 0.00932 0.00248 -0.0256 0.5568 1.0000
3.250 0.5154 0.00956 0.00263 -0.0252 0.5208 1.0000
3.500 0.5409 0.00987 0.00279 -0.0247 0.4729 1.0000
3.750 0.5659 0.01030 0.00301 -0.0243 0.4080 1.0000
4.000 0.5904 0.01088 0.00329 -0.0239 0.3301 1.0000
4.250 0.6130 0.01194 0.00376 -0.0236 0.1968 1.0000
4.500 0.6350 0.01330 0.00449 -0.0233 0.0743 1.0000
4.750 0.6601 0.01398 0.00511 -0.0230 0.0546 1.0000
5.000 0.6854 0.01455 0.00576 -0.0227 0.0444 1.0000
5.250 0.7102 0.01523 0.00650 -0.0223 0.0358 1.0000
5.500 0.7338 0.01616 0.00749 -0.0218 0.0295 1.0000
5.750 0.7587 0.01679 0.00826 -0.0214 0.0242 1.0000
6.000 0.7821 0.01775 0.00932 -0.0208 0.0207 1.0000
6.250 0.8037 0.01911 0.01076 -0.0199 0.0179 1.0000
6.500 0.8270 0.02017 0.01197 -0.0192 0.0160 1.0000
6.750 0.8494 0.02163 0.01358 -0.0183 0.0146 1.0000
7.000 0.8719 0.02322 0.01535 -0.0174 0.0135 1.0000
7.250 0.8945 0.02440 0.01666 -0.0168 0.0122 1.0000
7.500 0.9137 0.02684 0.01936 -0.0159 0.0108 1.0000
7.750 0.9329 0.02966 0.02261 -0.0148 0.0103 1.0000
8.000 0.9510 0.03242 0.02580 -0.0135 0.0100 1.0000
8.250 0.9658 0.03576 0.02961 -0.0121 0.0097 1.0000
8.500 0.9765 0.03964 0.03398 -0.0106 0.0096 1.0000
8.750 0.9825 0.04399 0.03880 -0.0090 0.0095 1.0000
9.000 0.9836 0.04859 0.04382 -0.0074 0.0094 1.0000
9.250 0.9806 0.05318 0.04878 -0.0060 0.0093 1.0000
9.500 0.9727 0.05774 0.05364 -0.0049 0.0092 1.0000
9.750 0.9578 0.06202 0.05815 -0.0039 0.0092 1.0000
10.000 0.9395 0.06677 0.06308 -0.0048 0.0092 1.0000
10.250 0.9205 0.07286 0.06932 -0.0087 0.0094 1.0000
10.500 0.9015 0.08071 0.07729 -0.0152 0.0096 1.0000
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Polar data table (+)
Polar graphs
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