NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M10 AIRFOIL (m10-il) Reynolds number: 200,000 Max Cl/Cd: 58.75 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m10-il-200000.txt Download as CSV file: xf-m10-il-200000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.5775 0.10085 0.09739 0.0120 1.0000 0.0356
-8.250 -0.5756 0.09698 0.09355 0.0086 1.0000 0.0368
-8.000 -0.5754 0.09302 0.08964 0.0026 1.0000 0.0377
-7.750 -0.5671 0.08829 0.08489 -0.0062 1.0000 0.0383
-7.500 -0.5558 0.08356 0.08008 -0.0121 1.0000 0.0386
-7.250 -0.5437 0.07883 0.07524 -0.0161 1.0000 0.0387
-7.000 -0.5418 0.07118 0.06759 -0.0184 1.0000 0.0395
-6.750 -0.5351 0.06807 0.06456 -0.0161 1.0000 0.0410
-6.500 -0.5226 0.06521 0.06168 -0.0163 1.0000 0.0437
-6.250 -0.5060 0.06101 0.05737 -0.0191 1.0000 0.0468
-6.000 -0.4745 0.05716 0.05290 -0.0248 1.0000 0.0512
-5.750 -0.4652 0.04995 0.04569 -0.0259 1.0000 0.0526
-5.500 -0.4494 0.04705 0.04284 -0.0257 1.0000 0.0546
-5.250 -0.4294 0.04418 0.03982 -0.0260 1.0000 0.0587
-5.000 -0.4049 0.03966 0.03475 -0.0273 1.0000 0.0662
-4.750 -0.3859 0.03754 0.03275 -0.0271 1.0000 0.0711
-4.500 -0.3616 0.03422 0.02896 -0.0274 1.0000 0.0807
-4.250 -0.3396 0.03193 0.02666 -0.0272 1.0000 0.0858
-3.750 -0.2835 0.02325 0.01648 -0.0254 1.0000 0.0663
-3.500 -0.2575 0.02059 0.01355 -0.0246 1.0000 0.0606
-3.250 -0.2319 0.01879 0.01142 -0.0238 1.0000 0.0612
-3.000 -0.2065 0.01761 0.00997 -0.0230 1.0000 0.0630
-2.750 -0.1818 0.01636 0.00852 -0.0221 1.0000 0.0634
-2.500 -0.1576 0.01546 0.00749 -0.0212 1.0000 0.0641
-2.250 -0.1342 0.01418 0.00616 -0.0205 1.0000 0.0669
-2.000 -0.1113 0.01350 0.00551 -0.0196 1.0000 0.0695
-1.750 -0.0888 0.01292 0.00496 -0.0187 1.0000 0.0715
-1.500 -0.0668 0.01248 0.00454 -0.0179 1.0000 0.0744
-1.000 0.0023 0.01145 0.00359 -0.0214 0.9913 0.0890
-0.750 0.0432 0.00982 0.00307 -0.0249 0.9844 0.3446
-0.500 0.0901 0.00808 0.00310 -0.0279 0.9840 1.0000
-0.250 0.1360 0.00803 0.00292 -0.0317 0.9712 1.0000
0.000 0.1791 0.00799 0.00278 -0.0349 0.9558 1.0000
0.250 0.2150 0.00796 0.00269 -0.0364 0.9350 1.0000
0.500 0.2453 0.00795 0.00261 -0.0366 0.9136 1.0000
0.750 0.2701 0.00800 0.00258 -0.0355 0.8899 1.0000
1.000 0.2936 0.00807 0.00257 -0.0341 0.8680 1.0000
1.250 0.3172 0.00817 0.00259 -0.0328 0.8462 1.0000
1.500 0.3411 0.00829 0.00264 -0.0316 0.8258 1.0000
1.750 0.3657 0.00841 0.00270 -0.0307 0.8050 1.0000
2.000 0.3904 0.00855 0.00277 -0.0297 0.7859 1.0000
2.250 0.4157 0.00869 0.00287 -0.0289 0.7643 1.0000
2.500 0.4402 0.00882 0.00294 -0.0278 0.7383 1.0000
2.750 0.4648 0.00895 0.00299 -0.0268 0.7084 1.0000
3.000 0.4899 0.00909 0.00306 -0.0259 0.6791 1.0000
3.250 0.5153 0.00926 0.00317 -0.0251 0.6496 1.0000
3.500 0.5400 0.00944 0.00325 -0.0242 0.6063 1.0000
3.750 0.5647 0.00968 0.00335 -0.0233 0.5550 1.0000
4.000 0.5893 0.01003 0.00351 -0.0225 0.4915 1.0000
4.250 0.6094 0.01119 0.00380 -0.0214 0.2968 1.0000
4.750 0.6514 0.01465 0.00601 -0.0202 0.0626 1.0000
5.000 0.6758 0.01551 0.00693 -0.0196 0.0548 1.0000
5.250 0.6981 0.01691 0.00834 -0.0187 0.0493 1.0000
5.500 0.7224 0.01787 0.00935 -0.0181 0.0438 1.0000
5.750 0.7454 0.01948 0.01093 -0.0172 0.0405 1.0000
6.000 0.7688 0.02224 0.01380 -0.0163 0.0385 1.0000
6.250 0.7945 0.02305 0.01485 -0.0156 0.0357 1.0000
6.500 0.8195 0.02505 0.01711 -0.0147 0.0344 1.0000
6.750 0.8436 0.02769 0.02011 -0.0137 0.0344 1.0000
7.000 0.8657 0.03125 0.02409 -0.0124 0.0359 1.0000
7.250 0.8873 0.03588 0.02892 -0.0115 0.0397 1.0000
14.250 0.6558 0.16055 0.15719 -0.0328 0.0386 1.0000
14.500 0.6516 0.16341 0.16003 -0.0349 0.0383 1.0000
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Polar data table (+)
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