NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M2 AIRFOIL (m2-il) Reynolds number: 200,000 Max Cl/Cd: 36.81 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m2-il-200000-n5.txt Download as CSV file: xf-m2-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M2 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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.7792 0.08418 0.08082 0.0020 1.0000 0.0158
-10.000 -0.8379 0.06515 0.06167 -0.0136 1.0000 0.0149
-9.750 -0.8940 0.05106 0.04707 -0.0150 1.0000 0.0146
-9.500 -0.9019 0.04549 0.04111 -0.0142 1.0000 0.0149
-9.250 -0.9011 0.04091 0.03610 -0.0131 1.0000 0.0153
-9.000 -0.8947 0.03692 0.03165 -0.0118 1.0000 0.0160
-8.750 -0.8841 0.03331 0.02755 -0.0106 1.0000 0.0169
-8.500 -0.8688 0.03045 0.02419 -0.0095 1.0000 0.0183
-8.250 -0.8492 0.02852 0.02181 -0.0086 1.0000 0.0197
-8.000 -0.8309 0.02630 0.01936 -0.0078 1.0000 0.0210
-7.750 -0.8090 0.02514 0.01808 -0.0073 1.0000 0.0223
-7.500 -0.7862 0.02402 0.01674 -0.0068 1.0000 0.0242
-7.250 -0.7627 0.02295 0.01543 -0.0062 1.0000 0.0267
-7.000 -0.7384 0.02200 0.01422 -0.0057 1.0000 0.0284
-6.750 -0.7161 0.02029 0.01236 -0.0051 1.0000 0.0307
-6.500 -0.6923 0.01944 0.01146 -0.0047 1.0000 0.0332
-6.250 -0.6680 0.01859 0.01047 -0.0042 1.0000 0.0354
-6.000 -0.6433 0.01786 0.00961 -0.0037 1.0000 0.0377
-5.750 -0.6186 0.01721 0.00882 -0.0032 1.0000 0.0393
-5.500 -0.5953 0.01621 0.00778 -0.0025 1.0000 0.0411
-5.250 -0.5713 0.01552 0.00709 -0.0020 1.0000 0.0438
-5.000 -0.5469 0.01497 0.00647 -0.0015 1.0000 0.0463
-4.750 -0.5225 0.01445 0.00588 -0.0009 1.0000 0.0484
-4.500 -0.4981 0.01399 0.00535 -0.0003 1.0000 0.0506
-4.250 -0.4740 0.01352 0.00490 0.0003 1.0000 0.0545
-4.000 -0.4497 0.01314 0.00453 0.0009 1.0000 0.0609
-3.750 -0.4255 0.01281 0.00425 0.0016 1.0000 0.0733
-3.500 -0.4011 0.01257 0.00403 0.0022 1.0000 0.0895
-3.250 -0.3768 0.01235 0.00381 0.0028 1.0000 0.1037
-3.000 -0.3526 0.01209 0.00359 0.0033 1.0000 0.1151
-2.750 -0.3284 0.01186 0.00338 0.0039 1.0000 0.1246
-2.500 -0.3040 0.01167 0.00320 0.0045 1.0000 0.1344
-2.250 -0.2799 0.01141 0.00302 0.0050 1.0000 0.1520
-2.000 -0.2557 0.01111 0.00286 0.0055 1.0000 0.1825
-1.750 -0.2315 0.00997 0.00265 0.0053 0.9991 0.4068
-1.500 -0.1971 0.00935 0.00257 0.0035 0.9943 0.5324
-1.250 -0.1626 0.00876 0.00259 0.0021 0.9876 0.6650
-1.000 -0.1280 0.00844 0.00264 0.0011 0.9790 0.7590
-0.750 -0.0937 0.00831 0.00272 0.0003 0.9676 0.8165
-0.500 -0.0608 0.00830 0.00283 0.0000 0.9540 0.8626
-0.250 -0.0293 0.00833 0.00290 -0.0002 0.9394 0.8969
0.000 0.0000 0.00834 0.00291 0.0000 0.9209 0.9209
0.250 0.0293 0.00833 0.00290 0.0002 0.8969 0.9394
0.500 0.0609 0.00830 0.00283 0.0000 0.8629 0.9540
0.750 0.0937 0.00831 0.00272 -0.0003 0.8165 0.9676
1.000 0.1280 0.00844 0.00264 -0.0011 0.7591 0.9790
1.250 0.1628 0.00876 0.00259 -0.0021 0.6650 0.9877
1.500 0.1972 0.00935 0.00257 -0.0035 0.5329 0.9944
1.750 0.2315 0.00997 0.00265 -0.0053 0.4069 0.9991
2.000 0.2557 0.01111 0.00286 -0.0055 0.1823 1.0000
2.250 0.2798 0.01141 0.00302 -0.0050 0.1521 1.0000
2.500 0.3040 0.01167 0.00320 -0.0045 0.1344 1.0000
2.750 0.3283 0.01186 0.00338 -0.0039 0.1246 1.0000
3.000 0.3526 0.01209 0.00359 -0.0033 0.1151 1.0000
3.250 0.3767 0.01235 0.00381 -0.0027 0.1038 1.0000
3.500 0.4010 0.01257 0.00403 -0.0021 0.0896 1.0000
3.750 0.4254 0.01281 0.00425 -0.0015 0.0732 1.0000
4.000 0.4496 0.01314 0.00453 -0.0009 0.0609 1.0000
4.250 0.4739 0.01352 0.00490 -0.0003 0.0545 1.0000
4.500 0.4981 0.01399 0.00535 0.0003 0.0506 1.0000
4.750 0.5225 0.01445 0.00588 0.0009 0.0484 1.0000
5.000 0.5469 0.01497 0.00647 0.0015 0.0463 1.0000
5.250 0.5713 0.01552 0.00709 0.0020 0.0438 1.0000
5.500 0.5953 0.01621 0.00778 0.0025 0.0411 1.0000
5.750 0.6186 0.01721 0.00882 0.0032 0.0393 1.0000
6.000 0.6434 0.01786 0.00961 0.0037 0.0377 1.0000
6.250 0.6680 0.01859 0.01047 0.0042 0.0354 1.0000
6.500 0.6923 0.01944 0.01146 0.0047 0.0332 1.0000
6.750 0.7162 0.02029 0.01236 0.0051 0.0307 1.0000
7.000 0.7385 0.02200 0.01422 0.0057 0.0284 1.0000
7.250 0.7628 0.02295 0.01543 0.0062 0.0267 1.0000
7.500 0.7863 0.02402 0.01675 0.0068 0.0242 1.0000
7.750 0.8091 0.02514 0.01808 0.0073 0.0223 1.0000
8.000 0.8311 0.02630 0.01937 0.0078 0.0210 1.0000
8.250 0.8493 0.02854 0.02183 0.0085 0.0197 1.0000
8.500 0.8689 0.03046 0.02421 0.0094 0.0183 1.0000
8.750 0.8843 0.03332 0.02756 0.0105 0.0169 1.0000
9.000 0.8948 0.03694 0.03167 0.0118 0.0160 1.0000
9.250 0.9013 0.04093 0.03612 0.0130 0.0153 1.0000
9.500 0.9020 0.04552 0.04114 0.0141 0.0149 1.0000
9.750 0.8942 0.05111 0.04712 0.0149 0.0146 1.0000
10.000 0.8377 0.06531 0.06183 0.0135 0.0149 1.0000
10.250 0.7796 0.08434 0.08098 -0.0023 0.0158 1.0000
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Polar data table (+)
Polar graphs
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