NACA M4 AIRFOIL (m4-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M4 AIRFOIL (m4-il) Reynolds number: 1,000,000 Max Cl/Cd: 73.25 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m4-il-1000000-n5.txt Download as CSV file: xf-m4-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M4 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6618 0.09794 0.09654 0.0323 1.0000 0.0032
-8.500 -0.6634 0.09331 0.09193 0.0292 1.0000 0.0033
-7.500 -0.7406 0.01818 0.01421 -0.0002 1.0000 0.0039
-7.250 -0.7168 0.01639 0.01215 0.0002 1.0000 0.0041
-7.000 -0.6915 0.01516 0.01071 0.0004 1.0000 0.0044
-6.750 -0.6654 0.01426 0.00965 0.0006 1.0000 0.0046
-6.500 -0.6384 0.01319 0.00838 0.0006 0.9531 0.0051
-6.250 -0.6166 0.01265 0.00763 0.0020 0.9099 0.0055
-6.000 -0.5924 0.01215 0.00695 0.0028 0.8807 0.0060
-5.750 -0.5669 0.01170 0.00631 0.0034 0.8555 0.0066
-5.500 -0.5407 0.01129 0.00572 0.0037 0.8321 0.0070
-5.250 -0.5142 0.01087 0.00517 0.0041 0.8102 0.0084
-5.000 -0.4875 0.01050 0.00471 0.0043 0.7898 0.0102
-4.750 -0.4601 0.01035 0.00452 0.0044 0.7715 0.0134
-4.500 -0.4322 0.01038 0.00452 0.0044 0.7557 0.0161
-4.250 -0.4045 0.01028 0.00433 0.0044 0.7414 0.0171
-4.000 -0.3766 0.01017 0.00416 0.0044 0.7284 0.0186
-3.750 -0.3483 0.01029 0.00427 0.0043 0.7162 0.0201
-3.500 -0.3203 0.01024 0.00418 0.0043 0.7051 0.0216
-3.250 -0.2923 0.01015 0.00403 0.0043 0.6946 0.0234
-2.750 -0.2361 0.01009 0.00386 0.0042 0.6743 0.0266
-2.500 -0.2081 0.01000 0.00372 0.0041 0.6648 0.0273
-2.250 -0.1800 0.00992 0.00358 0.0041 0.6553 0.0278
-2.000 -0.1520 0.00981 0.00342 0.0040 0.6459 0.0281
-1.750 -0.1239 0.00973 0.00331 0.0040 0.6368 0.0284
-1.250 -0.0688 0.00902 0.00247 0.0041 0.6183 0.0292
-1.000 -0.0410 0.00878 0.00218 0.0041 0.6092 0.0294
-0.750 -0.0131 0.00859 0.00193 0.0041 0.5998 0.0296
-0.500 0.0149 0.00842 0.00172 0.0041 0.5902 0.0298
-0.250 0.0429 0.00827 0.00153 0.0040 0.5806 0.0305
0.000 0.0711 0.00817 0.00139 0.0040 0.5705 0.0304
0.250 0.0993 0.00809 0.00127 0.0039 0.5601 0.0303
0.500 0.1276 0.00802 0.00117 0.0038 0.5499 0.0302
0.750 0.1558 0.00797 0.00109 0.0037 0.5393 0.0302
1.250 0.2125 0.00791 0.00097 0.0035 0.5167 0.0306
1.500 0.2408 0.00793 0.00092 0.0033 0.4988 0.0313
1.750 0.2690 0.00797 0.00090 0.0032 0.4746 0.0350
2.250 0.3196 0.00635 0.00099 0.0032 0.4305 0.7268
2.500 0.3430 0.00575 0.00109 0.0046 0.4108 0.9756
2.750 0.3823 0.00590 0.00119 0.0019 0.3914 0.9916
3.000 0.4212 0.00608 0.00130 -0.0008 0.3696 0.9970
3.250 0.4573 0.00640 0.00144 -0.0029 0.3208 0.9997
3.500 0.4851 0.00667 0.00157 -0.0032 0.2821 1.0000
3.750 0.5111 0.00708 0.00174 -0.0031 0.2234 1.0000
4.000 0.5370 0.00762 0.00200 -0.0031 0.1554 1.0000
4.250 0.5626 0.00832 0.00235 -0.0031 0.0722 1.0000
4.500 0.5884 0.00873 0.00261 -0.0030 0.0403 1.0000
4.750 0.6146 0.00894 0.00281 -0.0028 0.0345 1.0000
5.000 0.6407 0.00917 0.00303 -0.0026 0.0294 1.0000
5.250 0.6670 0.00935 0.00324 -0.0024 0.0271 1.0000
5.500 0.6932 0.00963 0.00349 -0.0022 0.0179 1.0000
5.750 0.7194 0.00993 0.00376 -0.0021 0.0136 1.0000
6.000 0.7458 0.01025 0.00409 -0.0019 0.0102 1.0000
6.250 0.7721 0.01054 0.00443 -0.0019 0.0089 1.0000
6.500 0.7983 0.01091 0.00481 -0.0018 0.0073 1.0000
6.750 0.8244 0.01133 0.00527 -0.0017 0.0065 1.0000
7.000 0.8506 0.01171 0.00570 -0.0016 0.0060 1.0000
7.250 0.8765 0.01212 0.00617 -0.0015 0.0054 1.0000
7.500 0.9022 0.01258 0.00668 -0.0014 0.0049 1.0000
7.750 0.9272 0.01323 0.00739 -0.0013 0.0044 1.0000
8.000 0.9525 0.01374 0.00796 -0.0012 0.0041 1.0000
8.250 0.9776 0.01426 0.00856 -0.0010 0.0038 1.0000
8.500 1.0022 0.01487 0.00925 -0.0008 0.0036 1.0000
8.750 1.0265 0.01550 0.00996 -0.0006 0.0034 1.0000
9.000 1.0505 0.01617 0.01069 -0.0004 0.0032 1.0000
9.250 1.0741 0.01684 0.01144 -0.0002 0.0030 1.0000
9.500 1.0968 0.01768 0.01237 0.0002 0.0029 1.0000
9.750 1.1169 0.01898 0.01384 0.0007 0.0027 1.0000
10.000 1.1373 0.02014 0.01515 0.0013 0.0026 1.0000
10.250 1.1579 0.02116 0.01631 0.0018 0.0025 1.0000
10.500 1.1772 0.02236 0.01767 0.0024 0.0024 1.0000
10.750 1.1950 0.02373 0.01922 0.0031 0.0024 1.0000
11.000 1.2110 0.02529 0.02097 0.0040 0.0023 1.0000
11.250 1.2247 0.02705 0.02294 0.0050 0.0022 1.0000
11.500 1.2361 0.02895 0.02505 0.0061 0.0021 1.0000
11.750 1.2451 0.03094 0.02725 0.0072 0.0020 1.0000
12.000 1.2512 0.03299 0.02950 0.0085 0.0020 1.0000
12.250 1.2495 0.03517 0.03186 0.0104 0.0019 1.0000
12.500 1.2456 0.03778 0.03464 0.0111 0.0019 1.0000
12.750 1.2396 0.04117 0.03822 0.0107 0.0019 1.0000
13.000 1.2334 0.04505 0.04226 0.0092 0.0018 1.0000
13.250 1.2217 0.05011 0.04752 0.0068 0.0018 1.0000
13.500 1.2044 0.05648 0.05409 0.0033 0.0018 1.0000
13.750 1.1846 0.06383 0.06162 -0.0012 0.0018 1.0000
14.000 1.1533 0.07400 0.07200 -0.0076 0.0019 1.0000
14.250 1.1037 0.08923 0.08747 -0.0168 0.0019 1.0000
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Polar data table (+)
Polar graphs
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