NACA M4 AIRFOIL (m4-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M4 AIRFOIL (m4-il) Reynolds number: 200,000 Max Cl/Cd: 55.43 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m4-il-200000-n5.txt Download as CSV file: xf-m4-il-200000-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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6068 0.09957 0.09637 0.0263 1.0000 0.0140
-8.000 -0.6059 0.09524 0.09207 0.0235 1.0000 0.0137
-7.750 -0.6057 0.09081 0.08767 0.0201 1.0000 0.0134
-7.500 -0.6018 0.08598 0.08285 0.0160 1.0000 0.0132
-7.250 -0.5949 0.08094 0.07779 0.0118 1.0000 0.0131
-7.000 -0.5856 0.07576 0.07257 0.0077 1.0000 0.0130
-6.750 -0.5736 0.07072 0.06747 0.0041 1.0000 0.0132
-6.500 -0.5589 0.06578 0.06243 0.0008 1.0000 0.0141
-6.250 -0.5422 0.06025 0.05676 -0.0024 1.0000 0.0149
-6.000 -0.5244 0.05451 0.05084 -0.0048 1.0000 0.0151
-5.750 -0.5053 0.04857 0.04465 -0.0065 1.0000 0.0154
-5.500 -0.4849 0.04204 0.03778 -0.0075 1.0000 0.0159
-5.250 -0.4631 0.03358 0.02869 -0.0074 1.0000 0.0171
-5.000 -0.4414 0.03233 0.02734 -0.0078 1.0000 0.0187
-4.750 -0.4173 0.03065 0.02545 -0.0079 1.0000 0.0202
-4.500 -0.3926 0.02710 0.02147 -0.0075 1.0000 0.0221
-4.250 -0.3662 0.02388 0.01769 -0.0070 1.0000 0.0263
-4.000 -0.3384 0.02216 0.01546 -0.0066 1.0000 0.0284
-3.750 -0.3124 0.02003 0.01313 -0.0070 0.9859 0.0317
-3.500 -0.2794 0.01893 0.01178 -0.0081 0.9416 0.0345
-3.250 -0.2514 0.01793 0.01044 -0.0078 0.9114 0.0371
-3.000 -0.2260 0.01716 0.00938 -0.0069 0.8857 0.0391
-2.750 -0.2009 0.01635 0.00830 -0.0059 0.8638 0.0399
-2.500 -0.1757 0.01575 0.00749 -0.0050 0.8440 0.0405
-2.250 -0.1508 0.01463 0.00620 -0.0042 0.8266 0.0419
-2.000 -0.1255 0.01399 0.00545 -0.0035 0.8100 0.0430
-1.750 -0.0998 0.01354 0.00493 -0.0029 0.7946 0.0447
-1.500 -0.0737 0.01321 0.00452 -0.0025 0.7800 0.0470
-1.250 -0.0475 0.01284 0.00407 -0.0020 0.7664 0.0477
-1.000 -0.0212 0.01251 0.00365 -0.0016 0.7534 0.0478
-0.750 0.0054 0.01224 0.00330 -0.0012 0.7410 0.0481
-0.500 0.0323 0.01202 0.00299 -0.0009 0.7287 0.0487
-0.250 0.0595 0.01185 0.00274 -0.0007 0.7168 0.0496
0.000 0.0868 0.01172 0.00253 -0.0005 0.7052 0.0511
0.250 0.1141 0.01161 0.00235 -0.0003 0.6940 0.0536
0.500 0.1413 0.01145 0.00222 -0.0001 0.6830 0.0676
0.750 0.1635 0.01015 0.00214 0.0004 0.6722 0.4577
1.250 0.2579 0.00906 0.00233 -0.0072 0.6477 1.0000
1.500 0.2841 0.00912 0.00232 -0.0068 0.6362 1.0000
1.750 0.3103 0.00920 0.00234 -0.0065 0.6249 1.0000
2.250 0.3628 0.00936 0.00241 -0.0058 0.6011 1.0000
2.500 0.3891 0.00945 0.00250 -0.0055 0.5889 1.0000
2.750 0.4154 0.00954 0.00259 -0.0051 0.5765 1.0000
3.000 0.4417 0.00965 0.00269 -0.0048 0.5638 1.0000
3.250 0.4677 0.00977 0.00279 -0.0044 0.5445 1.0000
3.500 0.4936 0.00991 0.00287 -0.0040 0.5170 1.0000
3.750 0.5196 0.01008 0.00298 -0.0037 0.4884 1.0000
4.000 0.5456 0.01028 0.00314 -0.0033 0.4611 1.0000
4.250 0.5715 0.01053 0.00333 -0.0030 0.4281 1.0000
4.500 0.5972 0.01086 0.00355 -0.0028 0.3813 1.0000
4.750 0.6230 0.01124 0.00382 -0.0026 0.3341 1.0000
5.000 0.6472 0.01215 0.00425 -0.0026 0.2218 1.0000
5.250 0.6697 0.01376 0.00513 -0.0029 0.0811 1.0000
5.500 0.6943 0.01464 0.00586 -0.0028 0.0495 1.0000
5.750 0.7192 0.01538 0.00664 -0.0026 0.0374 1.0000
6.000 0.7433 0.01629 0.00762 -0.0023 0.0291 1.0000
6.250 0.7681 0.01697 0.00848 -0.0020 0.0240 1.0000
6.500 0.7916 0.01795 0.00956 -0.0016 0.0203 1.0000
6.750 0.8135 0.01927 0.01097 -0.0010 0.0173 1.0000
7.000 0.8360 0.02041 0.01225 -0.0004 0.0157 1.0000
7.250 0.8577 0.02182 0.01380 0.0004 0.0145 1.0000
7.500 0.8793 0.02328 0.01540 0.0012 0.0134 1.0000
7.750 0.9012 0.02441 0.01664 0.0017 0.0120 1.0000
8.000 0.9199 0.02665 0.01911 0.0025 0.0107 1.0000
8.250 0.9381 0.02948 0.02226 0.0035 0.0103 1.0000
8.500 0.9564 0.03190 0.02504 0.0045 0.0100 1.0000
8.750 0.9720 0.03482 0.02837 0.0055 0.0097 1.0000
9.000 0.9839 0.03821 0.03220 0.0067 0.0095 1.0000
9.250 0.9913 0.04202 0.03645 0.0079 0.0094 1.0000
9.500 0.9945 0.04592 0.04076 0.0091 0.0092 1.0000
9.750 0.9934 0.04981 0.04501 0.0101 0.0089 1.0000
10.000 0.9871 0.05373 0.04923 0.0110 0.0087 1.0000
10.250 0.9733 0.05749 0.05322 0.0117 0.0086 1.0000
10.500 0.9566 0.06214 0.05808 0.0103 0.0086 1.0000
10.750 0.9384 0.06799 0.06411 0.0068 0.0087 1.0000
11.000 0.9193 0.07504 0.07132 0.0017 0.0089 1.0000
11.250 0.8993 0.08347 0.07987 -0.0047 0.0092 1.0000
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Polar data table (+)
Polar graphs
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