NACA M4 AIRFOIL (m4-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NACA M4 AIRFOIL (m4-il) Reynolds number: 500,000 Max Cl/Cd: 66.76 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m4-il-500000-n5.txt Download as CSV file: xf-m4-il-500000-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 = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.6353 0.09874 0.09672 0.0303 1.0000 0.0054
-8.250 -0.6349 0.09442 0.09243 0.0275 1.0000 0.0055
-8.000 -0.6354 0.09007 0.08810 0.0240 1.0000 0.0057
-7.750 -0.6327 0.08501 0.08304 0.0194 1.0000 0.0059
-7.500 -0.6266 0.07943 0.07744 0.0146 1.0000 0.0062
-7.250 -0.6186 0.07341 0.07137 0.0099 1.0000 0.0063
-7.000 -0.6084 0.06691 0.06479 0.0054 1.0000 0.0063
-6.750 -0.5966 0.05953 0.05728 0.0013 1.0000 0.0065
-6.500 -0.5849 0.04969 0.04717 -0.0024 1.0000 0.0067
-6.250 -0.5970 0.01914 0.01448 -0.0023 1.0000 0.0076
-6.000 -0.5716 0.01762 0.01273 -0.0022 1.0000 0.0083
-5.750 -0.5456 0.01627 0.01115 -0.0020 1.0000 0.0094
-5.500 -0.5197 0.01466 0.00917 -0.0017 1.0000 0.0106
-5.250 -0.4901 0.01401 0.00840 -0.0022 0.9603 0.0125
-5.000 -0.4639 0.01424 0.00859 -0.0016 0.9168 0.0141
-4.750 -0.4403 0.01397 0.00812 -0.0005 0.8864 0.0163
-4.500 -0.4155 0.01375 0.00775 0.0003 0.8611 0.0179
-4.250 -0.3893 0.01406 0.00801 0.0007 0.8382 0.0193
-4.000 -0.3628 0.01411 0.00794 0.0010 0.8174 0.0211
-3.500 -0.3091 0.01360 0.00710 0.0015 0.7823 0.0254
-3.250 -0.2818 0.01342 0.00678 0.0017 0.7669 0.0269
-3.000 -0.2544 0.01323 0.00645 0.0018 0.7529 0.0278
-2.750 -0.2268 0.01317 0.00628 0.0019 0.7398 0.0284
-2.500 -0.2006 0.01182 0.00475 0.0023 0.7281 0.0305
-2.250 -0.1735 0.01135 0.00419 0.0024 0.7165 0.0318
-2.000 -0.1461 0.01102 0.00379 0.0025 0.7056 0.0328
-1.750 -0.1187 0.01073 0.00343 0.0026 0.6950 0.0336
-1.250 -0.0637 0.01021 0.00277 0.0028 0.6743 0.0345
-1.000 -0.0360 0.00999 0.00250 0.0029 0.6646 0.0350
-0.750 -0.0083 0.00983 0.00228 0.0029 0.6547 0.0359
-0.500 0.0195 0.00965 0.00205 0.0029 0.6446 0.0362
-0.250 0.0474 0.00949 0.00185 0.0029 0.6348 0.0361
0.000 0.0753 0.00937 0.00168 0.0029 0.6250 0.0362
0.250 0.1034 0.00927 0.00154 0.0029 0.6148 0.0363
0.500 0.1315 0.00920 0.00142 0.0029 0.6045 0.0367
0.750 0.1596 0.00914 0.00134 0.0028 0.5944 0.0372
1.000 0.1877 0.00911 0.00127 0.0028 0.5838 0.0383
1.250 0.2159 0.00908 0.00122 0.0027 0.5727 0.0405
1.500 0.2434 0.00878 0.00124 0.0026 0.5618 0.1551
1.750 0.2675 0.00650 0.00131 0.0032 0.5515 0.9688
2.000 0.3178 0.00666 0.00143 -0.0019 0.5364 0.9958
2.250 0.3548 0.00681 0.00149 -0.0041 0.5134 1.0000
2.500 0.3813 0.00693 0.00152 -0.0038 0.4871 1.0000
2.750 0.4077 0.00707 0.00156 -0.0036 0.4623 1.0000
3.000 0.4342 0.00721 0.00163 -0.0034 0.4421 1.0000
3.250 0.4606 0.00734 0.00172 -0.0032 0.4227 1.0000
3.500 0.4870 0.00753 0.00182 -0.0030 0.3949 1.0000
3.750 0.5133 0.00776 0.00195 -0.0028 0.3619 1.0000
4.000 0.5394 0.00808 0.00211 -0.0027 0.3183 1.0000
4.250 0.5653 0.00851 0.00234 -0.0027 0.2608 1.0000
4.500 0.5904 0.00938 0.00276 -0.0028 0.1568 1.0000
4.750 0.6154 0.01029 0.00324 -0.0029 0.0658 1.0000
5.000 0.6412 0.01076 0.00361 -0.0028 0.0425 1.0000
5.250 0.6673 0.01108 0.00395 -0.0027 0.0351 1.0000
5.500 0.6934 0.01143 0.00432 -0.0025 0.0286 1.0000
5.750 0.7195 0.01175 0.00467 -0.0024 0.0221 1.0000
6.000 0.7455 0.01218 0.00512 -0.0022 0.0162 1.0000
6.250 0.7713 0.01263 0.00560 -0.0021 0.0130 1.0000
6.500 0.7964 0.01343 0.00650 -0.0018 0.0104 1.0000
6.750 0.8219 0.01392 0.00707 -0.0017 0.0095 1.0000
7.000 0.8472 0.01442 0.00762 -0.0015 0.0083 1.0000
7.250 0.8720 0.01503 0.00830 -0.0013 0.0075 1.0000
7.500 0.8949 0.01615 0.00952 -0.0009 0.0068 1.0000
7.750 0.9182 0.01705 0.01058 -0.0006 0.0065 1.0000
8.000 0.9411 0.01804 0.01170 -0.0001 0.0061 1.0000
8.250 0.9633 0.01913 0.01293 0.0004 0.0057 1.0000
8.500 0.9859 0.02002 0.01393 0.0009 0.0053 1.0000
8.750 1.0087 0.02074 0.01472 0.0011 0.0049 1.0000
9.000 1.0300 0.02175 0.01582 0.0016 0.0045 1.0000
9.250 1.0469 0.02376 0.01804 0.0025 0.0043 1.0000
9.500 1.0654 0.02537 0.01988 0.0033 0.0042 1.0000
9.750 1.0818 0.02733 0.02213 0.0042 0.0040 1.0000
10.000 1.0951 0.02978 0.02494 0.0053 0.0039 1.0000
10.250 1.1040 0.03278 0.02831 0.0066 0.0037 1.0000
10.500 1.1066 0.03643 0.03238 0.0082 0.0036 1.0000
10.750 1.1007 0.04074 0.03709 0.0098 0.0035 1.0000
11.000 1.0844 0.04501 0.04169 0.0117 0.0035 1.0000
11.250 1.0641 0.05008 0.04703 0.0112 0.0035 1.0000
11.500 1.0426 0.05634 0.05354 0.0084 0.0035 1.0000
11.750 1.0210 0.06358 0.06099 0.0038 0.0035 1.0000
12.000 0.9990 0.07194 0.06952 -0.0021 0.0035 1.0000
12.250 0.9774 0.08136 0.07904 -0.0090 0.0036 1.0000
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Polar data table (+)
Polar graphs
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