NACA M4 AIRFOIL (m4-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NACA M4 AIRFOIL (m4-il) Reynolds number: 1,000,000 Max Cl/Cd: 80.72 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m4-il-1000000.txt Download as CSV file: xf-m4-il-1000000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6640 0.13880 0.13728 0.0509 1.0000 0.0086
-10.500 -0.6596 0.13461 0.13308 0.0491 1.0000 0.0092
-6.500 -0.6448 0.01715 0.01315 0.0021 1.0000 0.0108
-6.250 -0.6166 0.01752 0.01358 0.0020 1.0000 0.0113
-6.000 -0.5886 0.01773 0.01381 0.0019 1.0000 0.0118
-5.750 -0.5631 0.01623 0.01206 0.0022 1.0000 0.0129
-5.500 -0.5364 0.01530 0.01095 0.0024 1.0000 0.0137
-5.250 -0.5105 0.01389 0.00931 0.0026 1.0000 0.0150
-5.000 -0.4817 0.01455 0.01007 0.0022 1.0000 0.0156
-4.750 -0.4514 0.01483 0.01036 0.0016 0.9626 0.0165
-4.500 -0.4292 0.01465 0.01001 0.0031 0.9164 0.0176
-4.250 -0.4050 0.01434 0.00952 0.0041 0.8867 0.0188
-4.000 -0.3787 0.01461 0.00966 0.0046 0.8612 0.0195
-3.750 -0.3541 0.01288 0.00765 0.0052 0.8400 0.0215
-3.500 -0.3272 0.01287 0.00756 0.0053 0.8189 0.0227
-3.250 -0.3001 0.01258 0.00716 0.0055 0.8005 0.0241
-3.000 -0.2728 0.01215 0.00659 0.0057 0.7834 0.0253
-2.750 -0.2451 0.01194 0.00627 0.0058 0.7679 0.0266
-2.500 -0.2173 0.01180 0.00604 0.0059 0.7534 0.0273
-2.250 -0.1892 0.01185 0.00600 0.0059 0.7402 0.0280
-2.000 -0.1634 0.00988 0.00386 0.0064 0.7287 0.0302
-1.750 -0.1360 0.00949 0.00341 0.0065 0.7170 0.0316
-1.500 -0.1084 0.00916 0.00303 0.0066 0.7058 0.0324
-1.250 -0.0808 0.00886 0.00268 0.0067 0.6952 0.0329
-1.000 -0.0531 0.00861 0.00237 0.0068 0.6851 0.0333
-0.750 -0.0252 0.00838 0.00208 0.0068 0.6747 0.0336
-0.500 0.0027 0.00819 0.00185 0.0068 0.6645 0.0344
-0.250 0.0308 0.00803 0.00165 0.0068 0.6546 0.0354
0.000 0.0589 0.00790 0.00147 0.0068 0.6446 0.0366
0.250 0.0871 0.00780 0.00133 0.0068 0.6343 0.0383
0.500 0.1154 0.00773 0.00124 0.0067 0.6243 0.0395
0.750 0.1437 0.00765 0.00111 0.0067 0.6140 0.0439
1.250 0.1908 0.00533 0.00102 0.0075 0.5936 0.8077
1.500 0.2281 0.00501 0.00115 0.0060 0.5820 0.9889
1.750 0.2890 0.00527 0.00134 -0.0015 0.5656 1.0000
2.000 0.3160 0.00534 0.00132 -0.0013 0.5430 1.0000
2.250 0.3428 0.00541 0.00131 -0.0011 0.5222 1.0000
2.500 0.3696 0.00548 0.00132 -0.0009 0.5059 1.0000
2.750 0.3963 0.00557 0.00134 -0.0007 0.4862 1.0000
3.000 0.4229 0.00569 0.00139 -0.0006 0.4622 1.0000
3.250 0.4493 0.00585 0.00144 -0.0004 0.4329 1.0000
3.500 0.4757 0.00599 0.00151 -0.0001 0.4077 1.0000
3.750 0.5021 0.00622 0.00160 0.0000 0.3697 1.0000
4.000 0.5282 0.00657 0.00174 0.0001 0.3130 1.0000
4.250 0.5540 0.00724 0.00200 0.0000 0.2147 1.0000
4.500 0.5793 0.00830 0.00248 -0.0002 0.0838 1.0000
4.750 0.6050 0.00887 0.00284 -0.0001 0.0392 1.0000
5.000 0.6311 0.00916 0.00311 0.0001 0.0331 1.0000
5.250 0.6572 0.00948 0.00347 0.0004 0.0276 1.0000
5.500 0.6833 0.00973 0.00372 0.0006 0.0236 1.0000
5.750 0.7093 0.01024 0.00427 0.0008 0.0192 1.0000
6.000 0.7355 0.01050 0.00455 0.0010 0.0175 1.0000
6.250 0.7616 0.01083 0.00491 0.0011 0.0156 1.0000
6.500 0.7864 0.01186 0.00606 0.0013 0.0133 1.0000
6.750 0.8127 0.01207 0.00629 0.0015 0.0126 1.0000
7.000 0.8386 0.01247 0.00672 0.0016 0.0117 1.0000
7.250 0.8642 0.01296 0.00725 0.0018 0.0108 1.0000
7.500 0.8894 0.01352 0.00786 0.0020 0.0101 1.0000
7.750 0.9125 0.01468 0.00912 0.0023 0.0093 1.0000
8.000 0.9331 0.01646 0.01108 0.0030 0.0088 1.0000
8.250 0.9577 0.01703 0.01173 0.0033 0.0085 1.0000
8.500 0.9811 0.01796 0.01277 0.0037 0.0080 1.0000
8.750 1.0042 0.01886 0.01377 0.0041 0.0076 1.0000
9.000 1.0276 0.01959 0.01460 0.0045 0.0071 1.0000
9.250 1.0500 0.02049 0.01559 0.0048 0.0068 1.0000
9.500 1.0714 0.02158 0.01679 0.0053 0.0065 1.0000
9.750 1.0911 0.02298 0.01833 0.0059 0.0063 1.0000
10.000 1.1072 0.02509 0.02065 0.0067 0.0061 1.0000
10.250 1.1165 0.02851 0.02442 0.0081 0.0059 1.0000
10.500 1.1157 0.03336 0.02975 0.0098 0.0058 1.0000
10.750 1.1032 0.03908 0.03595 0.0118 0.0058 1.0000
11.000 1.0865 0.04370 0.04089 0.0136 0.0058 1.0000
11.250 1.0671 0.04789 0.04530 0.0140 0.0057 1.0000
11.500 1.0515 0.05274 0.05034 0.0120 0.0058 1.0000
11.750 1.0351 0.05860 0.05637 0.0083 0.0058 1.0000
12.000 1.0209 0.06484 0.06276 0.0038 0.0058 1.0000
12.250 1.0077 0.07162 0.06966 -0.0013 0.0058 1.0000
12.500 0.9992 0.07795 0.07609 -0.0059 0.0058 1.0000
12.750 0.9863 0.08585 0.08409 -0.0115 0.0059 1.0000
13.000 0.9744 0.09417 0.09250 -0.0169 0.0059 1.0000
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Polar data table (+)
Polar graphs
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