NACA M4 AIRFOIL (m4-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M4 AIRFOIL (m4-il) Reynolds number: 200,000 Max Cl/Cd: 57.27 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m4-il-200000.txt Download as CSV file: xf-m4-il-200000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6123 0.10167 0.09844 0.0275 1.0000 0.0309
-8.000 -0.6094 0.09784 0.09464 0.0253 1.0000 0.0315
-7.750 -0.6075 0.09391 0.09074 0.0225 1.0000 0.0322
-7.500 -0.6037 0.08981 0.08664 0.0192 1.0000 0.0329
-7.250 -0.5965 0.08548 0.08230 0.0155 1.0000 0.0339
-7.000 -0.5869 0.08101 0.07780 0.0116 1.0000 0.0353
-6.750 -0.5730 0.07641 0.07313 0.0068 1.0000 0.0371
-6.500 -0.5464 0.07247 0.06892 -0.0005 1.0000 0.0390
-6.250 -0.4915 0.05255 0.04932 -0.0058 1.0000 0.0410
-6.000 -0.4811 0.04883 0.04560 -0.0053 1.0000 0.0426
-5.750 -0.4682 0.04475 0.04145 -0.0061 1.0000 0.0445
-5.500 -0.4529 0.04043 0.03700 -0.0074 1.0000 0.0474
-5.250 -0.4237 0.03733 0.03326 -0.0095 1.0000 0.0524
-5.000 -0.4166 0.03040 0.02637 -0.0099 1.0000 0.0542
-4.750 -0.4000 0.02754 0.02354 -0.0098 1.0000 0.0568
-4.500 -0.3759 0.02419 0.01963 -0.0101 1.0000 0.0670
-4.250 -0.3588 0.02140 0.01700 -0.0102 1.0000 0.0703
-4.000 -0.3344 0.01875 0.01387 -0.0101 1.0000 0.0813
-3.750 -0.3139 0.01656 0.01177 -0.0103 1.0000 0.0855
-3.500 -0.2902 0.01456 0.00951 -0.0105 1.0000 0.0976
-3.250 -0.2673 0.01309 0.00794 -0.0107 1.0000 0.1124
-3.000 -0.2368 0.01109 0.00581 -0.0124 0.9847 0.1280
-2.750 -0.2153 0.02069 0.01355 -0.0063 1.0000 0.0664
-2.500 -0.1880 0.01893 0.01159 -0.0065 1.0000 0.0664
-2.250 -0.1475 0.01665 0.00911 -0.0091 0.9739 0.0644
-2.000 -0.1102 0.01537 0.00760 -0.0107 0.9501 0.0643
-1.750 -0.0813 0.01470 0.00677 -0.0103 0.9255 0.0652
-1.500 -0.0581 0.01395 0.00593 -0.0089 0.9027 0.0673
-1.250 -0.0355 0.01331 0.00524 -0.0073 0.8831 0.0680
-1.000 -0.0124 0.01277 0.00466 -0.0059 0.8646 0.0694
-0.750 0.0114 0.01237 0.00424 -0.0047 0.8480 0.0721
-0.500 0.0358 0.01206 0.00388 -0.0037 0.8325 0.0762
-0.250 0.0609 0.01177 0.00354 -0.0028 0.8177 0.0824
0.000 0.0863 0.01148 0.00327 -0.0020 0.8034 0.0997
0.250 0.1544 0.00894 0.00330 -0.0102 0.7909 1.0000
0.500 0.1796 0.00901 0.00321 -0.0094 0.7765 1.0000
0.750 0.2050 0.00909 0.00316 -0.0088 0.7626 1.0000
1.000 0.2305 0.00917 0.00313 -0.0082 0.7488 1.0000
1.250 0.2561 0.00926 0.00312 -0.0076 0.7353 1.0000
1.500 0.2817 0.00936 0.00313 -0.0071 0.7218 1.0000
1.750 0.3074 0.00946 0.00316 -0.0065 0.7085 1.0000
2.000 0.3332 0.00956 0.00323 -0.0060 0.6952 1.0000
2.250 0.3590 0.00968 0.00329 -0.0055 0.6819 1.0000
2.500 0.3848 0.00979 0.00337 -0.0050 0.6686 1.0000
2.750 0.4107 0.00991 0.00347 -0.0045 0.6551 1.0000
3.000 0.4366 0.01003 0.00359 -0.0040 0.6409 1.0000
3.250 0.4622 0.01011 0.00365 -0.0034 0.6226 1.0000
3.500 0.4874 0.01017 0.00365 -0.0027 0.6019 1.0000
3.750 0.5132 0.01023 0.00373 -0.0021 0.5797 1.0000
4.000 0.5389 0.01032 0.00382 -0.0015 0.5568 1.0000
4.250 0.5644 0.01039 0.00385 -0.0009 0.5237 1.0000
4.500 0.5900 0.01053 0.00394 -0.0004 0.4883 1.0000
4.750 0.6157 0.01075 0.00410 0.0001 0.4439 1.0000
5.000 0.6400 0.01136 0.00431 0.0005 0.3345 1.0000
5.250 0.6597 0.01424 0.00571 0.0001 0.0725 1.0000
5.500 0.6843 0.01530 0.00680 0.0004 0.0577 1.0000
5.750 0.7076 0.01657 0.00809 0.0008 0.0494 1.0000
6.000 0.7312 0.01763 0.00919 0.0014 0.0436 1.0000
6.250 0.7519 0.01975 0.01128 0.0023 0.0399 1.0000
6.500 0.7758 0.02111 0.01278 0.0031 0.0380 1.0000
6.750 0.8001 0.02227 0.01407 0.0038 0.0346 1.0000
7.000 0.8239 0.02406 0.01601 0.0046 0.0332 1.0000
7.250 0.8477 0.02622 0.01839 0.0055 0.0324 1.0000
7.500 0.8708 0.02910 0.02166 0.0066 0.0330 1.0000
7.750 0.8920 0.03226 0.02526 0.0077 0.0333 1.0000
8.000 0.9123 0.03434 0.02747 0.0081 0.0310 1.0000
8.250 0.9237 0.04138 0.03539 0.0099 0.0374 1.0000
11.750 0.8089 0.12854 0.12489 -0.0254 0.0491 1.0000
12.000 0.8052 0.13447 0.13080 -0.0289 0.0471 1.0000
12.250 0.8045 0.13947 0.13578 -0.0313 0.0454 1.0000
12.500 0.6709 0.13005 0.12659 -0.0079 0.0516 1.0000
12.750 0.6620 0.13498 0.13150 -0.0109 0.0513 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M4 AIRFOIL (m4-il)