NACA M5 AIRFOIL (m5-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M5 AIRFOIL (m5-il) Reynolds number: 200,000 Max Cl/Cd: 57.23 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m5-il-200000-n5.txt Download as CSV file: xf-m5-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M5 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.6403 0.09004 0.08685 0.0133 1.0000 0.0191
-8.500 -0.6477 0.08439 0.08123 0.0085 1.0000 0.0192
-8.250 -0.6516 0.07858 0.07540 0.0043 1.0000 0.0194
-8.000 -0.6521 0.07236 0.06911 0.0003 1.0000 0.0198
-7.750 -0.6506 0.06568 0.06230 -0.0032 1.0000 0.0204
-7.500 -0.6484 0.05761 0.05400 -0.0061 1.0000 0.0217
-7.250 -0.6571 0.04218 0.03783 -0.0078 1.0000 0.0240
-7.000 -0.6358 0.04169 0.03727 -0.0078 1.0000 0.0249
-6.750 -0.6159 0.03976 0.03517 -0.0077 1.0000 0.0261
-6.500 -0.5973 0.03630 0.03139 -0.0074 1.0000 0.0286
-6.250 -0.5827 0.02949 0.02359 -0.0060 1.0000 0.0319
-6.000 -0.5612 0.02733 0.02116 -0.0057 1.0000 0.0336
-5.750 -0.5360 0.02675 0.02053 -0.0058 1.0000 0.0352
-5.500 -0.5111 0.02533 0.01889 -0.0056 1.0000 0.0367
-5.250 -0.4857 0.02378 0.01704 -0.0054 1.0000 0.0387
-5.000 -0.4595 0.02233 0.01528 -0.0052 1.0000 0.0406
-4.750 -0.4293 0.02090 0.01350 -0.0058 0.9726 0.0416
-4.500 -0.3980 0.01981 0.01211 -0.0065 0.9447 0.0422
-4.250 -0.3706 0.01898 0.01105 -0.0063 0.9198 0.0427
-4.000 -0.3463 0.01760 0.00950 -0.0055 0.8970 0.0443
-3.750 -0.3220 0.01696 0.00874 -0.0046 0.8763 0.0457
-3.500 -0.2973 0.01631 0.00796 -0.0037 0.8572 0.0465
-3.250 -0.2724 0.01570 0.00724 -0.0029 0.8396 0.0474
-3.000 -0.2473 0.01516 0.00659 -0.0021 0.8233 0.0485
-2.750 -0.2219 0.01467 0.00599 -0.0014 0.8081 0.0498
-2.500 -0.1961 0.01424 0.00546 -0.0008 0.7935 0.0515
-2.250 -0.1700 0.01386 0.00498 -0.0003 0.7797 0.0533
-2.000 -0.1436 0.01355 0.00457 0.0001 0.7663 0.0550
-1.750 -0.1174 0.01316 0.00412 0.0005 0.7537 0.0577
-1.500 -0.0909 0.01287 0.00380 0.0008 0.7414 0.0615
-1.250 -0.0640 0.01268 0.00354 0.0011 0.7297 0.0674
-1.000 -0.0370 0.01244 0.00330 0.0013 0.7179 0.0760
-0.750 -0.0099 0.01222 0.00310 0.0015 0.7062 0.0898
-0.500 0.0171 0.01197 0.00293 0.0017 0.6951 0.1204
0.000 0.0568 0.00954 0.00262 0.0044 0.6746 0.7170
0.500 0.1892 0.00988 0.00329 -0.0103 0.6502 1.0000
0.750 0.2155 0.00990 0.00320 -0.0100 0.6400 1.0000
1.000 0.2419 0.00992 0.00314 -0.0097 0.6297 1.0000
1.250 0.2682 0.00994 0.00310 -0.0095 0.6192 1.0000
1.500 0.2945 0.00998 0.00308 -0.0092 0.6092 1.0000
1.750 0.3206 0.01003 0.00306 -0.0089 0.5993 1.0000
2.000 0.3469 0.01007 0.00308 -0.0086 0.5887 1.0000
2.250 0.3731 0.01013 0.00311 -0.0083 0.5785 1.0000
2.500 0.3992 0.01021 0.00314 -0.0079 0.5685 1.0000
2.750 0.4253 0.01028 0.00320 -0.0076 0.5576 1.0000
3.000 0.4514 0.01036 0.00329 -0.0073 0.5467 1.0000
3.250 0.4774 0.01045 0.00338 -0.0069 0.5358 1.0000
3.500 0.5034 0.01056 0.00348 -0.0066 0.5242 1.0000
3.750 0.5292 0.01067 0.00360 -0.0062 0.5101 1.0000
4.000 0.5550 0.01079 0.00372 -0.0058 0.4923 1.0000
4.250 0.5806 0.01093 0.00383 -0.0054 0.4700 1.0000
4.500 0.6061 0.01111 0.00398 -0.0050 0.4435 1.0000
4.750 0.6314 0.01132 0.00413 -0.0046 0.4144 1.0000
5.000 0.6567 0.01157 0.00434 -0.0043 0.3874 1.0000
5.250 0.6816 0.01191 0.00458 -0.0039 0.3513 1.0000
5.500 0.7062 0.01234 0.00489 -0.0035 0.3083 1.0000
5.750 0.7298 0.01300 0.00529 -0.0033 0.2409 1.0000
6.000 0.7509 0.01427 0.00599 -0.0031 0.1372 1.0000
6.250 0.7725 0.01541 0.00677 -0.0028 0.0721 1.0000
6.500 0.7953 0.01623 0.00753 -0.0023 0.0512 1.0000
6.750 0.8183 0.01699 0.00832 -0.0018 0.0403 1.0000
7.000 0.8407 0.01781 0.00916 -0.0013 0.0324 1.0000
7.250 0.8633 0.01855 0.00999 -0.0007 0.0274 1.0000
7.500 0.8842 0.01957 0.01108 -0.0001 0.0242 1.0000
7.750 0.9059 0.02038 0.01200 0.0006 0.0215 1.0000
8.000 0.9268 0.02127 0.01301 0.0013 0.0194 1.0000
8.250 0.9458 0.02242 0.01423 0.0022 0.0179 1.0000
8.500 0.9608 0.02420 0.01608 0.0035 0.0167 1.0000
8.750 0.9799 0.02532 0.01734 0.0045 0.0159 1.0000
9.000 0.9987 0.02647 0.01864 0.0054 0.0147 1.0000
9.250 1.0164 0.02775 0.02006 0.0064 0.0138 1.0000
9.500 1.0330 0.02925 0.02170 0.0075 0.0132 1.0000
9.750 1.0487 0.03082 0.02346 0.0086 0.0126 1.0000
10.000 1.0631 0.03252 0.02531 0.0098 0.0122 1.0000
10.250 1.0754 0.03447 0.02742 0.0110 0.0118 1.0000
10.500 1.0845 0.03692 0.03008 0.0125 0.0114 1.0000
10.750 1.0877 0.04013 0.03357 0.0141 0.0111 1.0000
11.000 1.0910 0.04222 0.03596 0.0159 0.0108 1.0000
11.250 1.0884 0.04459 0.03860 0.0179 0.0105 1.0000
11.500 1.0833 0.04746 0.04174 0.0191 0.0102 1.0000
11.750 1.0759 0.05087 0.04541 0.0194 0.0100 1.0000
12.000 1.0659 0.05486 0.04966 0.0189 0.0099 1.0000
12.250 1.0533 0.05952 0.05457 0.0175 0.0098 1.0000
12.500 1.0385 0.06485 0.06013 0.0153 0.0098 1.0000
12.750 1.0217 0.07087 0.06638 0.0122 0.0098 1.0000
13.000 1.0031 0.07765 0.07335 0.0083 0.0098 1.0000
13.250 0.9828 0.08529 0.08118 0.0036 0.0099 1.0000
13.500 0.9610 0.09389 0.08994 -0.0018 0.0100 1.0000
13.750 0.9373 0.10374 0.09994 -0.0080 0.0102 1.0000
14.000 0.9115 0.11526 0.11156 -0.0149 0.0105 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M5 AIRFOIL (m5-il)