NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: NACA M13 AIRFOIL (m13-il) Reynolds number: 1,000,000 Max Cl/Cd: 115.33 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-m13-il-1000000.txt Download as CSV file: xf-m13-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: NACA M13 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
------ -------- --------- --------- -------- -------- --------
0.500 0.3289 0.00778 0.00142 -0.0464 0.6098 0.0324
0.750 0.3571 0.00770 0.00130 -0.0465 0.6004 0.0331
1.000 0.3854 0.00763 0.00122 -0.0466 0.5908 0.0335
1.250 0.4136 0.00761 0.00116 -0.0466 0.5811 0.0339
1.750 0.4701 0.00759 0.00109 -0.0468 0.5585 0.0358
2.000 0.4983 0.00760 0.00108 -0.0469 0.5466 0.0364
2.250 0.5264 0.00762 0.00108 -0.0470 0.5353 0.0380
2.500 0.5545 0.00763 0.00113 -0.0471 0.5243 0.0623
2.750 0.5826 0.00769 0.00121 -0.0472 0.5132 0.0740
3.000 0.6107 0.00772 0.00126 -0.0474 0.5019 0.0832
3.250 0.6388 0.00777 0.00132 -0.0475 0.4914 0.0883
3.500 0.6667 0.00784 0.00140 -0.0476 0.4772 0.0973
3.750 0.6909 0.00607 0.00162 -0.0474 0.4602 1.0000
4.000 0.7183 0.00624 0.00172 -0.0474 0.4393 1.0000
4.250 0.7455 0.00647 0.00183 -0.0475 0.4118 1.0000
4.500 0.7727 0.00670 0.00196 -0.0475 0.3837 1.0000
4.750 0.7992 0.00708 0.00215 -0.0475 0.3347 1.0000
5.000 0.8233 0.00802 0.00257 -0.0474 0.2263 1.0000
5.250 0.8430 0.00996 0.00357 -0.0470 0.0353 1.0000
5.500 0.8689 0.01049 0.00402 -0.0468 0.0192 1.0000
5.750 0.8953 0.01085 0.00444 -0.0466 0.0166 1.0000
6.000 0.9215 0.01122 0.00484 -0.0465 0.0148 1.0000
6.250 0.9472 0.01170 0.00536 -0.0463 0.0132 1.0000
6.500 0.9699 0.01279 0.00661 -0.0456 0.0116 1.0000
6.750 0.9956 0.01317 0.00703 -0.0454 0.0112 1.0000
7.000 1.0207 0.01363 0.00754 -0.0451 0.0105 1.0000
7.250 1.0453 0.01413 0.00808 -0.0448 0.0096 1.0000
7.500 1.0690 0.01478 0.00877 -0.0443 0.0090 1.0000
7.750 1.0917 0.01553 0.00957 -0.0438 0.0085 1.0000
8.000 1.1118 0.01667 0.01079 -0.0429 0.0081 1.0000
8.250 1.1231 0.01941 0.01369 -0.0406 0.0076 1.0000
8.500 1.1468 0.01987 0.01423 -0.0402 0.0074 1.0000
8.750 1.1686 0.02070 0.01516 -0.0394 0.0071 1.0000
9.000 1.1884 0.02196 0.01653 -0.0384 0.0067 1.0000
9.250 1.2068 0.02364 0.01836 -0.0372 0.0065 1.0000
9.500 1.2250 0.02528 0.02014 -0.0360 0.0062 1.0000
9.750 1.2431 0.02669 0.02169 -0.0349 0.0060 1.0000
10.000 1.2600 0.02815 0.02330 -0.0338 0.0057 1.0000
10.250 1.2722 0.03084 0.02624 -0.0321 0.0057 1.0000
10.500 1.2761 0.03543 0.03124 -0.0294 0.0058 1.0000
10.750 1.2657 0.04209 0.03844 -0.0255 0.0062 1.0000
11.000 1.2541 0.04686 0.04355 -0.0222 0.0064 1.0000
11.250 1.2358 0.05060 0.04752 -0.0183 0.0066 1.0000
11.500 1.2174 0.05441 0.05154 -0.0161 0.0067 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M13 AIRFOIL (m13-il)