NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M13 AIRFOIL (m13-il) Reynolds number: 200,000 Max Cl/Cd: 73.39 at α=5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m13-il-200000-n5.txt Download as CSV file: xf-m13-il-200000-n5.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 = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4491 0.10418 0.10085 0.0024 1.0000 0.0202
-8.000 -0.4431 0.10101 0.09771 -0.0001 1.0000 0.0213
-7.500 -0.4390 0.09188 0.08866 -0.0063 1.0000 0.0144
-7.250 -0.4312 0.08800 0.08479 -0.0088 1.0000 0.0142
-7.000 -0.4195 0.08351 0.08032 -0.0148 1.0000 0.0145
-6.750 -0.4040 0.08254 0.07934 -0.0161 1.0000 0.0217
-6.500 -0.3902 0.07834 0.07513 -0.0201 1.0000 0.0217
-6.250 -0.3818 0.07403 0.07084 -0.0210 1.0000 0.0190
-6.000 -0.3655 0.06918 0.06596 -0.0255 1.0000 0.0175
-5.750 -0.3466 0.06420 0.06093 -0.0300 1.0000 0.0167
-5.500 -0.3263 0.05960 0.05626 -0.0338 1.0000 0.0165
-5.250 -0.3025 0.05523 0.05180 -0.0376 0.9942 0.0168
-5.000 -0.2682 0.05112 0.04755 -0.0429 0.9703 0.0190
-4.750 -0.2316 0.04585 0.04202 -0.0483 0.9483 0.0202
-4.500 -0.1962 0.04035 0.03619 -0.0524 0.9277 0.0210
-4.250 -0.1647 0.03558 0.03106 -0.0548 0.9083 0.0236
-4.000 -0.1409 0.03434 0.02965 -0.0551 0.8875 0.0261
-3.750 -0.1131 0.03053 0.02545 -0.0556 0.8700 0.0272
-3.500 -0.0837 0.02626 0.02057 -0.0556 0.8544 0.0304
-3.250 -0.0572 0.02246 0.01614 -0.0554 0.8395 0.0318
-3.000 -0.0316 0.02080 0.01416 -0.0552 0.8240 0.0331
-2.750 -0.0058 0.02012 0.01330 -0.0550 0.8090 0.0356
-2.500 0.0214 0.01819 0.01090 -0.0546 0.7956 0.0363
-2.250 0.0487 0.01668 0.00897 -0.0542 0.7824 0.0369
-2.000 0.0760 0.01559 0.00757 -0.0538 0.7697 0.0376
-1.750 0.1031 0.01477 0.00650 -0.0535 0.7574 0.0382
-1.500 0.1303 0.01421 0.00573 -0.0532 0.7457 0.0394
-1.250 0.1573 0.01370 0.00506 -0.0530 0.7346 0.0403
-1.000 0.1843 0.01323 0.00446 -0.0527 0.7240 0.0401
-0.750 0.2114 0.01279 0.00392 -0.0525 0.7131 0.0401
-0.500 0.2385 0.01245 0.00351 -0.0523 0.7028 0.0400
-0.250 0.2657 0.01219 0.00318 -0.0521 0.6929 0.0400
0.000 0.2930 0.01200 0.00291 -0.0520 0.6828 0.0401
0.500 0.3480 0.01176 0.00254 -0.0518 0.6634 0.0408
0.750 0.3755 0.01169 0.00243 -0.0517 0.6536 0.0418
1.000 0.4031 0.01164 0.00237 -0.0516 0.6438 0.0435
1.250 0.4306 0.01161 0.00237 -0.0515 0.6344 0.0479
1.500 0.4582 0.01162 0.00238 -0.0515 0.6246 0.0667
1.750 0.4858 0.01164 0.00241 -0.0514 0.6146 0.0793
2.000 0.5133 0.01167 0.00243 -0.0514 0.6050 0.0890
2.500 0.5623 0.01028 0.00264 -0.0506 0.5853 0.7268
2.750 0.5972 0.00990 0.00272 -0.0517 0.5745 1.0000
3.000 0.6242 0.01005 0.00281 -0.0515 0.5643 1.0000
3.250 0.6513 0.01018 0.00293 -0.0514 0.5530 1.0000
3.500 0.6784 0.01033 0.00309 -0.0513 0.5415 1.0000
3.750 0.7054 0.01048 0.00324 -0.0512 0.5301 1.0000
4.000 0.7323 0.01065 0.00342 -0.0510 0.5194 1.0000
4.250 0.7592 0.01083 0.00364 -0.0509 0.5086 1.0000
4.500 0.7862 0.01101 0.00388 -0.0508 0.4979 1.0000
4.750 0.8130 0.01120 0.00413 -0.0507 0.4861 1.0000
5.000 0.8388 0.01143 0.00433 -0.0503 0.4593 1.0000
5.250 0.8635 0.01178 0.00456 -0.0498 0.4152 1.0000
5.500 0.8864 0.01240 0.00486 -0.0493 0.3355 1.0000
5.750 0.9005 0.01469 0.00592 -0.0484 0.1205 1.0000
6.000 0.9183 0.01650 0.00721 -0.0475 0.0293 1.0000
6.250 0.9411 0.01738 0.00817 -0.0468 0.0225 1.0000
6.500 0.9642 0.01813 0.00910 -0.0462 0.0187 1.0000
6.750 0.9858 0.01909 0.01023 -0.0454 0.0167 1.0000
7.000 1.0050 0.02035 0.01163 -0.0444 0.0151 1.0000
7.250 1.0181 0.02246 0.01385 -0.0425 0.0131 1.0000
7.500 1.0383 0.02345 0.01496 -0.0416 0.0123 1.0000
7.750 1.0562 0.02491 0.01654 -0.0402 0.0117 1.0000
8.000 1.0741 0.02653 0.01833 -0.0388 0.0111 1.0000
8.250 1.0926 0.02832 0.02026 -0.0375 0.0107 1.0000
8.500 1.1115 0.03024 0.02235 -0.0363 0.0102 1.0000
8.750 1.1298 0.03180 0.02405 -0.0352 0.0095 1.0000
9.000 1.1463 0.03328 0.02566 -0.0343 0.0087 1.0000
9.250 1.1603 0.03597 0.02850 -0.0330 0.0081 1.0000
9.500 1.1719 0.03962 0.03247 -0.0314 0.0079 1.0000
9.750 1.1820 0.04225 0.03545 -0.0297 0.0078 1.0000
10.000 1.1882 0.04518 0.03874 -0.0277 0.0077 1.0000
10.250 1.1898 0.04833 0.04223 -0.0255 0.0076 1.0000
10.500 1.1861 0.05149 0.04570 -0.0230 0.0076 1.0000
10.750 1.1768 0.05461 0.04908 -0.0203 0.0076 1.0000
11.000 1.1654 0.05806 0.05279 -0.0185 0.0076 1.0000
11.250 1.1525 0.06192 0.05694 -0.0177 0.0076 1.0000
11.500 1.1384 0.06625 0.06149 -0.0179 0.0076 1.0000
11.750 1.1235 0.07105 0.06650 -0.0190 0.0077 1.0000
12.000 1.1078 0.07635 0.07198 -0.0208 0.0077 1.0000
12.250 1.0914 0.08218 0.07799 -0.0234 0.0077 1.0000
12.500 1.0746 0.08851 0.08443 -0.0267 0.0077 1.0000
12.750 1.0579 0.09534 0.09139 -0.0305 0.0078 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M13 AIRFOIL (m13-il)