NACA M14 AIRFOIL (m14-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M14 AIRFOIL (m14-il) Reynolds number: 500,000 Max Cl/Cd: 98.13 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m14-il-500000-n5.txt Download as CSV file: xf-m14-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NACA M14 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4506 0.10241 0.10022 -0.0073 1.0000 0.0098
-9.000 -0.4510 0.09748 0.09530 -0.0097 1.0000 0.0102
-8.750 -0.4563 0.09136 0.08922 -0.0131 1.0000 0.0106
-8.500 -0.4504 0.08857 0.08644 -0.0148 1.0000 0.0108
-8.250 -0.4438 0.08602 0.08389 -0.0166 1.0000 0.0110
-8.000 -0.4382 0.08333 0.08124 -0.0186 1.0000 0.0113
-7.750 -0.4305 0.08005 0.07796 -0.0219 1.0000 0.0117
-7.500 -0.4191 0.07568 0.07359 -0.0270 0.9871 0.0123
-7.250 -0.4008 0.06915 0.06699 -0.0355 0.9525 0.0131
-7.000 -0.3893 0.05873 0.05636 -0.0455 0.9199 0.0143
-6.750 -0.3716 0.05687 0.05438 -0.0466 0.8951 0.0147
-6.500 -0.3532 0.05509 0.05248 -0.0475 0.8734 0.0151
-6.250 -0.3339 0.05219 0.04943 -0.0493 0.8537 0.0159
-6.000 -0.3137 0.04731 0.04433 -0.0521 0.8365 0.0169
-5.750 -0.2933 0.03632 0.03280 -0.0563 0.8229 0.0193
-5.500 -0.2706 0.03526 0.03160 -0.0565 0.8074 0.0198
-5.250 -0.2474 0.03363 0.02980 -0.0568 0.7933 0.0202
-5.000 -0.2238 0.03110 0.02703 -0.0571 0.7805 0.0210
-4.750 -0.1998 0.02785 0.02346 -0.0573 0.7687 0.0221
-4.500 -0.1759 0.02363 0.01876 -0.0571 0.7581 0.0234
-4.250 -0.1512 0.02069 0.01535 -0.0567 0.7474 0.0241
-4.000 -0.1251 0.01917 0.01350 -0.0565 0.7370 0.0247
-3.750 -0.0980 0.01868 0.01277 -0.0563 0.7267 0.0252
-3.500 -0.0724 0.01669 0.01038 -0.0562 0.7170 0.0258
-3.250 -0.0460 0.01522 0.00865 -0.0561 0.7076 0.0263
-3.000 -0.0191 0.01435 0.00760 -0.0560 0.6983 0.0267
-2.750 0.0083 0.01367 0.00678 -0.0559 0.6888 0.0270
-2.500 0.0358 0.01309 0.00608 -0.0559 0.6796 0.0273
-2.250 0.0632 0.01259 0.00545 -0.0558 0.6703 0.0276
-2.000 0.0909 0.01213 0.00489 -0.0557 0.6608 0.0278
-1.750 0.1185 0.01179 0.00447 -0.0557 0.6513 0.0284
-1.500 0.1461 0.01144 0.00405 -0.0556 0.6415 0.0288
-1.250 0.1737 0.01107 0.00360 -0.0556 0.6319 0.0289
-1.000 0.2013 0.01075 0.00321 -0.0555 0.6228 0.0290
-0.750 0.2289 0.01048 0.00288 -0.0554 0.6135 0.0292
-0.500 0.2566 0.01024 0.00260 -0.0554 0.6050 0.0296
0.000 0.3121 0.00988 0.00217 -0.0554 0.5878 0.0307
0.250 0.3399 0.00976 0.00201 -0.0554 0.5795 0.0315
0.500 0.3678 0.00966 0.00188 -0.0554 0.5711 0.0323
0.750 0.3957 0.00960 0.00179 -0.0554 0.5631 0.0330
1.000 0.4236 0.00955 0.00171 -0.0555 0.5545 0.0336
1.250 0.4516 0.00952 0.00166 -0.0555 0.5468 0.0341
1.500 0.4796 0.00951 0.00163 -0.0556 0.5388 0.0348
1.750 0.5076 0.00950 0.00159 -0.0556 0.5306 0.0374
2.250 0.5635 0.00952 0.00161 -0.0557 0.5148 0.0431
2.500 0.5908 0.00934 0.00171 -0.0558 0.5070 0.1607
3.000 0.6539 0.00759 0.00195 -0.0580 0.4901 1.0000
3.250 0.6807 0.00769 0.00204 -0.0578 0.4820 1.0000
3.500 0.7076 0.00780 0.00212 -0.0576 0.4737 1.0000
3.750 0.7343 0.00793 0.00222 -0.0575 0.4647 1.0000
4.000 0.7613 0.00805 0.00234 -0.0574 0.4562 1.0000
4.250 0.7881 0.00819 0.00247 -0.0573 0.4475 1.0000
4.500 0.8148 0.00835 0.00260 -0.0571 0.4349 1.0000
4.750 0.8410 0.00857 0.00275 -0.0569 0.4123 1.0000
5.000 0.8665 0.00887 0.00293 -0.0567 0.3816 1.0000
5.250 0.8920 0.00919 0.00314 -0.0564 0.3544 1.0000
5.500 0.9167 0.00961 0.00340 -0.0561 0.3196 1.0000
5.750 0.9412 0.01007 0.00371 -0.0558 0.2818 1.0000
6.000 0.9643 0.01071 0.00410 -0.0553 0.2291 1.0000
6.250 0.9831 0.01193 0.00483 -0.0544 0.1371 1.0000
6.500 1.0018 0.01313 0.00562 -0.0535 0.0621 1.0000
6.750 1.0234 0.01390 0.00625 -0.0528 0.0321 1.0000
7.000 1.0462 0.01450 0.00679 -0.0522 0.0199 1.0000
7.250 1.0697 0.01498 0.00730 -0.0517 0.0163 1.0000
7.500 1.0931 0.01545 0.00783 -0.0511 0.0146 1.0000
7.750 1.1154 0.01603 0.00844 -0.0505 0.0129 1.0000
8.000 1.1364 0.01674 0.00924 -0.0496 0.0114 1.0000
8.250 1.1584 0.01729 0.00986 -0.0489 0.0107 1.0000
8.500 1.1793 0.01792 0.01057 -0.0480 0.0101 1.0000
8.750 1.1993 0.01861 0.01132 -0.0471 0.0094 1.0000
9.000 1.2183 0.01934 0.01212 -0.0460 0.0089 1.0000
9.250 1.2356 0.02019 0.01305 -0.0448 0.0084 1.0000
9.500 1.2480 0.02140 0.01434 -0.0429 0.0078 1.0000
9.750 1.2651 0.02214 0.01516 -0.0416 0.0075 1.0000
10.000 1.2795 0.02303 0.01614 -0.0400 0.0072 1.0000
10.250 1.2899 0.02404 0.01725 -0.0378 0.0069 1.0000
10.500 1.2979 0.02511 0.01841 -0.0354 0.0066 1.0000
10.750 1.3055 0.02631 0.01971 -0.0332 0.0064 1.0000
11.000 1.3128 0.02764 0.02113 -0.0313 0.0062 1.0000
11.250 1.3199 0.02910 0.02269 -0.0297 0.0061 1.0000
11.500 1.3265 0.03069 0.02437 -0.0284 0.0059 1.0000
11.750 1.3321 0.03247 0.02625 -0.0272 0.0058 1.0000
12.000 1.3358 0.03453 0.02840 -0.0261 0.0056 1.0000
12.250 1.3365 0.03701 0.03099 -0.0251 0.0055 1.0000
12.500 1.3354 0.03976 0.03387 -0.0240 0.0054 1.0000
12.750 1.3390 0.04210 0.03634 -0.0234 0.0053 1.0000
13.000 1.3417 0.04456 0.03895 -0.0229 0.0052 1.0000
13.250 1.3436 0.04719 0.04172 -0.0225 0.0051 1.0000
13.500 1.3441 0.05001 0.04470 -0.0222 0.0050 1.0000
13.750 1.3436 0.05301 0.04786 -0.0221 0.0048 1.0000
14.000 1.3422 0.05619 0.05120 -0.0221 0.0047 1.0000
14.250 1.3394 0.05961 0.05476 -0.0223 0.0046 1.0000
14.500 1.3353 0.06326 0.05857 -0.0228 0.0045 1.0000
14.750 1.3302 0.06717 0.06263 -0.0234 0.0044 1.0000
15.000 1.3234 0.07146 0.06707 -0.0243 0.0043 1.0000
15.250 1.3154 0.07609 0.07187 -0.0255 0.0043 1.0000
15.500 1.3067 0.08102 0.07696 -0.0270 0.0042 1.0000
15.750 1.2967 0.08631 0.08240 -0.0288 0.0042 1.0000
16.000 1.2857 0.09195 0.08820 -0.0309 0.0042 1.0000
16.250 1.2733 0.09806 0.09448 -0.0334 0.0042 1.0000
16.500 1.2599 0.10457 0.10115 -0.0363 0.0041 1.0000
16.750 1.2454 0.11148 0.10821 -0.0395 0.0041 1.0000
17.000 1.2291 0.11905 0.11595 -0.0432 0.0041 1.0000
17.250 1.2121 0.12709 0.12415 -0.0473 0.0041 1.0000
17.500 1.1930 0.13597 0.13320 -0.0521 0.0041 1.0000
17.750 1.1707 0.14614 0.14354 -0.0578 0.0041 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M14 AIRFOIL (m14-il)