NACA M14 AIRFOIL (m14-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NACA M14 AIRFOIL (m14-il) Reynolds number: 100,000 Max Cl/Cd: 57.43 at α=6.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-m14-il-100000-n5.txt Download as CSV file: xf-m14-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3945 0.09275 0.08806 -0.0195 1.0000 0.0509
-7.500 -0.3918 0.08962 0.08499 -0.0215 1.0000 0.0526
-7.250 -0.3865 0.08624 0.08164 -0.0253 1.0000 0.0549
-7.000 -0.3779 0.08321 0.07856 -0.0361 1.0000 0.0575
-6.500 -0.3650 0.07519 0.07066 -0.0341 1.0000 0.0605
-6.250 -0.3571 0.07239 0.06787 -0.0339 1.0000 0.0630
-6.000 -0.3503 0.06952 0.06501 -0.0347 1.0000 0.0655
-5.750 -0.3263 0.06705 0.06216 -0.0431 0.9955 0.0709
-5.500 -0.3002 0.06095 0.05598 -0.0475 0.9858 0.0719
-5.250 -0.2796 0.05725 0.05239 -0.0480 0.9765 0.0754
-5.000 -0.2444 0.05379 0.04867 -0.0533 0.9643 0.0834
-4.750 -0.2068 0.04603 0.04037 -0.0585 0.9520 0.0562
-4.500 -0.1775 0.04226 0.03640 -0.0610 0.9396 0.0542
-4.250 -0.1452 0.03867 0.03240 -0.0634 0.9272 0.0552
-4.000 -0.1152 0.03538 0.02878 -0.0650 0.9148 0.0535
-3.750 -0.0852 0.03219 0.02518 -0.0661 0.9027 0.0518
-3.500 -0.0556 0.02934 0.02185 -0.0666 0.8906 0.0507
-3.250 -0.0275 0.02704 0.01911 -0.0667 0.8776 0.0502
-3.000 -0.0002 0.02556 0.01728 -0.0666 0.8645 0.0521
-2.750 0.0272 0.02410 0.01545 -0.0663 0.8518 0.0534
-2.500 0.0546 0.02266 0.01365 -0.0659 0.8397 0.0530
-2.250 0.0819 0.02143 0.01207 -0.0655 0.8279 0.0527
-2.000 0.1093 0.02037 0.01074 -0.0650 0.8167 0.0527
-1.750 0.1365 0.01948 0.00963 -0.0645 0.8051 0.0529
-1.500 0.1637 0.01871 0.00869 -0.0641 0.7934 0.0534
-1.250 0.1906 0.01805 0.00789 -0.0636 0.7824 0.0542
-1.000 0.2173 0.01748 0.00721 -0.0631 0.7721 0.0551
-0.750 0.2438 0.01701 0.00665 -0.0626 0.7611 0.0563
-0.500 0.2699 0.01660 0.00621 -0.0622 0.7501 0.0595
-0.250 0.2961 0.01631 0.00589 -0.0617 0.7398 0.0639
0.000 0.3225 0.01605 0.00554 -0.0611 0.7299 0.0671
0.250 0.3494 0.01587 0.00526 -0.0608 0.7189 0.0704
0.500 0.3765 0.01567 0.00503 -0.0604 0.7087 0.0773
0.750 0.4034 0.01547 0.00483 -0.0600 0.6994 0.1001
1.000 0.4281 0.01454 0.00483 -0.0599 0.6887 0.4037
1.500 0.5012 0.01346 0.00473 -0.0627 0.6690 1.0000
1.750 0.5271 0.01361 0.00478 -0.0623 0.6585 1.0000
2.000 0.5530 0.01377 0.00484 -0.0618 0.6488 1.0000
2.250 0.5789 0.01392 0.00488 -0.0613 0.6400 1.0000
2.500 0.6048 0.01409 0.00503 -0.0610 0.6296 1.0000
2.750 0.6308 0.01427 0.00515 -0.0606 0.6201 1.0000
3.000 0.6568 0.01444 0.00526 -0.0602 0.6113 1.0000
3.250 0.6827 0.01464 0.00548 -0.0598 0.6010 1.0000
3.500 0.7087 0.01484 0.00566 -0.0595 0.5917 1.0000
3.750 0.7347 0.01504 0.00584 -0.0591 0.5825 1.0000
4.000 0.7606 0.01526 0.00611 -0.0587 0.5723 1.0000
4.250 0.7865 0.01549 0.00637 -0.0584 0.5632 1.0000
4.500 0.8124 0.01572 0.00664 -0.0580 0.5536 1.0000
4.750 0.8381 0.01598 0.00697 -0.0576 0.5434 1.0000
5.000 0.8639 0.01623 0.00729 -0.0572 0.5341 1.0000
5.250 0.8895 0.01650 0.00764 -0.0569 0.5243 1.0000
5.500 0.9150 0.01679 0.00805 -0.0565 0.5138 1.0000
5.750 0.9405 0.01709 0.00846 -0.0561 0.5040 1.0000
6.000 0.9658 0.01736 0.00882 -0.0556 0.4929 1.0000
6.250 0.9903 0.01762 0.00918 -0.0549 0.4782 1.0000
6.500 1.0137 0.01785 0.00955 -0.0541 0.4572 1.0000
6.750 1.0344 0.01801 0.00967 -0.0527 0.4200 1.0000
7.000 1.0523 0.01844 0.00990 -0.0511 0.3624 1.0000
7.250 1.0698 0.01919 0.01044 -0.0496 0.3030 1.0000
7.500 1.0789 0.02090 0.01144 -0.0475 0.1818 1.0000
7.750 1.0812 0.02361 0.01331 -0.0448 0.0716 1.0000
8.000 1.0908 0.02546 0.01501 -0.0427 0.0482 1.0000
8.250 1.1021 0.02698 0.01662 -0.0407 0.0395 1.0000
8.500 1.1130 0.02840 0.01817 -0.0387 0.0349 1.0000
8.750 1.1202 0.03004 0.01993 -0.0363 0.0325 1.0000
9.000 1.1284 0.03148 0.02155 -0.0340 0.0304 1.0000
9.250 1.1336 0.03294 0.02314 -0.0315 0.0282 1.0000
9.500 1.1371 0.03457 0.02484 -0.0291 0.0264 1.0000
9.750 1.1376 0.03663 0.02694 -0.0267 0.0252 1.0000
10.000 1.1432 0.03854 0.02895 -0.0248 0.0245 1.0000
10.250 1.1521 0.04038 0.03092 -0.0230 0.0237 1.0000
10.500 1.1637 0.04226 0.03297 -0.0214 0.0229 1.0000
10.750 1.1761 0.04418 0.03505 -0.0200 0.0218 1.0000
11.000 1.1868 0.04618 0.03722 -0.0187 0.0206 1.0000
11.250 1.1957 0.04831 0.03950 -0.0175 0.0196 1.0000
11.500 1.2045 0.05066 0.04201 -0.0163 0.0189 1.0000
11.750 1.2126 0.05330 0.04483 -0.0152 0.0184 1.0000
12.000 1.2182 0.05623 0.04797 -0.0141 0.0181 1.0000
12.250 1.2206 0.05954 0.05150 -0.0131 0.0178 1.0000
12.500 1.2192 0.06318 0.05538 -0.0122 0.0176 1.0000
12.750 1.2139 0.06716 0.05967 -0.0117 0.0175 1.0000
13.000 1.2053 0.07138 0.06415 -0.0115 0.0174 1.0000
13.250 1.1940 0.07595 0.06898 -0.0119 0.0173 1.0000
13.500 1.1808 0.08079 0.07406 -0.0127 0.0173 1.0000
13.750 1.1662 0.08595 0.07946 -0.0141 0.0173 1.0000
14.000 1.1505 0.09143 0.08517 -0.0160 0.0172 1.0000
14.250 1.1342 0.09726 0.09121 -0.0186 0.0173 1.0000
14.500 1.1177 0.10341 0.09756 -0.0217 0.0173 1.0000
14.750 1.1005 0.11015 0.10451 -0.0255 0.0174 1.0000
15.000 1.0829 0.11768 0.11224 -0.0301 0.0175 1.0000
15.250 1.0622 0.12685 0.12158 -0.0362 0.0178 1.0000
15.500 1.0299 0.14098 0.13596 -0.0459 0.0185 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NACA M14 AIRFOIL (m14-il)