NACA 25112 (naca25112-jf)
NACA 25112 - NACA 25112 5 digit reflex airfoil
| Details | Dat file | Parser | |
|
(naca25112-jf) NACA 25112 NACA 25112 5 digit reflex airfoil Max thickness 12% at 29.5% chord. Max camber 2% at 27.2% chord Source Javafoil generated Source dat file The dat file is in Selig format |
NACA 25112 0.99998910 0.00125995 0.99930326 0.00135018 0.99724778 0.00162080 0.99382887 0.00207156 0.98905678 0.00270208 0.98294579 0.00351178 0.97551412 0.00449986 0.96678377 0.00566530 0.95678046 0.00700680 0.94553345 0.00852271 0.93307538 0.01021099 0.91944213 0.01206912 0.90467264 0.01409400 0.88880874 0.01628186 0.87189499 0.01862815 0.85397851 0.02112738 0.83510878 0.02377304 0.81533754 0.02655743 0.79471859 0.02947152 0.77330766 0.03250484 0.75116223 0.03564536 0.72834145 0.03887939 0.70490595 0.04219149 0.68091774 0.04556448 0.65644007 0.04897934 0.63153732 0.05241534 0.60627487 0.05585006 0.58071903 0.05925955 0.55493688 0.06261851 0.52899619 0.06590052 0.50296532 0.06907829 0.47691311 0.07212407 0.45090876 0.07500994 0.42499472 0.07770679 0.39916635 0.08016081 0.37348845 0.08230129 0.34803325 0.08406170 0.32287644 0.08538169 0.29809696 0.08620863 0.27377671 0.08649907 0.25000000 0.08621990 0.22685286 0.08534941 0.20442216 0.08387798 0.18279457 0.08180842 0.16205540 0.07915600 0.14228736 0.07594802 0.12356937 0.07222304 0.10597530 0.06802970 0.08957302 0.06342521 0.07442341 0.05847353 0.06057978 0.05324341 0.04808739 0.04780621 0.03698327 0.04223375 0.02729620 0.03659618 0.01904695 0.03095991 0.01224859 0.02538575 0.00690697 0.01992730 0.00302119 0.01462953 0.00058415 0.00952771 -0.00041692 0.00464660 0.00000000 0.00000000 0.00178739 -0.00430862 0.00489395 -0.00818444 0.00929047 -0.01163946 0.01494543 -0.01469111 0.02182558 -0.01736166 0.02989654 -0.01967744 0.03912337 -0.02166809 0.04947127 -0.02336572 0.06090608 -0.02480405 0.07339482 -0.02601768 0.08690603 -0.02704127 0.10140999 -0.02790881 0.11687873 -0.02865302 0.13328581 -0.02930468 0.15060586 -0.02989205 0.16881400 -0.03044028 0.18788504 -0.03097086 0.20779259 -0.03150106 0.22850810 -0.03204336 0.25000000 -0.03260495 0.27223279 -0.03318723 0.29516640 -0.03378549 0.31875562 -0.03438866 0.34294976 -0.03497924 0.36769251 -0.03553345 0.39292196 -0.03602161 0.41857082 -0.03640876 0.44456277 -0.03665542 0.47075094 -0.03673113 0.49703468 -0.03663598 0.52333977 -0.03637474 0.54959159 -0.03595239 0.57571543 -0.03537425 0.60163682 -0.03464597 0.62728173 -0.03377371 0.65257692 -0.03276424 0.67745021 -0.03162506 0.70183069 -0.03036453 0.72564905 -0.02899202 0.74883777 -0.02751801 0.77133138 -0.02595418 0.79306666 -0.02431345 0.81398285 -0.02261004 0.83402183 -0.02085939 0.85312827 -0.01907817 0.87124983 -0.01728411 0.88833722 -0.01549584 0.90434436 -0.01373271 0.91922844 -0.01201453 0.93295002 -0.01036128 0.94547307 -0.00879280 0.95676500 -0.00732844 0.96679666 -0.00598674 0.97554240 -0.00478506 0.98298003 -0.00373925 0.98909082 -0.00286334 0.99385947 -0.00216922 0.99727411 -0.00166643 0.99932628 -0.00136193 1.00001090 -0.00125995 |
Dat file parser warnings
|
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for NACA 25112 (naca25112-jf)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca25112-jf | 50,000 | 9 | 17.7 at α=2.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca25112-jf | 50,000 | 5 | 28.6 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca25112-jf | 100,000 | 9 | 42.5 at α=10° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca25112-jf | 100,000 | 5 | 46.7 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca25112-jf | 200,000 | 9 | 62.9 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca25112-jf | 200,000 | 5 | 63.9 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca25112-jf | 500,000 | 9 | 89.9 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca25112-jf | 500,000 | 5 | 85.9 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca25112-jf | 1,000,000 | 9 | 109.4 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca25112-jf | 1,000,000 | 5 | 98.1 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||